Inspired By Reality

These are some spacecraft designs that are based on reality. So they appear quite outlandish and undramatic looking. In the next page will appear designs that are fictional, but much more breathtaking. Obviously the spacecraft on this page are all NASA style exploration vehicles, they are not very suited for interplanetary combat (well, most of them at least).

For slower-than-light star ships, go here.

Many of these spacecraft have a table of parameters. You can find the meaning of many of them here. Numbers in black are from the documents. Numbers in yellow have been calculated by me using the document numbers, these might be incorrect.


I'm toying with the idea of making some spacecraft "trading cards."

S.S. Absyrtis

ABSYRTIS
PropulsionFictional Thermo-catalyst
Isp2,680 sec
Exhaust Velocity26,300 m/s
Propellant mass flow935 kg/sec
Thrust2,500,000 kg
24,600,000 N
Inert mass281,000 kg
Payload mass56,000 kg
Dry mass337,000 kg
Propellant mass672,000 kg
Wet mass1,013,000 kg
Mass Ratio3.0
ΔV28,900 m/s
Initial Accel7.4 g
(2.5 g?)
Burning time720 seconds
(12 minutes )
Optimum chamber temperature3,450° C
Height60 m
Maximum dia6 m

Yes this is a fictional spacecraft, but it was designed by G. Harry Stine, former project engineer on the Viking and Aerobee rocket programs at White Sands Proving Ground. Mr. Stine used the Absyrtis in his novel CONTRABAND ROCKET written under the nom de plume "Lee Correy".

For purposes of the novel the Absyrtis needed a specific impulse higher than standard LOX/LH2 (450 sec) but lower than a full blown nuclear thermal rocket engine. In time-honored fashion Mr. Stine created some handwavium out of his imagination. He postulated a catalyst propellant: but nuclear, not chemical. His "thermal-catalyst" fuel (aka "thermo-juice") is perfectly stable under normal conditions, but when subjected to a certain neutron flux the molecule explodes. The resultant high-velocity gas is further accelerated by neutron heating. Mr. Stine chose a specific impulse of 2,680 seconds, which is better than a closed-cycle gas core nuclear rocket but less than an open-cycle gas core.

I did some cross-checking on the performance numbers and they all seem to check out. Except I calculate an initial acceleration of only 2.5 g instead of the listed 7.4 g. No doubt I made a mistake in aritmetic.

I figure that the delta-V is about 29 km/sec, which is pretty good actually. Enough for a one-way orbit-to-orbit Hohmann transfer to pretty much anywhere in the solar system. Although many of those transfer will take more than a decade of transit time.

You can see a size comparison of the Absyrtis and other spacecraft here.

CONTRABAND ROCKET

(ed note: in the year 2050, our heros are members of the Southwestern Rocket Society (SRS) fan club. The fans want to travel in space in the worst way, but civilians are not allowed to fly in their own ships.

On a field trip to Luna Louis' rocket junkyard they are stunned to find the space ship Absyrtis sitting in the lot. As it turns out that ship was Mr. Louis' last command when he was in the UN Space Force, and when the ship was decommissioned he managed to obtain it at scrap metal prices.

Club president Chubb Delany has an insane idea. He tells Mr. Louis that the club would love to refurbish the old ship, and fly it on a short hop to Luna. With Mr. Louis as captain.

Mr. Louis says if the club will promise that, he will give the ship to them free, along with any used rocket parts in the lot needed for the refurbishing.)

THE SPACE SHIP ABSYRTIS

UN Space Force interplanetary cruiser, missile launching, universal type, Argonaut Class

The space ship Absyrtis of the Argonaut Class saw twenty-two years of service as a ship of the UN Space Force line fleet, and an additional five years in support and reserve capacity. She was a member of the fleet which put down the Asgard Space Station Rebellion. After her modification to an express cargo vessel, she was instrumental in the sustenance of the outposts and colony on Venus. Subsequent modifications enabled her to serve the Jovian moons. She operated as a support ship in the Titan expedition to Saturn, but obsolescence forced her to confine her operations to the Earth-Luna area in which she served in many capacities until being mothballed and placed in circum-terrestrial orbit. She was finally returned to White Sands Spaceport, sealed, decommissioned, and left to stand for two years before being sold to a local salvage yard.

She was the first ship to bear the name and the last of her class to be decommissioned and removed from the Big Book. Her reliability earned for herself and her crews three ratings of excellence, two efficiency awards, a UN citation, numerous national commendations, and the International Astronautical Federation plaque in connection with her work in the Jovian moons area.

  • Manufacturer: Hueco Spacecraft Inc., White Sands Spaceport, America, Terra.
  • Commissioned: May 2018

     “Have you, now?” Louis said quizzically. “And how do you like the farce space flight is now?”
     “Farce?” Chubb echoed.
     “Farce, son. They’re too sloppy these days. It’s too easy. Automatic controls. Nuclear drives. There was a time when space flight was an art! Yes, sir An art! Not button pushing! Used to load her up with thermo-propellants, hit the firing button when the clock said so, and fly her by the seat of your pants and the astrostat! All the time wondering if she was going to blow! … That was space flight! Pilots, they call themselves! Bah! Bus drivers is what they arel” He settled back in his chair and jerked his thumb over his shoulder. “Now, in the good old days, it was different. Take the old Absyrtis back there on my lot …”
     The junk yard was old stuff to Chubb, but LeRoy was utterly amazed at the terrific amounts of junk of all types. As the old skipper led them out to where the Absyrtis towered seventy meters over the low sheds, the real estate man found himself making mental estimates of the combined worth of this desert land and the tremendous inventory on it. It was plain to see that Luna Louis was not a down-and-out old spacebum. “Captain, there seems to be a little bit of everything here. How’d you come by it all?”
     “Ships are scrapped all the time, mate,” Louis replied.
     “Why? Do they wear out?”
     “No, sir! They just get obsolete,,and it becomes cheaper to build a new ship than to modify the old one,” Luna Louis explained. “The Bureau of Space Commerce has some pretty strict rules about the condition of space craft; when a ship reaches a certain age, they usually down-check it on principle …”
     “How’s that?”
“They figure it’s old enough that if something hasn’t happened to it yet, it will. But with a little decent maintenance and repair, a space ship’s good for over a hundred years … and a power plant’s good for a lot longer than that because its operating time is only a fraction of that of its ship.” Louis paused for a moment. “Of course, we get a good deal of equipment from wrecked or damaged ships. Got one lad who does nothing but sit up on the roof with a pair of binoculars watching for ships that don’t make the grade …” He let it drop at that because they had reached the boat-tail of the Absyrtis.     Chubb stopped to catch his breath and looked up. In addition to the rust streaking her sides, there was no doubt that this ship was old. The tall, slim, almost regal lines were not those of a modern ship. Modern ships looked efficient; they were. The Absyrtis was merely beautiful, a work of art, the result of a designer with a sense of line and sweep and proportion who had labored over his drawing boards doing work which he must have loved. It was reflected starting from the parabola of revolution of her nose cone down her sleek, unbroken sides to the graceful curve of her boat-tail with its six gaping thrust chambers, and in the swallow-like profile of her drooping wings. It belonged to another day of space flight.
     “Shipped many a ton of lunar ore in this bucket,” Louis said in recollection. “But she was a bitch to handle under thrust. Shake? Man, she’d shake your teeth right out! And the center of pressure would tend to wander forward of the center of gravity if you didn’t watch the mass distribution. Let’s go aboard.” He grabbed a rope ladder hanging down the side of the ship and clambered spryly aloft with an agility which amazed Chubb and LeRoy.
     LeRoy followed and Chubb waited until the other had gained the lock high on the ship before he entrusted his full weight to the ropes. He didn’t look down; if he had, he would have frozen to the ladder with vertigo. He kept his eyes aloft and climbed steadily, hand-foot-hand-foot. He was out of breath when he stepped through the air lock and looked around.
     The tour of the old ship was fascinating. Chubb’s eyes were alight the entire time. It was like a childhood dream come true. It brought up memories entombed by the years and Chubb remembered the toy spaceships which looked like the Absyrtis and the drawings he had hopefully sent to the Space Force at White Sands, crude sketches of a “Sooper Space Combat Rocket”. And there were forgotten memories of a chubby little boy playing spaceman in that pile of boxes in the back yard, dreaming of a space ship the image of which was the Absyrtis.
     Just being in her gave him a feeling of satisfaction he had not experienced for years. Feeling the cold metal of her companionways and smelling the ancient, musty odors of far-off worlds which still lingered in her made him suddenly realize with a pang of sorrow and regret that this could have been his—could have, if he had had a different gene makeup (Chubb's genetic makeup predisposes him to be overweight, not allowed in space crews).
     The Absyrtis was far from a complete space ship. Most of the power plant essentials were missing, the electronics had been stripped, and there were no astrogation instruments. The Absyrtis had seen hand tools, but not a cutting torch.
     “Give her just a few essentials and she’s ready to lift,” Louis remarked, sitting down on an acceleration couch in the barren, echoing control room far forward in the nose. “Many’s the time I’ve sweated it out on this couch, mates. But this old bucket never failed me. A taut, reliable old ship she was. After we converted her to thermo-juice, she saw Mars and Venus and Ganymede. Bailey took her out to Titan once after I got stuck on dirt for keeps. But she knows her way into Dianaport by heart; hardly have to lay a finger on the board for a landing. She just sniffs her way in.”
     “What are you going to do with her, skipper?” Chubb ventured to ask.
     “She’s the last of her kind, mate. The pure-nuclear ships have taken over now. And a new kind of spaceman is flying them. We’re both obsolete, so she stays here with me. Oh, maybe one of these days I’ll get me a red-hot crew together. We may not get high enough to crash, but we’ll still get oft the ground again. The regulation hounds will try to stop us, but to hell with them! It’s a sad thing, mates, when the laws won’t let a man do what he wants or even kill himself as long as he doesn’t hurt anybody else in the process.” The old man’s eyes were on the empty holes in the control panels where instruments, lights, and switches should have been.

(ed note: Mr. Louis takes them up on their offer. In exchange for refurbishing the Absyrtis and keeping Louis as captain, he will give them the ship for free)

     Refitting the Absyrtis turned out to be quite a task. The old ship lacked more than was apparent on a cursory inspection. As a result, Chubb closed the doors to his consulting office in order to devote his full energies to the project.
     So he moved in with Luna Louis, sharing the old bachelor’s quarters with him. It was far from being luxurious, but Chubb was having the time of his life. He didn’t really care where he slept or when he ate; he had his hands on a space ship at last.
     Louis turned out to be less senile than any of them expected. He seemed to snap out of his dreamy moods. The transition was strange to behold. Once again, he stood straight and his voice carried the tone of authority and casual competence. His eyes became alert, and his mind sharpened like a rusty knife edge that has been put to the whetstone at long last.
     The youngsters of the SRS were by far the most persistent at the work site. They came in droves on Friday afternoon and stayed on the job until Monday morning when they dragged back across the desert to classes or to their jobs. Many of them came out during the week to perform the many tasks at hand.
     Their first job was a complete and minute inspection of the ship as she stood. No manuals on the Argonaut Class could be found, but Luna Louis turned out to be a man of remarkable memory.
     “Hey, skipper, this valve seat mikes a tenth of a millimeter less than the blade. What gives here?” LeRoy called from the power room on the temporary intercom Bert Eggstrom had rigged.
     Louis answered from the forward radar blister, “Where did it come from?”
     “The feed heater just abaft of the forward tank bay.”
     “That sounds about right, Mister Finch. What’s the condition of the seat and gland packing?”
     “Packing’s shot. But how can this valve seal?”
     “Don’t worry about it. It gets hot in that forward feed heater. Thermal expansion of the seat causes that valve to seal tighter than your old britches. Get the part number off that valve, and we’ll see if maybe I’ve got some packing for it. Pull the whole valve and take it down to the shop.”
     “Right-o, skipper!”
     “Hey, skipper?” Chubb’s voce echoed up the main ship well. “Got a minute?”
     Louis turned to the youngster who was working in the blister with him. “Yank that sweep selsyn, Jimmy. The rotor’s shot. I think maybe one of the units from that old Mark Fourteen radar out in the yard will fit. Don’t bother with those cap screws; knock it loose with a hammer, because you’ll have to drill and tap new holes anyway."
     “How about these waveguide junctions, skipper?”
     “Put the torque wrench to them. They’ll warp back,” the old skipper told him, handing him his tools and crawling out of the little hatch into the main portion of the ship. Wiping the sweat from his neck with a piece of waste, he yelled down the well, “Up here, mate!”
     Chubb came puffing up the ladder from below. “Here’s a survey of the equipment in the boat-tail, skipper.”
     Leafing slowly through the sheaf of papers handed to him, Louis mused, “Not as bad as I expected.”
     “What do you mean, skipper? Half the structural members back there are bent, broken, or missing! Engineering-wise, it’s flimsy as a paper bag!”
     “And just what do you know about space ship structures, Mister Delany?” Louis asked sarcastically. “The tail of this bucket was grossly over-designed. We ripped out those members years ago to make room for the thermo-juice drive.” He handed the papers back to Chubb and told him, “Take them down and give them to that gal who’s doing the consolidation. I’ve got most of these missing parts—or something that will do the job.”
     “Check, skipper. Tank bays and radars are the only lists we need now. Maybe we’d better start thinking about moving the ship out to a launching pad.”
     “Why move her?”
     “Huh?”
     “A fine “engineer you are! What would the costs be? I've got the parts, the shops, and the tools right here.”
     Chubb thought about this. “You mean refit and lift from here?
     “Is there a better place?”
     “But it’ll wreck your yard when we lift, skipper.”
     “So it will. But once we raise ship, mate, I’ll not be needing this yard any longer.”
     Chubb stared at him for a moment, then quietly went out the hatch and clambered down the hastily-rigged servicing tower. A month ago, he would have paled at the thought of hanging on a slender ladder fifty meters up. He had in fact done so. But it didn’t bother him now, and he was in much better physical shape. It was a matter of pride to him that he had managed to lose five kilos.
     Wandering back through the yard toward the hut they were using for an office, he noticed the change in Luna Louis’ junk yard. Old tools had been ressurected from the heaps, cleaned up, and placed in sheds. Under a ragged tarpoline, three youths were hydrostating valves and pressure vessels; beside them was a jury-rigged flow bench. Farther down the line, he passed a leaning shack in which Bert and several other men were working over old radar gear. A sign over the door proudly announced, “Department of Witchcraft and Sorcery. Slightly Used Pentacles and Klystrons For Sale.”

     It took five long months filled with scrounging for old parts, digging around in junk yards all over North America, and draining of funds. Chubb’s savings were long gone. Luna Louis had converted everything he could to cash. Al Olson, being independently wealthy, kept the project on its feet.
     Everybody worked their hearts out. There were long hours. There were the inevitable minor accidents. There were daily crises which threatened to wreck the whole thing. Louis had his hands more that full working with a very green crew. Everybody made mistakes—but nobody made the same one twice. But the day finally arrived when they could start making dry runs of the ship and her equipment.
     The old Absyrtis didn’t look the same at all, Sporting a new coat of brown and yellow paint—a purposely difierent color and marking scheme than that used anywhere else —she was practically a new ship inside and out.
     Chubb stood surveying her in the late afternoon sunlight, taking a break in his schedule for a cigarette. Yes, every one of the hundreds of men and boys who had worked on her could take real pride in her now, he knew. A few more checks, radioactive bricks for her reactor, and propellants were all she needed—plus a trained crew.
     Olson was out pulling the legal strings for the reactor bricks. Chubb had no idea how they were going to be pried loose from the Bureaus of Nuclear Energy, but Al had assured him that there would be no trouble.
     The propellants? Well, Chubb was expecting ten tank cars into El Paso any day. There was no problem there. The “go-juice” for which the Absyrtis had been designed was a commercially-available chemical which would release its energy by thermo-catalytic action. It was cheap, but it was no longer used for space flight.
     The rocket engine is a basically useful device. Rocket engineers found this out many, many years before when they became aware that the military subsidies following the Second World War might not last forever. A rocket can do more than push. It can generate tremendous volumes of gas. It is an essential device for high-speed, high-temperature chemistry. And the jet of hot gases man dig holes. In the open-pit copper and iron mines all over the world, the snarl of rocket engines was a common thing as their exhausts dug holes faster and more economically than the best carbide bits.
     The crew was his only real worry. He knew they were still green as grass, himself included. Space Commander McLaughlin had been right on one point: you don’t learn it all out of the books. Some people had picked up Louis’ training with little effort; others just couldn’t understand the difference between a fitting and a flange or between a selsyn and a klystron, no matter how high their enthusiasm had been.
     “All hands clear the ship!” Louis’ voice came from a portable megaphone from the lock high on the side of the ship. “Stand clear for pressurizing and water-flow checks!”
     Chubb was joined by Bert, who was handling the electronics and had no part in this check of the propulsion system. “Ran the final checks on the radar today, Chubb. That doppler system is all hot to go. Same with the guidance and control.”
     “Good! Did you get the running rabbits off the surveilance screens?”
     “Yeah, found a mis-matched waveguide in the antenna system. How’s Greg doing with the air system?”
     “Had chlorogel all over Deck D the last I saw him,” Chuhb replied. “Sprung a leak in the irradiation chambers.”
     “Tough luck.”
     “He’ll get it fixed. He’s good.” Chubb watched the silent ship for a moment, then asked, “Say, Bert, maybe it’s none of my business, but how come a sharp electronic engineer like you never got into space in the first place?”
     “Oh,” Bert said offhandedly, “eyes for one thing. Plus the fact I’m a lunger.”
     “T-B? You don’t look like it!”
     “Hell, man, I’ve only got one lung—and that’s full of calcification. Why do you think I came to this country? Same reason as Greg: climate.”
     Great space! Chubb thought. What a crew this ship’s got! Greg with arthritis, Bert with one lung, LeRoy with a heart, and the skipper ripe for the grave! And me, twenty kilos overweight!
     “Stand by to pressurize!” came the call from the ship. Through the thick hull of the Absyrtis the two men on the ground heard the slam of valves and the high-pitched, ringing hiss of pressurized gas filling the propellant tanks.
     Nothing ruptured; the tanks held their pressure.
     “What are they doing?” Bert wanted to know.
     “We’ve got a dummy propellant load of water in the tanks,” Chuhb explained, rocking back on his heels with his arms on his hips. “They’ve pressurized to detect leaks and to see if the system will hold pressure. Next they’ll pop the main propellant valves and run the water out through the rocket nozzles to check flow rates and pressure drops.”
     “How can they run the propellant pumps without the reactor to drive them?”
     “They won't need the pumps. LeRoy and the skipper just want flow characteristics. They know what effect the pumps will have and they … Hold it! There they go!”
     It was quite a show. A terrific roar came from the stem of the ship, but no flame lashed out. Instead, the rocket nozzles sprayed solid streams of water which ran off onto the desert sands in a small flood. Thousand of gallons of water spewed out before the flood suddenly ceased with a bang and a hammering sound.
     “Wow! I’ll bet that shut-down opened a dozen joints!” Chubb took off across the desert like a huge ballon being driven before a gale.
     The power room was a mess when he climbed into it. LeRoy and his crew were trying to tighten fittings and stem the gush of water. There was still considerable water remaining in the tanks. Everybody was soaking wet. Chubb grabbed a box wrench, snugged up a leaking fitting, and shouted to LeRoy “Vents open?”
     “Hell, yes! Get that flange tight before we drown!”
     “Open your dump valves and drain those tanks! You’ll never get these fitting tight with ten meters of hydraulic head on them”
     LeRoy leaped for the jetman’s couch and threw switches. “Electrical system’s shorted out by water! Open that hand valve next to you, Chubb!”
     Once the situation was under control, Chubb—looking like a water-loogged whale—sat down on an auxiliary generator and observed. “I thought you guys knew this power room. What a sad show! What would you have done in a real emergency?”

(ed note: A torrential thunderstorm undermines the landing pad andthe ship tips. The ship is saved, but...)

     The Absyrtis was canted over at a five-degree angle, and that was that. Chubb and Luna Louis surveyed the situation the next morning and came to the conclusion that any attempt to right the ship might cause her to fall even farther. The only answer was to secure the ship in its present position and proceed. Taking LeRoy’s suggestion, they fastened quick-release clamps on the guys, then promptly safetied them against accidental release.
     The rest of the check-outs on the “Leaning Tower of White Sands”—the name hung on the ship by Greg Shearer —went off more or less as scheduled. They weren’t all successful at first. Bert Eggstrom was the only one who didn’t have more than his share of troubles, but he had them nonetheless. The communication gear worked like a charm, and Bert was very proud the night he logged his first contact with Asgard Space Station as it went over. The computer and the autopilot finally made four consecutively successful dry runs. He had trouble with the radar; instead of tracking the high-flying evening antipodal rocket as intended, it locked onto a flight of ducks migrating south. But it tracked.
     LeRoy kept on finding leaks, sticky valves, broken welds, and loose nuts everywhere. Most of his trouble was with a very green crew. The college students working for him did extremely well, but he had trouble with other kinds of people. Imagine trying to teach a dry-goods salesman how to run a smooth weld.

     Greg Shearer was having trouble with the air system, the water recovery system, and the hatches and locks. Being a bachelor like Chubb, he didn’t have the same kind of trouble LeRoy was having, but he had trouble enough nonetheless.
     He had recruited every member of the SRS who would work and who had, like himself, a green thumb and a knowledge of organic and catalytic chemistry. On the first pressure test, gaskets leaked all over the place, but the worst part came when Greg replaced the standard air with the oxy-helium space mix from the air system. It drove everybody choking from the ship. The ship air from the blowers smelled something like a cross between a garbage dump, a stable, and a locker room. In disgust, he and his crew had to replace every bit of chlorogel solution in the system—while wearing respirators.
     In addition, the ship’s water came out a putrid brown for five days while he fiddled with the old and finnicky water recovery system.
     He was still working with it when Luna Louis came around with Chubb for a final inspection on the refitting. Louis inspected the ship with a critical eye, finding things that nobody expected. He suggested here, corrected there, bawled out ninety-percent of the crew for blunders and oversights, but finally pronounced the ship as ready as it ever would be for final checks, provisioning, and space.
     It should have been a day of rejoicing, but for Chubb it was one of anxiety. During the quiet evening hours after supper when everybody sat around listening to Luna Louis spin old space tales of faraway worlds, Chubb could not keep his mind off the subject.
     The final check of the ship would require that the reactor be activated. This meant heavy water (moderator) and thorium, and as far as he was concerned that might take an act of God to get. The Bureau of Nuclear Power didn’t pass that kind of stuff around like tin pennies.

     They went in the personnel hatch. The interior of the ship was considerably different than it had been the day they had first walked her decks. New paint glistened on the bulkheads, and the smell of oil and solvents and men was in her again. After going up the main ship well, they emerged into the control room. The new pastahide cushions on the couches shone in the light of the resurrected flouro-units, and the gaping holes in the panels had been filled with instruments, gauges, switches, and winking lights. The computer rack now held a small electronic brain which was capable of flying the ship; Bert had managed to pick it up from Space Force surplus and had rebuilt it. It was capable of reading-in data in any number of ways: from sensing elements in the ship, from ground radio commands, from control panel commands, and from a self-programming keyboard. Its read-out was equally as versatile. Bert regarded it as being several grades smarter than any of the SRS men working on the ship.
     “About time you showed up,” Louis remarked caustically. “Grab a cup of joe and sit down. You might favor me by pouring me another cup while you’re at it. Can’t operate without coffee.” Luna Louis, true to Space Force tradition, had set up three indispensable things immediately on the Absyrtis; in order they were a coffee mess, a loudspeaker system, and a wardroom of division officers who really ran things.

     “It has to be that way, Mister Olson,” Louis pointed out. He took a long swig of coffee and went on, “And now we have a slight logistics problem: thorium and heavy water. What do you have to report on that, Mister Olson?”
     “It’s on its way. Be here tomorrow.”
     Chubb sat up and knocked his head against the bottom of the couch above. “Along with the BSC boys who will promptly hang a red tag on the lock?”
     Al Olson smiled knowingly. “Not yet. It seems the UN once passed a Nuclear Energy Act which has been on the books since 2005. It guarantees the delivery of available radioactive substances to non-profit organizations utilizing it for other than commercial purposes. After a little talk with the BNP (Bureau of Nuclear Power) Regional Director in Albuquerque, the way was paved.”
     “What did it cost?” Bert wanted to know.
     “It’s still BNP property under consignment lease to us,” Al replied.

     Luna Louis had to supervise the installation of the thorium and heavy water. The current BNP manuals possessed by the two men did not include the procedure for the Argonaut Class, the reactor being a long-obsolete model. This caused Chubb to ask Louis anxiously, “Skipper, are you sure this old reactor has enough soup to lift this ship?”
     “Mate,” Louis said huffily, “if the engineers hadn’t decided to go to the pure nuclear drive, they’d still be using this type of reactor. This old Mark Seven’s a damn-sight more reliable and efiicient per kilogram of mass and has a lower operating count so the ship shielding is lighter. What licked this type of drive years ago was the propellant mass you have to carry … and thermo-juice was more expensive then. The nuclear stuff cost less when they changed over, but this bucket’s flown for twenty years with that fish bowl, and it hasn’t had so much as a pin-hole leak in the heat exchanger. The Baja California power pile can't boast that record, mate!”
     At last, they were ready for the final checks. LeBoy, who had hand-picked his engineering gang, treated the reactor with a great deal of respect now. The power crew started using the particle counters which had been racked at the ready on the power room bulkhead for months. They raised the temperature to a stable plateau capable of running the ship’s generators and of charging the batteries, although ground power was still available to help them out. The other divisions began a gradual phase-over to reactor power but only as an emergency measure.

From CONTRABAND ROCKET by Lee Correy (G. Harry Stine) (1956)

Aerojet Rocketdyne LEU NTP

This is from LEU NTP Engine System Trades and Mission Options (2019) and Why NTP Works for Mars Missions (2020)

Aerojet Rocketdyne has been studying manned Mars missions using spacecraft propelled by solid-core nuclear thermal propulsion (NTP) using low-enriched uranium (LEU). They are using LEU because the military and the International Atomic Energy Agency gets very paranoid about civilians getting their hands on uranium that is anywhere near weapons-grade. Specifically Aerojet Rocketdyne wants to work with High-assay low-enriched uranium (HALEU). LEU is enriched from 2% to 20%, most commercial reactors use enriched from 2% to 5%, HALEU is enriched from 5% to 20%. There had been some other NTP proposals from other companies using HEU, but those proposals had been shelved due to said paranoia.

By using HALEU, Aerojet hopes to design an engine with a specific impulse above 900 seconds (exhaust velocity above 8,800 m/s), which is an improvement over chemical rocket's pathetic 450 seconds (The BWXT company is looking into 19.75% enriched HALEU as well). Aerojet has been working on this for a couple of decades, patiently altering the ground rules to accomodate curve balls thrown by NASA (the most recent being structuring the mission around NASA's questionable Lunar "Gateway" aka the Lunar Operations Platform-Gateway or LOP-G).

Aerojet figures that their engine has a variety of applications:

  • Moving heavy cargo (e.g., large landers) to Mars within 200 days
  • Delivering larger orbiter spacecraft to Jupiter (e.g., two to three times the size of Juno)
  • Delivering orbiters (as opposed to a mere fly-by) the size of New Horizons or larger at the outer planets with transit times less than ten years
  • Fly-by space probe missions to the outer planets using just the LEU NTP core stage flown directly from the faring of the Space Launch System (if the SLS ever sees the light of day), without needing multiple launches and space assembly of a full LEU NTP spacecraft

The LEU Nuclear Thermal Propulsion System

Aerojet goal was to optimise the engine for:

  • Maximum possible specific impulse
  • Minimum possible reactor mass
  • Longest operating life of nuclear criticality, meaning maximum number of hours of thrust you can get out of the engine before the blasted thing clogs up with nuclear poisons and stops working.

Both Aerojet and NASA have done numerous studies that suggest for a crewed Mars mission, the optimum propulsion system is an array of three NTP engines with a thrust of 25,000 lbf each (111,200 N) for a total thrust of 75,000 lbf (333,600 N). Each engine has a reactor with a thermal power of 550 MWt which heats the propellant to greater than 2,600 K for a specific impulse of 890 seconds or greater (exhaust velocity of 8,700 m/s or greater). Each reactor will require 50 to 100 kilograms of LEU. A single engine is comparable in size to a standard chemical RL10B-2 engine.

LEU NTP Mission Trade Studies

Aerojet wants to assure the reader that they have been continually doing trade studies on Mars Mission spacecraft using this system since 2016. Because NASA has made the requirements a freaking moving target. First the study using NASA's initial ground rules, then a new study when NASA updated to the Evolvable Mars Campaign (EMC) in 2016, and then yet another new study when NASA incorporated the information from the Mars Capability Studies (MSC) team in 2018 (“In-Space Transportation for NASA’s Evolvable Mars Campaign”, and “Transit habitat Design for Mars Exploration”). The LEU NTP had many advantages:

  • the spacecraft mass can be reduced by using cis-lunar aggregation Near-rectilinear halo orbit (NRHO) orbit (coincidentally the orbit of the NASA Lunar Gateway, surprise surprise) and a lunar-distant high-Earth-orbit (LD-HEO) type orbit for Terra departure and return.
  • using the above orbits still allows a transit time from Terra to Mars of five to six months
  • the spacecraft can be boosted into orbit piecemeal in as little as four or five SLS launches using the 8.4m payload faring

However, the priority was to reduce the Terra to Mars to Terra transit times. Because the longer the trip, the more radiation exposure suffered by the crew due to deadly Galactic Cosmic Rays. It is impractical to carry enough radiation shielding for full protection (meaning the a viable spacecraft might be impossible to design), so the fallback position is to reduce expsure time.

In Figure 4 below, the Low Thrust (LT) option is a conventional solar-powered ion drive rocket while the High Thrust (HT) option is the LEU NTP rocket. As you can see the LEU NTP rocket has a drastically reduced radiation expose in all its options.

Note that there are two types of trajectory:

TrajectorySurface
Stay
(days)
ΔV
(km/s)
DrawbackTransit
time
(days)
Wet Mass
(MT)
ConjunctionLong
~600
ΔV 4 to 5
1/3 ΔV of Opp
Needs bigger
surface base
~325~200
OppositionShort
~30
ΔV 10 to 13
x3 ΔV of Conj
Needs more
powerful engine
~680~720

Aeronutronic EMPIRE

Aeronutronic EMPIRE
EngineNERVA
solid core NTR
Thrust200,000 N
ΔV5,300 m/s
Length47.6 m
Crew6
Mission
Duration
611 days
Wet Mass170,100 kg

Information for this entry are from EMPIRE Building: Ford Aeronutronic's 1962 Plan for Piloted Mars/Venus Flybys, Humans to Mars: Fifty Years of Mission Planning, 1950-2000 by David Portree, The Empire Dual Planet Flyby Mission by Franklin Dixon, EMPIRE: Background and Initial Dual-Planet Mission Studies by Fred Ordway et al. and the entry in Astronautix.


Back in 1962, NASA's Marshall Space Flight Center's Future Projects Office (FPO) decided to get serious about manned exploration of other planets. They commissioned a study with the contrived name Early Manned Planetary-Interplanetary Roundtrip Expeditions (EMPIRE). Three mission study contracts were awarded. General Dynamics would study Mars orbital missions. Lockheed would study Mars flyby and orbital missions. And Aeronutronic would study Mars-Venus flybys.

Doing a flyby instead of a landing was disappointing, but the FPO figured you need to start with baby steps. A flyby would require less than half the delta V of a full blown Mars orbital or landing mission. Nowadays we would wonder why bother to send astronauts when you could just use an unmanned space probe. However, back in the 1960s automated probes were nowhere near reliable enough for such a mission.

As a consolation the studies were allowed to include NERVA nuclear thermal rockets. A mission to Mars using honest-to-jonny atomic rockets, by Jove!


Aeronutronic examined the work of Dr. Gaetano Crocco. In 1956 he published a mission that would require only one burn to inject the spacecraft into the mission trajectory, it would coast for the rest of the mission. The spacecraft would do a flyby reconnaissance of Mars and arrive back at Terra exactly one year to the day (so Terra would be back at the starting point). All with no additional engine burns. Naturally the spacecraft will need an additional 13.5 km/sec delta V in order to brake into Terran capture and landing, but this can be done without fuel by using aerobraking. This mission was called the Unperturbed Non-Symmetrical Trajectory which was immediately shortened to the Crocco Trajectory.

The astronauts would observe Mars through telescopes during the brief flyby. Annoyingly, if the ship came closer to Mars than about 1,300,000 kilometers, the gravity well would bend the trajectory such that the ship would miss Terra and the astronauts would die a lonely death in deep space. After going to all this trouble for a Mars space mission it is frustrating to be prevented from getting any closer than three times the Terra-Luna distance.

Dr. Crocco had a solution. The ship could get closer to Mars. As long as the trajectory was designed so that the spacecraft did a bank-shot off of Venus' gravity well to correct for Martian bend. The opportunity to do observations of Venus was a nice bonus. It did, however, increase the mission duration from 365 days to about 396 days.


However Aeronutronic found a major drawback to the Crocco Trajectory. The spacecraft (in a 300 kilometer LEO) would need a sizable 11.95 km/s delta V to use it (I know the table says 10.1, ignore it).

There was another option: the Unperturbed Symmetrical Trajectory. This would need less than half the delta V, only a mere 5.3 km/sec. The drawback here was the mission would increase by a proportional amount, to 611 days.

Aeronutronic went with the Symmetrical trajectory because a lower delta V means a lower propellant requirement, which means a much lower total ship mass to be boosted into LEO. Such is the tyranny of the rocket equation. The increase in required oxygen and food was relatively minor.

Another drawback is the aerobraking delta V increases from 13.5 km/sec to 15.8 km/sec, but again the required increase in reentry vehicle mass was worth it.


How much spacecraft mass exactly do you save by reducing the delta V from 11.95 to 5.3 km/sec? A metric butt-load, which in this case means a reduction from 1,017,000 kg to only 170,100 kg! The nuclear symmetric spacecraft is only 17% the size of the nuclear Crocco ship.

Aeronutronic did briefly look at chemical rockets, but they would have even more mass. They were rejected.

The spacecraft would use a single NERVA engine with 200,000 newtons of thrust. To kick the spacecraft for 5.3 km/sec of delta V it would have to burn for a whopping 48 minutes. This was perilously close to the operational lifetime of such an engine. The burn time could be reduced if a larger engine with more thrust was designed, but Aeronutronic figured this could not be done in time for the 1970 launch window.



The first stage is the NERVA engine, a core tank, and six perimeter tanks clustered around the core. First stage injection consumes 56.2 metric tons of propellant. After all of the first stage propellant is burnt, the perimeter tanks are jettisoned (3.3 metric tons). The empty core tank is retained because that is the only thing connecting the NERVA engine to the rest of the spacecraft. The ship's mass has dropped from 170.1 metric tons to 119.1 tons.


The second stage is the NERVA, the empty core tank, and eight tanks clustered around the habitat module. Second stage injection burns all the 34.7 metric tons of propellant. Then the NERVA and the empty core tank are jettisoned (11.9 metric tons) creating a orbiting artifact that will be dangerously radioactive for several thousand years. The 8 second stage tanks are retained as meteoroid shielding for the habitat module. The ship's mass has dropped to 69.1 metric tons.

The spacecraft no longer needs a main engine since it is in the arms of Saint Kepler.


The ship is now reconfigured into orbit mode.

The twin habitat modules extend on telescoping arms and the ship spins at 3 rpm to create 0.3 g of artificial gravity (SpinCalc tells me each habitat module has to be 29.8 meters from the spin axis). Sixteen-meter-diameter communication dish antennas blossom from the ends of each habitat module, aimed at Terra.

One of the SNAP-8 radioisotope thermal power generator (RTG) unfurls its heat radiator and energizes. The spacecraft's power budget is 300 kW. The second SNAP-8 is held in reserve as a backup. I am wondering if this is a mis-print, since I was under the impression that SNAP-8 was a nuclear reactor, not RTG. I was also under the impression that RTGs were hard pressed to produce more than 1 kW.

The core contains the 20-metric ton command center/storm cellar clad in 50 centimeters of polyethylene plastic for radiation shielding from solar proton storms. The core also contains the navigational stable platform, a small compartment for weightless experimentation, 10.9 tons of chemical fuel for the trajectory correction rockets (packed around the storm cellar to provide extra shielding), and the Terra aerobraking re-entry vehicle on top of a two stage retro-pack.

The habitat modules have 126 m3 of space, giving a luxurious 21 cubic meters per crew person instead of the bare minimum 17 m3. The storm cellar is only 8.4 m3 giving a miserly 1.4 m3 per crew person, but storm cellars are always cramped.


The watch-bill does its best to keep the crew busy during the 21 month mission.


After the reconnaissance pass by Mars, and the course correction pass by Venus, the spacecraft approaches Terra. The crew enters the re-entry vehicle, and moves away from the abandoned spaceship (which sails into an eccentric solar orbit). The two stage retro-pack slows the re-entry vehicle by 2.8 km/s, reducing the relative velocity to Terra down to 13 km/s. The remains of the retro-pack are jettisoned.

The re-entry vehicle slams into Terra's atmosphere and aerobrakes at a brutal 10 gravities until it slows enough to deploy parachutes. The astronauts are rescued and are transported to a hero's welcome, while NASA quickly asks Congress for a budget increase.

AFFRE Mars DRA

AFFRE MARS DRA 5.0
EngineAFFRE
Engine Mass
(reactor)
107,000 kg
Engine Mass
(mod oil)
91,000 kg
Reactor Power2.5 GW
Thrust4,651 N
Thrust Power730 MW
Specific
Impulse
32,000 sec
Exhaust
Velocity
313,900 m/s
Mass Flow
(FF)
3.12×10-5 kg/s
Mass Flow
(Hydrogen)
0.0179 kg/s
Mass Flow
(Total)
0.018 kg/s
MASS SCHEDULE
RCS925 kg
Propulsion268,961 kg
Structure5,899 kg
Heat
Radiators
280,816 kg
Power6,200 kg
Avionics3,118 kg
INERT MASS565,865 kg
Payload170,000 kg
DRY MASS735,919 kg
Propellant345,599 kg
WET MASS1,081,518 kg
Mass Ratio1.47
ΔV120,900 m/s
MARS MISSION
Outbound104 days
Mars Stay60 days
Return128 days
Total Trip292 days

This is from Final Report: Concept Assessment of a Fission Fragment Rocket Engine (FFRE) Propelled Spacecraft (2012), Fission Fragment Rocket Engine (FFRE) Technology and Status (2014), Opening the Solar System: An Advanced Nuclear Spacecraft for Human Exploration (2014)

Robert Werka helped design a fission-fragment rocket engine (FFRE), which was used in an NAIC study for a HOPE mission. Due to the miniscule thrust of the engine, the spacecraft performance was disappointing.

Robert Werka later figured out a new configuration for his FFRE.

As with most engines that have high specific impulse and exhaust velocity, the thrust of a FFRE is pitifully small. Ah, but there is a standard way of dealing with this problem: shifting gears. What you do is inject cold propellant into the exhaust ("afterburner"). The fission fragment exhaust loses energy while the cold propellant gains energy. The combined exhaust velocity of the fission fragment + propellant energy is lower than the original pure fission fragment, so the specific impulse goes down. However the propellant mass flow goes up since the combined exhaust has more mass than the original pure fission fragment. So the thrust goes up.

Now you have an Afterburner fission-fragment rocket engine (AFFRE).

As you are probably tired of hearing, this means the engine has shifted gears by trading specific impulse for thrust.

Shifting Gears
EngineIspThrust
FFRE527,000 sec43 Newtons
AFFRE32,000 sec4,651 Newtons

Robert Werka and Thomas Percy took the standard Human Exploration of Mars Design Reference Architecture 5.0 and designed it using an AFFRE, which is the spacecraft displayed here.

The payload is 170,000 kg: a 35,000 kg habitat module and a 135,000 kg Mars ascent/descent vehicle.


AFTERBURNING FISSION FRAGMENT ROCKET ENGINE

The heart of the engine is a standard "dusty plasma" fission fragment engine. A cloud of nanoparticle-sized fission fuel is held in an electrostatic field inside a neutron moderator. Atoms in the particles are fissioning like crazy, spewing high velocity fission products in all directions. These become the exhaust, directed by a magnetic nozzle.

The AFFRE alters this a bit. Instead of a cylindrical reactor core it uses half a torus. Each end of the torus has its own magnetic nozzle. But the biggest difference is that cold hydrogen propellant is injected into the flow of fission fragments as an afterburner, in order to shift gears.

In the diagram above, the magnetic nozzles are the two frameworks perched on top of the reactor core. It is a converging-diverging (C-D) magnetic nozzle composed of a series of four beryllium magnetic rings (colored gold in the diagram). Note how each frame holding the beryllium rings is shaped like an elongated hour-glass, that is the converting-diverging part. The fission fragment plume emerges from the reactor core, is squeezed (converges) down until it reaches the midpoint of the magnetic nozzle, then expands (diverges) as it approaches the end of the nozzle. At the midpoint is the afterburner, where the cold hydrogen propellant is injected.

The semi-torus has a major and minor radius of 3 meters. The overall length of the engine is 13 meters. The reactor uses 91 metric tons of hydrocarbon oil as a moderator. This means the heavy lift vehicle can launch the engine "dry" with no oil moderator. In orbit the oil moderator can be easily injected into the reactor, at least easier than building the blasted thing in free fall out of graphite bricks.

The shadow shield is only composed of tungsten, to stop gamma rays. I presume that the liquid hydrogen propellant tanks and the 260-odd meter spine distance take care of the neutron radiation, since tungsten doesn't do diddly-squat to stop neutrons.

The afterburner did not quite make up for the low thrust, so they also had to switch fuel from Uranium-235 (500 barns of cross section) with Americium-242m (7,200 barns). This raises the thrust from fission fragments from a disappointing 3% of total thrust to a whopping 40% of total thrust. They figure this engine design can produce about 50 Newtons per gigawatt.

Keep in mind that Americium-242m is a rare nuclear isomer of ordinary Americium-242 occuring in only 0.4% (0.004) of all Am242, so you are going to have process a metric arseload to get enough of the isomer for a Mars mission. Small price to pay for a reusable spacecraft that can do much better than a Hohmann transfer.

The Americium fuel is stored in nine 4,000 kg crash-proof containers. In each container 80 kg of Americium dust is suspended in a concentrated boric acid solution which acts as a neutron poison to keep the fuel inert until needed. The nuclear fuel pumps transfer the solution to the engine, where upon injection the boric acid is flash evaporated, leaving the Americium dust.

The trade-off between thrust, specific impulse, vehicle mass, and mission delta-V can be adjusted to an optimum value due to the magic of afterburners.

The thrust frame the engine pushes against doubles as a radiation shadow shield. It casts a shadow of 22.5° to protect the rest of the spacecraft and the crew. The shield is 5 centimeters of tungsten to reduce the gamma-ray flux. It is assumed that the neutron radiation will all be caught by the oil moderator.


HEAT RADIATORS

Anytime a spacecraft has a nuclear reactor, and it is NOT totally cooled by open cycle-cooling (i.e., all the heat goes out the exhaust jet), it is going to need lots of heat radiators. Or the ship will melt. The AFFRE reactor generates 2.5 gigawatts of power and only about a third of that is exiting in the exhaust (thrust power is 0.73 gigawatts which is 29% of 2.5 GW). Some of the heat escapes as infrared energy out the reactor, but that still leaves about 450 megawatts of heat energy that the radiators will have to take care of. Due to the different temperature levels of various systems there are four separate cooling loops.

Loop 1 operates at 140K and cools the superconducting beryllum magnets. Loop 2 operates at 590K and cools the moderator oil. Loop 3 operates at 1200K and cools the reactor's internal heat shield. Loop 4 operates at 400K and is part of the Brayton power conversion units that convert the reactor heat gradient into electricity.

All four loops use different sections of the 22,791 square meters of double-sided heat radiator array. Looking at the mass schedule you can see the radiators is the most massive system of the entire ship, with the propulsion system a close second. Nothing else even comes close. The radiator is of course trimmed to stay within the radiation-safe shadow.


BRAYTON POWER CONVERSION UNITS

The Brayton units convert the temperature gradient from the reactor heat into electricity. The design was developed by the Glenn Research Center for the HOPE study.

Each of the four units can crank out a whopping 100 kilowatts of electricity. The spacecraft needs 300 kWe, the fourth Brayton is a spare.

This is a luxurious amount of electrical power. Most NASA deep space exploration ship designs have no nuclear electric power. They make do with solar cell arrays and fuel cells, so they have a Spartan power budget of about 15 kWe or so. The AFFRE ship uses much of its spare power to run the cryo-coolers that keep the liquid hydrogen propellant from boiling away. Other designs either use their hydrogen quickly or use inferior propellant like ammonia because liquid hydrogen cryo-coolers are power hogs.


STRUCTURE


PERFORMANCE

The AFFRE has such a spectactular specific impulse that most designs have outrageous amounts of delta-V.

Other engines such as NERVA are so weak that they must need to resort to staging (with entire NERVA engines jettisoned) and even then the remaining part of the spacecraft is about the size of the Apollo command module. Everything else is thrown away.

The AFFRE ship on the other hand returns to Terra basically intact, so you can reuse the entire thing for multiple missions. It has enough delta-V to return and brake into low Terra orbit. NERVA Mars ships typically have the crew bail out in Apollo modules which frantically aerobrake to land on Terra. The perfectly good spacecraft is abandoned into an eccentric solar orbit due to lack of delta-V.

A AFFRE ship can do the Terra-Mars plus Mars-Terra segments of the mission in half the time of a NTR ship. This drastically reduces the required life support consumables mass, and the crew's space radiation exposure.


COMPARISON WITH HOPE FFRE SPACECRAFT

The NIAC study is the FFE ship performing the HOPE Jupiter mission. The CIF study is the AFFE performing the DRA 5 Mars mission (the ship in this section).

The Jupiter ship uses the low-thrust/high-Isp FFE, so the propellant load was quite modest. The reactor was only 1 gigawatt. This made the spacecraft much smaller. However, the low thrust meant the round-trip for the mission lasted 15 freaking years.

The Mars ship uses the higher-thrust/lower-Isp AFFE, so both the propelland load and the spacecraft were quite a bit larger. And the reactor was 2.5 gigawatts. On the plus side the round trip was only 292 days, and it needs less nuclear fuel because it is using Americium instead of Uranium. It also carries three times as much payload mass. Most of the extra inert mass is from the radiator array.

AIST-NTR

This is from Affordable In-Space Transportation (1996)

The study was aimed at how to lower the cost of delivering satellites to geosynchronous orbit (GEO) since that is the bulk of near-term commercial space industrialization. Ariane, Atlas, and Titan IV can cost on the order of $55,000 US per kilogram transported to GEO (in 1996 dollars). This includes payload transport from surface of Terra to low Earth orbit (LEO) and payload transport from LEO to GEO.

They estimated that future reusable launch vehicles (RLV) could reduce by 50% the cost to LEO down to $2,200 to $4,400/kg for payloads in the 9,000 to 18,000 kg range (pretty good estimate, the reusable SpaceX Falcon Heavy has an estimated cost of $2,968/kg to LEO). The report figures that using a resuable first stage and a second stage using the old technology would reduce the total cost of delivering payload to GEO to about $22,000 US, using math they don't bother to explain. They figure that when comparing delivery to LEO with delivery to GEO, one-third to one-half of the price increase of the GEO stage is just because the upper stage is more expensive. The rest is because the maximum payload is lower for GEO, increasing the cost-per-kilogram value because the value for kilograms is smaller.

Bottom line is if you are trying to reduce the total cost of payload delivered to GEO, you will get more bang-for-your-buck if you focus on opimizing the GEO stage of the rocket. The study's goal is to reduce the payload-to-GEO-cost of a rocket with a RLV first-stage by an order of magnitude (to about $2,200/kg to GEO) for payloads in the range of 1,400 to 4,500 kilograms.

They found this is very hard to do.

The top candidtates (lowest life-cycle cost) were expendable solid chemical, expendable cryogenic-liquid/solid chemical, resuable cryogenic chemical, reusable solar electric, reusable solid-core nuclear thermal, and expendable solar thermal. Because this is the Atomic Rocket website, I am going to focus on that. Details about the others can be found in the report.

The report states that the nuclear thermal rocket was initially eliminated due to having too many negatives in the scoring. However "The advanced nuclear systems scored very low, but at the request of some team members that insisted past studies showed this concept to be viable and should be investigated further, the advanced nuclear concepts were also advanced to the next phase." Translation: some of the team members were nuke fans and begged to let the nuclear thermal rocket pass.


Ground Rules:

  • Resuable launch vehicles deliver payloads to LEO
  • LEO is defined as a circular orbit with an altitude of 185 km (100 nautical miles) with an inclination of 28.5° (due to the unfortunate location of the Kennedy Space Center).
  • The In-space transportation system (ISTS) hauls the payload from LEO to GEO.
  • GEO is defines as a circular orbit with an altitude of 35,786 km (19,323 nmi) with an inclination of zero.
  • In-space transporation technology must be available at NASA technology readiness level of 6 or higher by year 2005.
  • For this study payload masses are 1,400 and 4,500 kg
  • A single RLV launch transports 11,000 kg to and from LEO. LEO transportation weight is defined as LEO delivery weight plus associated airborne support equipment (ASE) weight.
  • Cost for ground to LEO with RLV is $440/kg
  • ISTS will be serviced by the RLV. So a resuable ISTS may need two RLV flights: one to carry ISTS propellant, one to carry payload.
  • If the ISTS can only deliver payload to geosynchroneous transfer orbits (GTO), an apogee kick motor can be used to insert payload into GEO.
  • GTO is defined as an elliptical orbit with a periapsis of 185 km (LEO), an apoapsis of 35,786 km (GEO), and an inclination of 28.5°. Obviously.

NUCLEAR THERMAL IN SPACE TRANSPORT

AIST-NTR
EngineSolid core NTR
Thrust67,000 N
Specific
Impulse
900 s
Propellant
Mass Flow
7.6 kg/s
PropellantLH2
Engine Mass2,450 kg

This is one of the high-thrust systems, especially compared to the solar electric. So the payload will be delivered quite rapidly.

The estimated operating life of the engine is 36,000 seconds (ten hours) total. The report notes that the ten hour operating life is several times that predicted for the cryogenic chemical engine, and they suspect optimism on the part of the nuclear propulsion specialists.

For the 1,400 kg payload this will allow the rocket to perform 50 missions (I calculate roughly 720 seconds of engine life used per mission). The report says a 374 second burn is used to travel from LEO to GTO. After ejecting the payload with the AKM, the rocket does a 203 second burn to return to LEO (and perform a small plane change maneuver to correct for differential nodal regression). Following each burn, the upper stage shuts down the nuclear reactor, but continues to flow fuel (4 percent of that burned) for several minutes to cool the engine.

The 4,500 kg payload would restrict the rocket to 32 missions (I calculate roughly 1,125 seconds of engine life used). The report says a 695 second burn moves to GTO and a 248 second burn returns to LEO.

The engine is capable of 67,000 newtons of thrust. The design goal was only for an initial thrust-to-weight ratio of about 0.2 This would only require about 11,000 N for the 1,400 kg payload mission and only 22,000 N for the 4,500 kg payload mission. Sadly the study decided that downsizing the engine would not reduce the cost very much, since there is a minimum size set by need to have a critical mass of nuclear fuel.

A quick analysis indicates that to get the payload from GTO to GEO it is optimal to use an apogee kick motor (AKM) instead of adding extra propellant mass. Eliminating the AKM would require doubling the propellant mass, increasing the number of RLV resupply flights.

Both of the items below are designed to be boosted into LEO by the reusable launch vehicle.

The first is the NTR transport vehicle, fully loaded with payload and propellant. It delivers the payload into GTO, where the apogee kick motor part of the payload inserts the customer payload into its slot in GEO. The empty NTR transport vehicle uses the remainder of its propellant for the return to LEO. There it enters sleep mode and awaits its next mission. Remember the transport cannot land back on Terra. When a fresh Refuel/Resupply package arrives, the transport will expend 100 m/s to rendevous with it.

The Refuel/Resupply Package gives an empty transport all it needs to perform a new mission. It has a new customer payload with a fully fueled AKM, replacement parts, and a refill for the transport's propellant tanks. The radioactive fuel elements inside the nuclear reactor are good for 32 to 50 missions, so they do not need to be replaced. Once they are spent the entire transport is decommissioned by being sent into a "grave-yard orbit" somewhere between LEO and GEO. Replacing reactor fuel elements is a nightmare on the ground, trying to do this in orbit is just too dangerous.

ASE is "Airborne Support Equipment". This is the struts and fittings required to hold the transport or resupply package in the RLV, and to safely eject it from the RLV's cargo bay or whatever. The ASE mass is estimated to be 15% of the item mass. Example: if the transport has a mass of 12,377 kg, the ASE will be an additional 1,857 kg of struts and fittings.

Avionics-C&DH is command and data handling. Avionics-GN&C is guidance, navigation, and control.

NTR Transport Mass Budget
SystemSmall
Payload
(kg)
Large
Payload
(kg)
INERT WEIGHT SCHEDULE
Structure1,7642,641
Mechanism1551
Passive thermal control202279
Avionics-Power136136
Avionics-C&DH8383
Avionics-GN&C7373
RCS101108
Propulsion subsystem7979
Nuclear Rocket Engine2,4542,454
TOTAL STAGE INERT WEIGHT4,9075,904
PAYLOAD WEIGHT SCHEDULE
Customer Payload1,3614,536
Apogee Kick motor91302
AKM Propellant1,2054,016
TOTAL PAYLOAD WEIGHT1,2954,318
TRANSPORT WEIGHT SCHEDULE
TOTAL STAGE INERT WEIGHT4,9075,904
TOTAL PAYLOAD WEIGHT4,9075,904
TOTAL DRY MASS7,56414,758
Stage Fuel4,8147,830
TOTAL WET MASS12,37722,588
TOTAL LEO DELIVERY WEIGHT12,37722,588
Stage Delivery ASE weigh1,8573,388
TOTAL LEO TRANSPORT WEIGHT14,23425,977
PERFORMANCE
Mass Ratio1.6361.531
Exhaust Velocity8,829 m/s8,829 m/s
delta V4,348 m/s3,758 m/s
Refuel/Resupply Package
SystemSmall
Payload
(kg)
Large
Payload
(kg)
INERT WEIGHT SCHEDULE
structure1,4472,219
mechanism1551
Passive thermal control202279
Propulsion subsystem7979
TOTAL STAGE INERT WEIGHT1,7442,628
PAYLOAD WEIGHT SCHEDULE
Apogee Kick motor91302
AKM Propellant1,2054,016
Resupply Fuel Weight4,8147,830
Replacement Parts Weight78109
Customer Payload weight1,3614,536
TOTAL PAYLOAD WEIGHT1,2954,318
TRANSPORT WEIGHT SCHEDULE
TOTAL STAGE INERT WEIGHT1,7442,628
TOTAL PAYLOAD WEIGHT1,7442,628
TOTAL DRY MASS9,29119,421
TOTAL LEO DELIVERY WEIGHT9,29119,421
Stage Delivery ASE weight1,3942,913
TOTAL LEO TRANSPORT WEIGHT10,6822,334

Antares Dawn Battlecruiser

Battlecruiser Discovery
Engine
EnginePhoton drive
(with gears)
ΔV10,500,000 m/s
(10,500 km/s)
Thrust Power2.36×1014 W
(236 terawatts)
Photon
Power req.
4.71×1014
(471 terawatts)
Powerfusion
(deuterium
+ hydrogen)
Fusion fuel
burn rate
0.73 kg/sec
Engine High Gear
Initial Accel9.81 m/s2 (1 g)
Thrust1,570,000 N
Exhaust
Vel
3×108 m/s
Specific
Impulse
30,600,000 sec
Engine Low Gear
Initial
Accel
39 m/s2
(4 g)
Thrust6,240,000 N
Exhaust
Vel
75,500,000 m/s
Specific
Impulse
7,690,000 sec
Ship
Length110 m
Body Dia14.5 m
Centrifuge
major radius
22 m
Centrifuge
minor radius
4 m
Centrifuge
volume
6,950 m3
Centrifuge
spin
1.0 g: 6.4 RPM
0.5 g: 4.5 RPM
0.1 g: 2.0 RPM
Centrifuge
type
dependant
Ship volume≅32,000 m3
Ship density≅5 kg/m3
Ship wet mass≅160,000 kg
Parasite
craft
x4 armed scouts
Weaponsantimatter proj
particle beam
missiles
lasers
FTL energy10% fuel/jump

The Derringer-class heavy battlecruiser Discovery is from Antares Dawn by Michael McCollum. Yes, the spacecraft has a hand-waving faster-than-light drive but the rest of the details are impressively hard. This might have something to do with the fact that Mr. McCollum has a major in aerospace propulsion and a minor in nuclear engineering. He work on the precursor to the Space Shuttle main engine.

One of my preferences for including a given spacecraft in the Realistic Designs pages is that I can calculate the ship's delta-V. For the Discovery, I did not have to calculate it, it is actually given in the novel.


Having said that, understand that this thing is a freaking torchship. Both the thrust and delta V are outrageous.


At the start of the novel, the Battlecruiser Discovery is in a 1,000 km orbit around the planet Alta with full fuel tanks. To everybody's surprise, a large starship appears at the star system's sole jump point and takes off accelerating at one half gee heading away from Alta. Everybody is surprised because the jump point vanished 120 years ago, and nobody knew it had reappeared. This is linked to the Antares supernova, but I digress.

The Discovery is dispatched to intercept the large starship. This will be a challenge since the jump point is 250 million kilometers away from Alta and the large starship is showing no sign of stopping its burn. The Discovery has a total delta V of 10,550,000 m/s (10,500 km/s) so things are going to be tight. They don't realize it yet but the large ship is a full blown Blastship, and it has an order of magnitude more delta V.

     000h: Blastship appears 250 million km from Alta. Blastship velocity is 0 km/s

     022h: Discovery departs Alta to intercept blastship. 10,500 km/s ΔV in tanks. Starts Burn 1 (33 hours at 3.5g). Blastship velocity is 388 km/s

     055h: End of Burn 1. 4,079 km/s ΔV expended, 6,421 km/s ΔV left in tanks. Discovery does skew-flip and starts deceleration Burn 2 (21 hours at 3.5 g). Blastship velocity is 970 km/s

     076h: End of Burn2. 2,596 km/s ΔV expended, 3,825 km/s ΔV left in tanks. Discovery rendezvous with blastship. Both velocity are 1,300 km/s. Discovery matches blastship acceleration of 0.5g. Discovery can do this for only 12 hours before it has to abandon the chase or not have enough fuel to return to Alta.

     084h: Discovery has 4 hours before forced to abandon chase. Both velocity are 1,480 km/s. Blastship's fuel tanks are identified by thermal imaging. Discovery punctures all six fuel tanks using secondary laser weapons.

     085h: Discovery has 3 hours before forced to abandon chase. Both velocity are 1,500 km/s. Blastship's fuel tanks finally run empty through punctures and blastship stops accelerating, as does Discovery. 159 km/s ΔV expended, 3,666 km/s ΔV left in tanks.

     253h: The blastship turns out to have a dead crew, lots of battle damage, and is running on autopilot. After a week of studying the blastship, Discovery receives a recall message from home base. Blastship will be intercepted later by a tanker and repair ship. Both ships have a velocity of 1,500 km/s and are 1.5 billion kilometers from Alta. Start of deceleration Burn 3 (21 hours at 2g).

     274h: End of Burn 3. 1,483 km/s ΔV expended, 2,183 km/s ΔV left in tanks. Discovery has a velocity of 0 km/s. Start of homeward Burn 4 (14 hours at 2g)

     288h: End of Burn 4. 989 km/s (book says 1000 km/s) ΔV expended, 1,194 km/s ΔV left in tanks. Discovery has a velocity of 1000 km/s. Start of 17 day coast phase.

     689h: End of coast phase. Discovery still has a velocity of 1000 km/s. Start of braking Burn 5 (14 hours at 2 g)

     703hh: End of Burn 5. 989 km/s (book says 1000 km/s) ΔV expended, 205 km/s ΔV left in tanks. Discovery has a practical velocity of 0 km/s in Alta orbit with only 2% of its original fuel load.

ANTARES DAWN

The landing boat overtook Discovery from below and behind, giving Drake a good look at his ship. The battle cruiser consisted of a torpedo-like central cylinder surrounded by a ring structure. The central cylinder housed the ship’s mass converter, photon drive, and jump engines — the latter needing only an up-to-date jump program to once more hurl the ship into the interstellar spacelanes. In addition, within the cylinder were fuel tanks filled with deuterium and tritium enriched cryogen; the heavy antimatter projectors that were Discovery’s main armament; and the ancillary equipment that provided power to the ship’s outer ring.

The surrounding ring was supported off the cylinder by twelve hollow spokes — six forward and six aft. It contained crew quarters, communications, sensors, secondary weapons pods, cargo spaces, and the hangar bay in which auxiliary craft were housed.

Unlike the interplanetary vessels built during the years of isolation, which all tended to be haphazard collections of geometric shapes, the battle cruiser’s shape was streamlined. Its sleek form was more concerned with the need to keep the jump charge from bleeding off the hull before a foldspace transition than to any requirement for the ship to transit a planetary atmosphere.

Drake listened to the communications between the landing boat and the cruiser all through the approach. As they drew close, he noticed the actinic light of the ship’s attitude jets firing around the periphery of the habitat ring. When in parking orbit, the cruiser was spun about its axis to provide half a standard gravity on the outermost crew deck. The purpose of the attitude jets was to halt the rotation in preparation for taking the landing boat aboard.

Drake was well pleased with what he heard on the intercom during the approach — mostly silence punctuated by a few terse exchanges of information. The complete absence of chatter was evidence of a taut ship and a good crew. He was suffused with a warm feeling of pride as he watched hangar doors (on ship's nose) open directly in front of the hovering boat just as the cruiser’s spin came to a halt.

     “Landing Boat Moliere. You may secure your reaction jets!” came the order from Discovery approach control.
     “Securing now,” the pilot said as he reached down to throw a large, red switch next to his right knee. The message ‘REAC JET SAFE’ flashed on a screen on the control panel.
     “Prepare to be winched aboard.”
     “Hook extended.”

A torpedo-like mechanism exited the open hatch and jetted across the dozen meters of open space to where the landing boat hovered. Attached to the torpedo was a single cable. The torpedo disappeared from view for several seconds, then the approach controller said, “All right, Moliere. Stand by to be reeled in!”

There was a barely perceptible jolt as the cable took up slack, then the landing boat slid smoothly forward. The curved hull of the cruiser and the open maw of the vehicle hatch swelled to fill the windscreen. The boat passed out of Val’s direct rays and into shadow. The dark was short lived, however. As soon as the bow passed into the hangar bay, the windscreen fluoresced with the blue-white glow of a dozen polyarc flood lamps.

There was a harder bumping sensation as the bow contacted the recoil snubber inside the bay. Then the boat was being pulled completely inside by giant manipulators and lifted to its docking area while a steady stream of orders issued from the bulkhead speaker.

“Close outer doors. Stand by to repressurize.”


There is a common belief among the uninitiated that a spaceship’s control room is located somewhere near the ship’s bow. In truth, that is almost never the case. Discovery, with its cylinder-and-ring design, was particularly unsuited to such an arrangement. Like most warships, the cruiser’s control room was located in the safest place the designers could find to put it — at the midpoint of the inside curve of the habitat ring.

Actually, Discovery possessed three control rooms, each capable of flying or fighting the ship alone should the need arise. For normal operations, however, there was a traditional division of labor between the three nerve centers. Control Room No. 1 performed the usual functions of a spacecraft’s bridge (flight control, communications, and astrogation); No. 2 was devoted to control of weapons and sensors; and No. 3 was used by the engineering department to monitor the overall health of the ship and its power-and-drive system.


An auxiliary screen lit up as a camera mounted on the habitat ring caught the glow that suddenly erupted from the aft end of Discovery’s central spire. Theoretically, the cruiser’s photon drive should have been invisible in the vacuum of space. However, waste plasma from the ship’s mass converters was dumped into the exhaust (gear-shifting the drive into low gear), causing the drive plume to glow with purple-white brilliance as Discovery broke from her parking orbit and headed out into the blackness of deep space.


An hour later, the ship was accelerating along a normal departure orbit at one standard gravity while crewmen rushed to convert compartments from the “out is down” orientation of parking orbit, to the “aft is down” of powered boost. The only compartments that did not need conversion were the control rooms (which were gimbaled to automatically keep the deck horizontal) and the larger compartments (hangar bay, engine room), which had been designed to allow access regardless of the direction of “down.”

From ANTARES DAWN by Michael McCollum (1986)
ANTARES PASSAGE

At the word “zero,” the apparition dramatically changed appearance.  Suddenly, the mirror-sheen (of the anti-radiation protective shield) was gone and a hull of armored steel took its place.  The ship thus revealed was a twin of Discovery.  Its central cylinder jutted from the center of a habitat ring.  Twelve spokes joined the central cylinder to the ring.  A focusing mechanism for the ship’s fusion powered photon engines jutted from the back of the central cylinder, while the business ends of lasers, particle beams, and antimatter projectors jutted from various places on the hull.  The outlines of hatches marked the positions of internal cargo spaces and hangar bays in which auxiliary craft were housed.

The Derringer-class heavy battle cruiser was a design that went back nearly two centuries.  Designed for speed and acceleration, the ring-and-cylinder design was a compromise between a good thrust-to-mass ratio and an adequate low speed spin-gravity capability.  The design was ungainly and fragile looking, but proven in battle.  One advantage the cylinder-and-ring ships had over purely cylindrical designs, if a ship were severely damaged, the habitat ring could be jettisoned whole, or in as many as six separate pieces.


Ten minutes after departing City of Alexandria, Landing Boat Moliere drew abreast of His Majesty’s Blastship Royal Avenger.  The view through the starboard viewports was awesome.  At the blastship’s stern were the focusing rings and field generators of three large photon engines.  Even quiescent, the engines that drove the flagship gave the impression of unlimited power.  Just in front of the engine exhausts were the radiators and other piping associated with the ship’s four massive fusion generators.  In front of the generators were the blastship’s fuel tanks; heavily armored and insulated to keep the deuterium enriched hydrogen fuel as close to absolute zero as possible.

Drake let his gaze move forward along the blastship’s flank.  The cylindrical hull was pierced in places by large hangar doors through which armed auxiliaries could sortie into battle.  Forward of these were the snouts of a dozen antimatter projectors, Royal Avenger’s primary anti-ship weapons.  The business ends of other weapon systems also jutted from the heavily armored hull.  Interspersed with the weaponry were all manner of sensor gear.

As the landing boat slipped past the blastship’s flanks, they were rewarded with ever changing vistas since Avenger was rotating about its axis at the rate of several revolutions per minute.  So close was landing boat to blastship that it was easy to imagine oneself in a small aircraft flying over an endless plain.  The optical illusion came to an abrupt end when the landing boat passed abeam of the blastship’s prow.

Like most starships, little or no effort had gone into streamlining Avenger.  In fact, the prow was actually slightly concave, and its surface covered with arrays of electronic and electromagnetic sensors.  A hangar door outwardly identical to those that dotted the blastship’s flanks was set flush with the hull at the giant ship’s axis of rotation.

As quickly as the bow portal came into view, Moliere’s pilot fired the attitude control thrusters to halt the landing boat’s forward speed.  Once Moliere had halted in space, he began firing his side thrusters to align the landing boat with the central portal.  A popping noise echoed through the passenger cabin each time the thrusters fired.  When Moliere was lined up with Royal Avenger’s axis portal, the thrusters fired twice more to match the flagship’s rate of rotation.  The hangar door retracted, and Moliere’s pilot nudged his boat toward the lighted opening.  Within seconds, the boat passed into a spacious cavern lighted by million-candlepower polyarc lamps.  There followed a series of bumping and scraping noises, and a gentle tug of deceleration as the landing boat’s forward velocity was halted.  After that, there came a long span of silence interrupted by the sudden sound of air swirling outside the hull.

Moliere had arrived.

From ANTARES PASSAGE by Michael McCollum (1998)

Artemis 8

The following memo was sent by the author to NASA administrator Jim Bridenstine and Scott Pace, executive secretary of the National Space Council, on June 30, 2020.

A mission equivalent to Apollo 8—call it “Artemis 8”—could be done, potentially as soon as this year, using Dragon, Falcon Heavy, and Falcon 9.

The basic plan is to launch a crew to low Earth orbit in Dragon using a Falcon 9. Then launch a Falcon Heavy, and rendezvous in LEO with its upper stage, which will still contain plenty of propellant. The Falcon Heavy upper stage is then used to send the Dragon on Trans Lunar Injection (TLI), and potentially Lunar Orbit Capture (LOC) and Trans Earth Injection (TEI) as well.

There are two options for how to do it:

  • A. Do mission only using the Dragon and the Falcon Heavy upper stage as flight elements, with the Falcon Heavy upper stage doing all maneuvers, as described above.
  • B. Do the mission using the Dragon, the Falcon Heavy upper stage for TLI, and a small propulsion stage (SPS) lifted to orbit by the Falcon Heavy upper stage for LOC and TEI.

Assumptions:

TLI ΔV = 3.1 km/s
LOC and TLI ΔVs = 1 km/s each for capture into Low Lunar Orbit, but less for capture into higher lunar orbits.
Dragon mass = 9.5 metric tons
FH upper stage dry mass = 10 tons
FH upper stage propellant capacity = 109 tons
FH engine specific impulse (Isp) = 348 s = 3.41 km/s exhaust velocity
SPS engine (Isp) = 378 s (LOX/CH4) = 3.7 km/s exhaust velocity
FH upper stage mass on reaching LEO = 75 tons = 10 ton dry mass + payload, with rest residual propellant. (This number results directly out of SpaceX data that its payload to LEO is 65 tons, and its payload to GTO is 26 tons.)

Option A

Falcon Heavy is launched without payload, resulting in LEO mass of the 10-ton dry stage and 65 tons propellant. After rendezvous and mate with Dragon, the assembled spacecraft has a dry mass of 19.5 tons and 65 tons of propellant. So mass ratio is 84.5/19.5 = 4.33. With the Falcon Heavy exhaust velocity of 3.41 km/s, this translates into a total ΔV capability of 5.0 kilometers per second. After 3.1 km/s used for TLI, this leaves 1.9 km/s for two 0.95 km/s ΔVs for LOC and TEI, enabling capture into a “lowish” lunar orbit and return to Earth.

Option B

The Falcon Heavy is launched with SPS as payload. The SPS includes 7.9 tons of LOX/CH4 propellant and 1.5 tons of dry mass. Together with Dragon, it has a total mass of 18.9 tons. With a mass ratio of 18.9/11 = 1.717 it has a total ΔV capability of 2 km/s, allowing it to do LOC and TEI going into and coming back from low lunar orbit. The Falcon Heavy upper stage reaches LEO with the 10-ton dry mass Falcon Heavy upper stage, 9.4 tons SPS mass, and 55.6 tons of propellant. After rendezvous and mate with Dragon, the assembled spacecraft will have a total mass of 84.5 tons, with 55.6 tons of that available in the FH upper stage to perform TLI (the remaining maneuvers will be done by the SPS).

The mass ratio of this assembly with respect to the TLI burn is 84.5/28.9 = 2.92. With the Falcon Heavy upper stage exhaust velocity of 3.41 km/s, this means that the Falcon Heavy upper stage will be able to executive a ΔV of 3.65 km/s, or 0.55 km/s more than the 3.1 km/s required. So, there is plenty of margin in this design, and in fact lower-performing propellants such as LOX/RP (348 s Isp) or NTO/MMH (320 s Isp) could be employed in the SPS and the mission would still be feasible as described, with the only penalty being a modest reduction in margin.

Life support and reentry issues

Travel to the Moon and back requires a minimum of six days, and the Dragon should be good for that. However, let’s say we want to add ten extra days to the Dragon’s endurance. A crew member uses one kilogram of oxygen per day. Thus, with a crew of two, ten days would require transporting an extra 20 kilograms of oxygen. If stored in gas cylinders at 3000 psi (as is done in SCUBA tanks), this would require a total volume of 0.075 cubic meters. The Dragon’s internal volume is 9.3 cubic meters, so that less than 1% of the available volume would be required to accommodate such tankage.

Dragon’s thermal protection is designed for reentry from return from Mars. This is a higher thermal protection requirement than return from the Moon.

Other observations

The Artemis program is advancing too slowly. As matters currently stand, it will have no visible accomplishments by the time of the election. Should administrations change, there is an excellent chance it will be cancelled. The Nixon Administration was not sympathetic to NASA’s plans for the human exploration of the Moon and Mars, and in fact cancelled NASA’s post-Apollo Moon base and Mars mission plans. But after Apollo 8, the only actual Moon mission done while LBJ was still president, it became unthinkable to abort the Apollo program short of landing. The best defense that Artemis will have in the event of a change of administrations is real tangible accomplishment: either actually done, or at least clearly imminent. Otherwise, it will be orphaned and likely go the way of Constellation and SEI. This must be prevented.

With Artemis 8, NASA can inspire the nation, restore our space program’s can-do spirit, and astonish the world with what free people can do. We should not miss this chance.

From “ARTEMIS 8” USING DRAGON by Robert Zubrin (2020)

Asaph-1 Mission to Phobos

This is from A Design Proposal for Asaph-1: A Human Mission to Phobos (2014).

ASAPH-1

I. Introduction

     THE moons of Mars are an excellent option for human exploration prior to the exploration of Mars. The moons provide a test-bed for many essential technologies that are required for a manned mission to Mars, while removing some of the complex issues that also must be addressed, such as Martian atmospheric entry of very large payloads and the prevention of forward contamination. Further, the moons are a good place to investigate the potential for insitu resource utilization (ISRU), which is an essential element for long-duration missions and possible colonization of Mars. Aside from these advantages, the moons also offer the unique opportunity to study asteroid-like small bodies in the solar system without having to undertake the risk of going into the asteroid belt itself. The study of small bodies will help in answering important questions about the formation of the solar system and the presence of life on other planets. A human mission to these moons will enable the performance of in-situ studies and also the return of samples to Earth, which can be analyzed with all the resources we have at hand without the constraints introduced by deep space operations.

     Due to its larger size and interesting surface morphology, including the presence of numerous craters and at least one large monolith, we believe that exploring Phobos offers the greatest scientific returns for a given cost. Nevertheless, a concurrent study of Deimos’ composition and structure via remote and/or robotic experimentation will provide vital information about the differences between the moons and may shed additional light on the formation of the moons.

     Motivated by these points, the Asaph-1 mission (named for Asaph Hall, the discoverer of the Martian moons in 1877), a manned mission to Phobos, was proposed by this 16-member “Team Voyager” as part of the Caltech Space Challenge held March 25-29, 2013, at the California Institute of Technology, Pasadena, California, USA. The mandate from the senior scientists, engineers, and organizers to the students was to design a manned mission to one of the Martian moons with a launch date no later than January 1, 2041. During the workshop, Team Voyager divided into subsets of student-experts to address such considerations as science objectives, remote-sensing instrumentation, trajectory, propulsion, communications, habitation design, human health, sample return, biologic contamination, and risk. The group arrived at a consensus on key design items by, first, discussing their merits with scientists and engineers from JPL, NASA, Lockheed Martin, and SpaceX; second, voting on them as a group; and third, affixing them to our “wall of truth.” Once a design consideration reached the wall of truth, it became a permanent part of the mission plan. This paper is a summary of the results of our detailed mission plan, including: (1) scientific motivation for the mission, (2) a summary of the mission architecture, (3) first-order details of the mission, such as trajectory design, propulsion systems and habitat design, and (4) a brief discussion of the long term impact of such a mission. Owing to the condensed, intense nature of the workshop some contingencies and peculiarities of the mission, such as abort trajectories, alternative lower-ΔV trajectories, and multiple re-entry scenarios (i.e., aerobraking and aerocapture maneuvers), could not be evaluated.

II. Scientific Motivation for the Mission

     The Asaph-1 mission is motivated by scientific discovery and demonstration of novel technologies, including those needed to support the extended duration of humans in space. Several outstanding physical and biological science questions that may be answered by the mission include:

     (1) What are the compositions, ages, and origins of Phobos and Deimos?

     Phobos is the larger, closer moon with approximate dimensions of 26.8 x 22.4 x 18.4 km. Deimos is the smaller, more distant moon with approximate dimensions of 15 x 12.2 x 10.4 km. To date, only a limited amount of visible imagery and infrared spectroscopic data has been acquired to determine the compositions of either of the moons, which, at least at the surface, consist of phyllosilicates (serpentine and/or kaolinite) with lesser feldspars or feldspathoids. The ages and origins of the moons are unknown. Both moons are very similar in composition to C- and D-type asteroids, which leads to the hypothesis that they are captured asteroids. However, they both have nearly circular and equatorial orbits around Mars, which would necessitate an explanation for the circularization and adjustment of the inclination of their orbits after capture. Additional hypotheses for their origin(s) are: (1) They are remnant debris left over from the Martian accretionary process, (2) They are second generation solar system objects that coalesced in orbit after Mars formed, (3) They are two of many small bodies that were ejected from the Martian surface by collision with a large bolide, (4) They are captured cometary nuclei. Measuring radiometric ages on the moons will help to constrain the formational history of the moons and, by extension, Mars itself.

     (2) Are there any compounds—particularly water, hydrocarbons, or metals—on Phobos or Deimos that could be used for human habitation in space (e.g., to establish a station on one of the moons)?

     Current data suggest that there is no free water (ice) on the surface of the moons. To date, all ‘water’ observed by spectroscopy occurs in hydroxyl groups bound within phyllosilicate minerals. A temperature on the order of ~500°C is required to dehydroxylate phyllosilicates, thereby liberating free water. Such a process may represent an engineering challenge but does not preclude the use of phyllosilicates as a source of water. The estimated densities of Phobos and Deimos are 1.87 and 1.54 g/cm3, respectively. A back-of-the-envelope average of seven common phyllosilicate minerals yields a density of ~2.61 g/cm3. Because the bulk density of the moons is significantly less than the average density of common phyllosilicate minerals, there must be a significant amount of lower density material present within the moons, i.e., various ices or potentially clathrate-like combinations of light hydrocarbons and water. If clathrates were to be found, they could prove useful for human habitation and transportation. The low bulk density of the moons argues against the presence of significant quantities of metals.

     (3) Are there any compounds that may indicate the presence of life?

     As yet, the answer to this question is unknown. Based on the criteria discussed by Clark et al. for small bodies within 2 A.U. of the Sun, it is likely that the Martian moons are sterile. Nevertheless, samples returned from the moons should be carefully shielded from organic contaminants, as these samples may yield important data to help answer this question.

     (4) What are the surface characteristics of the Martian moons, especially with regard to landing a spacecraft there?

     Images captured by MRO suggest that craters on both Phobos and Deimos are partly to completely filled with what appears to be powdery, fine-grained regolith. The craters of Deimos appear to be more filled with powder than those on Phobos. It is critical to know the depth and nature of the powdery regolith in order to make informed decisions about landing a spacecraft on either of the moons.

     (5) What physiological and psychological anomalies can be characterized using scans and samples from our crew during their incursion into deep space?

     Pre-, mid-, and post-mission analyses of crew health indicators will clarify the effects of radiation exposure, extended mission stress, and other, possibly unforeseen, factors on humans.

     (6) What will be the profile of radiation exposure encountered during the mission?

     Radiation data from the Mars Science Laboratory (MSL) cruise phase and the Asaph-1 precursor mission will better define the quantity and intensity of radiation that the crew must endure during the Asaph-1 manned mission to Phobos and, eventually, during the first human mission to the surface of Mars.

III. Mission Architecture

     This section details the mission architecture intended for the Asaph-1 mission, including: the benefits of implementing a precursor mission for such a program, the over-arching mission structure, a general timeline to achieve the scientific and operational goals, and other important engineering considerations, such as technological considerations and strategic knowledge gaps.

A. Phase One: Precursor Mission, Motivation and Benefits

     Just as the Surveyor program evaluated landing sites for the Apollo missions, a robotic precursor mission to Phobos and Deimos will reduce the risks involved in a manned mission by surveying potential landing sites and demonstrating technological feasibility. Phase One of the mission consists of an orbiting, remote-sensing Phobos-Deimos Surveyor (PDS), an impactor-lander Phobos Explorer (PE) and an identical impactor-lander Deimos Explorer (DE). The PDS-PE-DE system (Fig. 1) will launch from Earth in 2026 in a Falcon 9 and will use solar-electric propulsion to spiral out to Mars slowly over the course of two years. Upon reaching Mars in 2028, the PDS system will survey Phobos, its primary objective, and then Deimos and will deploy the PE and DE packages near their respective landing points. Having completed those missions, the PDS will remain in orbit around Mars to act as a communications relay for the Phase Two manned mission.

     At Mars, the PDS-PE-DE system will enter an areocentric orbit below Phobos with an inclination of 20°. This orbit will cause the PDS-PE-DE system to gradually overtake Phobos, giving surveillance coverage of both the north and south pole regions. Then the surveyor will transition to an orbit above Phobos, which will allow for the mapping of over 80% of the moon’s surface. From this higher vantage point, the PDS will release the PE, which contains an impactor experiment and lander. The impactor package will release four penetrometers to strike widely-spaced sites on each moon (Fig. 2). Using the results from the impactor experiment and the orbiting surveyor, the PE lander will reconnoiter the site most suitable for the landing of the manned mission, with sites ‘A’ and ‘B’ being the priority sites.

     The PDS system will then enter a higher-altitude Mars orbit, just below Deimos, and will release the DE. Again, this orbit will be slightly inclined from the ecliptic. The PDS will slowly move from below Deimos to trailing it, and then to a higher altitude orbit, thus mapping up to 50% of the moon’s surface. As on Phobos, the DE will release four penetrometers, which will be viewed from this higher altitude. Immediate results from the impacts will determine the landing site for the DE lander. Once the impact experiments have been performed, the PDS will move back down to an areocentric orbit slightly below Phobos, thereby maintaining sufficient communications with both landers. Because both moons are tidally locked to Mars, all of the impactor sites on Deimos, and all but site ‘B’, within Stickney crater on Phobos, have full view of the Martian surface at all times3. Although the explorer packages will necessarily be highly autonomous, this will allow windows for the explorers to communicate information to the orbiting PDS system.

     Although the movements to raise and lower the PDS system do complicate the precursor mission plan, the movement is required in order to survey both moons with the understanding that Phobos is the principal target of interest. Because Phobos is the priority, if the PE were to fail to initialize or if it yielded unsatisfactory results, the DE could be substituted for the faulty PE. If this were to be the case, the secondary raise to Deimos’ orbit would be obviated.

     Both landers will collect scientific data over the course of several years, until their power supplies run out. Meanwhile, the orbiting PDS will make remote sensing observations before, during, and probably after the Phase Two manned mission, and will also act as a key communications relay during Phase Two activities.

B. Phase Two: Primary Mission Overview

     Phase Two, the manned mission, is planned to be an operation with a human crew in which surface operations, including sample collecting, will be conducted on a Martian moon, nominally Phobos, depending on favorable results from the Precursor mission. The crew is anticipated to return to Earth with geological samples and other data collected at the surface. A human crew was chosen to carry out sample collection and operations of this mission, as opposed to a teleoperated robotic system, because human astronauts on the ground are uniquely suited to make rapid decisions about geologic sample collection, and they possess a situational awareness necessary to meet mission goals at Phobos.

     Phase Two is achieved using a sequence of launches from Earth to LEO, where the modules will rendezvous to form the mothership (MS). Once the assembly is complete, the MS will use an impulsive propulsion maneuver to reach the Martian system within six months. At Mars, the MS will enter a parking orbit for approximately one month. During this period, the Space Exploration Vehicle (SEV) will approach the surface of Phobos to perform scientific activities. After returning the crew to the MS, the SEV will return again to the surface of Phobos as a probe to continue autonomous scientific operations over the course of several years. The rest of the MS will depart from Mars using another impulsive propulsion maneuver to return to Earth. A bat chart of the mission (Fig. 3) is provided for easier visualization of the primary mission.

     The primary mission will utilize the following modules: (1) Propulsion Systems 1 and 2 (PROP1 and PROP2), which contains a nuclear thermal propulsion (NTP) system including liquid hydrogen tanks, (2) SEV, a vehicle that will bring astronauts from the MS to Phobos proximity and back, (3) Deep Space Habitat (DSH), a module that provides additional habitable volume for the crew. (4) Multi-purpose Command Vehicle (MPCV) with Orion Crew Module (CM), a vehicle that serves as the habitable volume shuttle from Earth to the MS, and will be used for the reentry of the crew.

C. Primary Mission Timeline and Considerations

     The primary mission has a nominal duration of 465 days, including a 185-day-outbound transfer, a 30-day stay at Mars and a 250-day-inbound transfer. The crew will leave Earth’s orbit in April, 2033, arrive at Mars during October of the same year, and return back to Earth in July, 2034.

     The determination of the key dates and trajectories for the mission is based on multiple factors. The first trade-off is between undertaking a short-stay (opposition class) mission versus a long-stay (conjunction-class) mission. Considering crew safety issues due to radiation exposure in deep space and taking into account that a longer round-trip duration will lead to a higher probability of contingencies, we prefer the opposition-class mission concept.

     The total ΔV from LEO to Mars, as a function of round-trip time and departure date during the ideal launch window, is plotted in Figure 4A. The investigated departure dates are a result of the time needed to develop the required technologies (leading to a highly optimistic early departure in 2020), with a launch date no later than January 1, 2041, as defined in the mission statement.

     Concerning radiation exposure, it is most favorable to perform a deep space mission during solar maximum. The first solar maximum within the shown departure dates will peak around 2022 (solar cycle 25); the following solar cycle 26 peaks between 2033 and 2035. Solar cycle 25 is predicted to be one of the weakest in centuries. Additionally, there are only a few possible launch dates in 2022 for an opposition-class mission. For these reasons, April, 2033 is selected for further investigation.

     The total ΔV has been calculated from LEO as a function of round-trip duration (Fig. 4B). It shows that a shorter round-trip duration automatically leads to an increase in the total ΔV required. It should be noted that the smallest ΔV , (i.e., longest round-trip duration) corresponds to the earliest departure date (April 7, 2033). With later departure dates, the round-trip duration decreases while ΔV increases. As a result, April 7, 2033 is determined to be the optimal departure date. This date selection allows for a launch slip of up to 25 days. Choosing this trajectory, there is calculated to be a constant line of sight from the spacecraft to Earth while in transit to and from Mars. Such a line of sight will be highly beneficial for flight control communications and crew safety.

     Trajectories were calculated using a robust Lambert solver, with ephemerides from JPL. At Mars, a bi-elliptic transfer is chosen to transport the crew safely from the mothership to Phobos. In theory, a Hohmann transfer would be more efficient to do this, where the red diamond marks the used transfer’s position on the graph (Fig. 4C). However, as Mars is only just beginning to capture the spacecraft when starting this transfer, the actual ΔV required to do a Hohmann transfer is a factor of ten larger than the ΔV using the bi-elliptic transfer.

D. Technological Requirements and Strategic Knowledge Gaps

     It is important to understand the technology required for the accomplishment of the mission. The technologies employed in the mission are currently at various readiness levels. Development time is taken into account in the mission architecture, and some of them will be discussed in detail in later sections of this report. A few of the key technological requirements are as follows: (1) A safe habitat needs to be designed for astronauts to survive for about 500 days in deep space. This includes radiation shielding, smart resource utilization, and comfortable living space for the astronauts. (2) Efficient propulsion systems that provide reasonable thrust at high Isp are required to transport the crew and supplies. (3) Multiple, carefully-timed launches are required to transport all the modules to the Martian system. (4) The capability to abort the mission safely at various stages needs to be assessed.

     In addition, there are certain important strategic knowledge gaps (SKG) that need to be retired before the undertaking of the manned phase of the mission: (1) The surface properties of Phobos and Deimos, such as regolith thickness and strength, are completely unknown and must be characterized before humans can be sent to either moon, (2) Deep space vehicles need to be tested for survival in deep space conditions prior to usage by astronauts, (3) Custom fairings need to be developed in order to accommodate high volume payloads on launch vehicles, (4) Methods for faster turnaround times for launch vehicles need to be developed in order to facilitate more launches in shorter periods of time, which allows for faster assembly of deep space cargo in LEO, (5) On-orbit assembly on a large scale needs to be perfected through research and testing, and (6) Improved thermal protection must be developed to protect spacecraft from the heat generated by reentry velocities in the range of 14-16 km/s.

IV. Details of the Phase Two Primary Mission

     In this section we present pertinent details about specific aspects of the primary mission, including: (1) trajectory, (2) propulsion and vehicle selection, (3) habitation design and considerations for human success in deep space, (4) surface mission operations that will realize the scientific goals of the mission, (5) systems engineering, (6) planetary protection, (7) risk matrices for the mission and program as a whole, and (8) anticipated costs and partnerships.

A. Trajectory

     The proposed trajectory (Fig. 5) is designed for an opposition-class mission with a round-trip duration of 465 days. Neither PROP1 nor PROP2 can be assembled and launched as a whole from Earth. Therefore, the launch campaign for the unmanned modules will start approximately five months prior to crew departure. Both PROP1 and PROP2 will each be brought into LEO through multiple launches over a period of several weeks. The design choice for LEO is further explained in Section IV.B.: Launch Vehicle Selection and Propulsion. After both modules have established a stable orbit of 300 km and are successfully assembled, the DSH will be launched to the same position and docked to the PROP1-PROP2 assembly. Only then, the crew, along with SEV, CM, and Service Module (SM), will be launched on April 7, 2033. The crew will enter LEO to rendezvous with the PROP1-PROP2-DSH assembly. Altogether, these six major components (PROP1-PROP2-DSH-SEV-CM-SM) comprise the MS, which will depart LEO later in April, 2033 for arrival at Mars in October, 2033. The MS will remain there for 30 days before beginning its return to Earth in November, 2033 with a planned Earth arrival in July, 2034. Figure 5A provides an overview of the heliocentric trajectories and the respective dates.

     Upon successful assembly of the MS, PROP1 and PROP 2 will provide a ΔV of 3.5 km/s in order to achieve a C3 energy of 6.15 km2/s2. This C3 will place the spacecraft on a hyperbolic trajectory for arrival at Mars on October 10, 2033. The Earth escape trajectory will have an outgoing asymptote right ascension of 272°, a declination of -23°, and a velocity azimuth at the periapsis of 90° in the Earth inertial reference frame.

     At Mars, the MS will burn with a ΔV of 2.2 km/s to achieve a Mars Orbit Insertion (MOI) and enter a 250 x 33,813 km parking orbit around Mars (orbital period of 1 sol) with an inclination of 34°. The eccentricity of this orbit (white dashed line in Figs. 5B, C) will aid in the transfer to Phobos’ orbit (dark blue line in Figs. 5B, C) by losing much of the velocity from the approach. What follows is the phasing period, which could require a minimum of twelve hours to a maximum of 14 days. Phasing ends once two criteria are fulfilled: (1) MS and Phobos have a phase difference of 180°, and (2) the first condition is met when the MS is located at the parking orbit apoapsis (Point 2 in Figs. 5B, C).

     As soon as these two criteria are met, the crew will board the SEV and depart for Phobos rendezvous. The desired orbit will be reached through a bi-elliptic Hohmann transfer with apse rotation requiring a ΔV of 0.4 km/s, which will change the SEV orbit inclination to 8° and raise the periapsis to 9377 km (light blue line in Figs. 5B, C). The periapsis will then match the radius of the Phobian orbit. After a 15-hour transfer, the SEV will perform a ΔV of -0.7 km/s to place the crew in a circular Phobos trailing orbit with an inclination of 1° (Point 3 in Figs. 5B, C). The SEV will trail Phobos for a minimal duration of 14 days. This duration can be increased if the initial MS-Phobos orbit phasing requires less than 14 days to complete. The SEV will visit several sites on the Phobian surface, which are described in Section IV. F: Science Mission and Surface Operations.

     Upon completion of all surface operations, the MS will exit the highly eccentric parking orbit and enter the Phobos trailing orbit of the SEV, requiring a total ΔV of 1.1 km/s. Docking of MS and SEV will occur on November 6, 2033. After the crew transfers from the SEV back to the MS, the SEV will return to the Phobian surface. It will use the same anchoring system previously used during EVA activities to attach itself to Phobos. The crew will continue to collect scientific data from Phobos using tele-robotic systems during their return to Earth.

     After two preparatory burns requiring a total ΔV of 0.7 km/s (Points 4 and 5 in Figs. 5B, C), a ΔV burn of 3.7 km/s will send the MS on a hyperbolic return trajectory on November 8, 2033. Arrival at Earth will be on July 16, 2034 with an Earth-relative velocity 16.2 km/s. An additional burn or aerobraking maneuver will reduce reentry velocity to approximately 14 km/s. A summary of the proposed trajectory with ΔV requirements can be found in Table 1.

Table 1. ΔV summary for MS and SEV mission operations
DescriptionΔV (km/s)
Place MS on hyperbolic trajectory3.5
Mars orbit insertion (MOI)2.2
SEV burn at apoapsis when Phobos- HEV phase difference
is 180° with plane change of 11.6° from ecliptic to 1.1°
with respect to Mars’ equatorial plane
0.4
SEV Phobos trailing orbit insertion for astronaut EVA0.7
MS departure from parking orbit0.4
Phobos trailing orbit insertion for MS0.7
Phobos trailing orbit exit when EVA is complete0.5
Burn at apoapsis to prepare for escape trajectory0.2
ΔV for Mars sphere of influence escape for return to Earth3.7
Total ΔV requirement for SEV1.1
Total ΔV requirement for MS11.2

B. Launch Vehicle Selection and Propulsion

3. Detailed Design

     The mass of the NTP module is approximated using the same assumptions as the NASA Human Spaceflight Architecture Team. The propulsion system comprises two different modules. The first module consists of the engine and nuclear core as well as some propellant. The second module is a tank carrying the bulk of the liquid hydrogen.

     The first stage consists of two engine cores generating a total thrust of 444 kN, which results in a thrust-to-weight ratio of 0.09. It is favorable to achieve a ratio of 0.1 for an impulsive burn, though in the case of starting from a circular orbit (LEO), it is not as critical as launching from an elliptical orbit. The burn duration is 82 min. The second stage (return trip) generates a thrust of 222 kN, resulting in ratio of 0.12 with a burn duration of 45 min. The elliptical orbit at Mars requires the increased ratio.

     The main propellant for the NTP is liquid hydrogen with a very low density of 70.85 kg/m3. To be able to exploit the full launch mass capacity, modifications to the fairing diameter, as well as length, are required. A simple increase in the diameter has significant implications for launcher performance. In order to increase the payload volume while still meeting the structural and control requirements, a shroud optimized for aerodynamics is proposed (Fig. 6). The optimized configuration allows for a near-doubling of the payload volume while still achieving the same launcher performance.

     The mass increase of the fairing due to the additional structure is approximated to be 36% of the standard payload fairing design. The standard Atlas-V HLV payload fairing has a mass of 4,400 kg, which results in an increase of 1,600 kg. This increase is subtracted from the launcher performance. Analogously, the Falcon Heavy fairing and performance is adapted. Finally, one has to take into account the additional cost for the development of the new shroud.

C. Habitation Elements

     In developing the habitation elements for the mission, the following general systems architecture guidelines were followed to maximize system and operational reliability and flexibility, and, ultimately, the safety of the crew: (1) Leverage systems that are currently in use or development to minimize development cost and risk, (2) Maximize commonality across all mission elements to increase system robustness, lowering the number of spares required, and decreasing the costs of system development and manufacturing, (3) Maximize multifunctionality and synergies among systems, yielding increased functionality for less mass, (4) Account for crew safety during all mission modes, and (5) Implement lessons learned from past programs. Based on these guidelines, the following architectural choices were made.

1. Deep Space Habitat (DSH)

     ISS-derived habitat structures were chosen as a baseline architecture for the DSH (Fig. 7), with modifications most notably made in the radiation protection to protect the crew for a long duration mission. Using modified ISS modules for the habitat is advantageous as the development work will be minimal, the system reliability has been demonstrated, ISS hardware is already flight-qualified, and ISS infrastructure such as payload racks and MPCV integration can be easily incorporated. The habitable volume is 76.3 m3, which is about 25% greater than the optimal recommended habitable volume for a crew of three for a mission duration of this length, according to the Celentano curve. This habitat will be configured for both on- and off-duty use.

     The primary Environmental Control and Life Support System (ECLSS) in the DSH is a closed-loop system similar to what is used on the ISS to minimize consumable mass. The secondary ECLSS is a passive system, known as Water Walls, that filters waste products through a series of forward osmosis treatment bags. Including both of these systems in the DSH design provides redundancy and increased radiation protection. The primary ECLSS design was validated for our crew size and mission duration using the software tool Environment for Life-Support Systems Simulation and Analysis developed at the Institute for Space Systems (Institut für Raumfahrtsysteme) at the University of Stuttgart, Germany.

2. Space Exploration Vehicle (SEV)

     The SEV (Fig. 8) is a pressurized “roving vehicle” currently being developed at NASA Johnson Space Center capable of short duration missions. It facilitates flexible exploration by the astronaut in both the intravehicular and extravehicular environments through the use of robotics and spacewalks, respectively. Moreover, the use of suitports in the vehicle enables the rapid transition of crew members between intravehicular and extravehicular activities when required. The SEV has a pressurized volume of 54 m3 and is capable of sustaining a two person crew for a maximum duration of 30 days. Due to the short mission duration for this vehicle, an open-loop ECLSS system architecture has been chosen to ensure high reliability, reduced complexity, and commonality between the vehicle and the Portable Life Support System (PLSS) of the spacesuits.

3. Extravehicular Mobility Unit (EMU)

     The NASA-ILD Dover Mark III Spacesuit will be used for exploration outside of the SEV. This spacesuit has been baselined by NASA as the next generation spacesuit design, and has been designed to interface with the suitports onboard the SEV. The PLSS, which interfaces with the Mark III suit, will provide life support for the astronauts during extravehicular operations.

4. Orion Multipurpose Crew Vehicle (MPCV)

     The Orion MPCV has been chosen as the baseline Earth reentry vehicle. This vehicle has been under extensive development by Lockheed Martin to support future NASA exploration missions, and has been designed with safety during all mission phases as its primary objective.

5. Spacecraft Atmospheres

     Spacecraft atmospheres were chosen based on those suggested by the NASA Exploration Atmospheres Working Group to ensure atmospheric capability between spacecraft elements while ensuring that pre-breath time for the required EVA frequency is properly accounted for. Table 5 lists the atmospheres selected for each habitation element to be used in the mission. It should be noted that nitrogen was chosen as the diluent gas in each atmospheric composition. The design presented here for the EMU requires no pre-breathe time.

Table 5. Atmospheres selected for each habitation
Habitation
Element
Atmospheric Pressure
and
Composition
DSH101.3 kPa (14.7 psi), 21% O2 nominally
70.3 kPa (10.2psi)
26.5% O2 during pressurization with the SEV
SEV70.3 kPa (10.2 psi)
26.5% O2
EMU57 kPa (8.3 psi)
100% O2 (Mark III suit)
MPCV101.3 kPa (14.7 psi)
21% O2 nominally
70.3 kPa (10.2 psi)
26.5% O2 during depressurization
prior to EVA from the vehicle

D. Human Factors

3. Crew Health Care

     a. Medical care

     Medical equipment and supplies consist of a standard ISS medical kit scaled up from 460 kg to 1000 kg of equipment, including a high-resolution ultrasound imager and expanded surgical supply kit. Medical consumables will also resemble those used in the ISS Health Maintenance System, expanded from 260 kg to 500 kg of pharmaceuticals and other consumable supplies. This provides a total medical supply kit for the mission with a mass of 1500 kg and an approximate volume of 6.5 m3, a size that fits comfortably into the larger mission design.

     b. Psychological considerations

     Long-term spaceflight produces extreme psychological stress, which, if ignored, can result in serious degradation of mental health that puts the mission and crew at risk but, if recognized in advance, can be mitigated. Major sources of psychological stress include isolation, interpersonal conflict, physical deterioration, separation from family, and lack of privacy. To improve the psychological well-being of the astronauts, it is important to provide them with nutritious food, communication with family, entertainment, and exercise throughout the duration of the mission. Typically, astronauts are provided with a variety of dehydrated food for their meals. To supplement their nutritional intake, small plants that serve as a food source may be included in the mission. Psychological benefits may be gained by both maintaining plants and harvesting them to obtain fresh food. When not conducting on-board science experiments, the crew members will be able to spend leisure time much as they would on Earth, reading books, listening to music, and emailing with friends and family.

     c. Countermeasures and mitigation strategies

     As much as possible, deleterious effects of space travel will be minimized through various countermeasures and mitigations strategies (Table 6). Spinning the habitat to create an artificial gravitational force is unreasonable due to the size of the spacecraft. There is a level of risk accepted in astronauts developing long-term adverse side effects due to the microgravity deconditioning. Once the mission is complete, the crew will have access to a full range of medical facilities to regain pre-flight levels of health and fitness.

4. Radiation

     a. Monitoring

     Tissue Equivalent Proportional Counters (TEPCs)--currently in use on the ISS--measure radiation doses for complex radiation fields and should be deployed in several locations in the DSH and SEV to measure radiation levels during transit, exploration, and EVA. Radiation levels throughout the DSH can be actively evaluated using portable TEPCs, allowing the crew to move to the most highly protected region of the vehicle during a solar particle event (SPE). An instrument similar to the Radiation Assessment Detector on MSL (also soon to be deployed on the ISS) will be deployed by the science team on the exterior of the DSH to record charged particle and neutron incidence for scientific use. SPE monitoring will be conducted using the existing network of solar observatories (i.e., SDO, SOHO, GOES) and any future expansion.

     b. Mitigation

     The mission architecture provides for 20 g/cm2 of uniform radiation shielding in the DSH. This degree of shielding is referenced in NASA documentation as the convergent design option for human missions based on an SPE mitigation/ mass trade33. Radiation shields that incorporate low atomic mass materials are capable of suppressing damaging secondary radiation in the form of neutrons that are ejected during particle transit through the aluminum hull of a module.

     c. Exposure estimates

     NASA has calculated the safe number of days that a person can travel in space when their vehicle is designed as mentioned above. These values are based on the need to prevent astronauts from exceeding an increased risk of 3% for REID (at the 95% confidence level). Failure to mitigate the effects of GCR and SPEs could lead to acute health effects including radiation sickness leading to incapacitation or death. Long-term risks include carcinogenesis, neural tissue damage, stem cell disturbances, and cataracts.

E. Science Mission and Surface Operations

1. Precursor Operations

     The goals of the precursor mission, in order of importance, are as follows: (1) to determine if humans can safely land on Phobos during the primary mission, or on Deimos in the event that Phobos is not feasible, (2) to establish a communications relay system that will facilitate the primary mission, (3) to gain important information regarding the nature and composition of the primary landing site to plan for a landing of the primary mission, and (4) to acquire remote sensing data on both of the moons to be used to understand their composition.

     In order to carry out the primary mission of landing humans on one of the Martian moons, we must characterize the structure and surface properties of Phobos and Deimos. This will be achieved on both moons using four impactor experiments at preselected sites, in-situ sampling and analysis conducted remotely using an immobile lander with an extendable arm, and a combination of remote observations from the PDS. In-situ sampling sites will be determined based on the findings from initial remote sensing surveys conducted by the PDS and from the impactor experiments.

     Impactor sites were selected in order to target sites of geologic interest, sites where future missions might land, and other widely-spaced sites to learn more about the distribution of the surface characteristics. The surface characteristics revealed by the four impactor tests will be a strong driver in determining how and where the lander is deployed and how and where the manned mission will dock and operate.

     To reduce complexity, DE and PE will be identical lander and impactor packages. The impactor package will be modeled after the one planned for the Japanese Lunar-A mission, but with four penetrometers instead of two. The Deimos and Phobos landers could be modeled after the Philae lander used in the Rosetta mission.

2. Science Instrumentation

     The primary instrument objective is to assess the surface environment to optimize human interactions with the surface environment of Phobos. In order to do this it is important to execute a comprehensive study of the planetary bodies to ensure the safety of the astronauts and the completion of mission objectives. The instrument suites have been designed to investigate the nature of the surface and subsurface of the Martian moons. This is a useful investigation for several reasons: (1) Determination of the nature of the regolith (uppermost, loose ‘soil’) allows assessment of the mechanical and chemical properties of the surface, (2) Identifying the strength and porosity of the surface provides critical information to help plan docking and anchoring maneuvers during the manned component of the mission, and (3) Studies of the flux of interstellar material and radiation levels will help to develop shielding techniques.

     Three unique science instrument suites (Surveyor, Explorer, and Expedition; Table 7) have been designed to achieve the aforementioned science objectives during the mission.

     The Surveyor suite is comprised of a number of heritage spectrometers and cameras, configured to investigate regolith properties remotely from orbit. Instruments in the Surveyor suite will provide valuable data on the topography of Phobos and Deimos, the flux of interplanetary material crossing the orbital plane of Phobos and Deimos, surface mineral composition, volatile abundance (such as water and CO2), and the strength of the magnetic fields on the moons.

     The Explorer suite is modeled after instruments from past NASA and JAXA missions35-42. The penetrometer device contained within the impactor package is modeled after the piezoelectric sensing element used in the Huygens probe. It was uniquely calibrated to withstand cryogenic temperatures, and future development will allow impact velocities of 300 m/s. The voltammetry, spectroscopy, and x-ray diffraction/fluorescence instruments (Wet Chemistry Lab, LIBS, and CheMin, respectively) do not require modification and are replicas of the original instruments. The Phobos and Deimos landers will employ robotic arms built on 360° swivels to deliver multiple regolith samples to the experiment chamber. Lastly, an additional micrometeroid detector (modified for a lander spacecraft) will be deployed in the Explorer suite to assess impact rates and material deposition. This instrument will provide details about the nature of the landing environment in which astronauts will execute future EVAs. Once positioned on the surface of Phobos or Deimos, micrometeoroid detector panels will deploy along the sides of the lander.

     Astronauts will manually deploy the Expedition suite of science instruments during EVA sorties. Seismic and radiation studies will be undertaken using heritage instruments. The PRSC (Planetary Retrieval of Subsurface Cores) will be based on core drilling that was done on the moon during the Apollo 15-17 missions but will have a somewhat larger core diameter for increased sample return. The ChipSat instruments will be developed as a public outreach effort to achieve the aims of independent science groups from around the world. Together, the employment of these instruments represents an innovative approach to meet a principal science objective of both on-the-ground data collection and sample return; they also promote the use of scientific equipment that is smaller in scale and lighter in weight.

3. Manned Phobos Operations

     Although one could argue that many of the mission’s scientific goals could be achieved through robotic means, there is a decided advantage to having humans “on the ground” to collect samples and to deploy instruments. Humans are versatile in that they can evaluate samples for quality and quantity in real time and can troubleshoot instruments on the fly. Up to this point in time, no robotic mission has returned extraterrestrial samples to Earth, and those samples that have been returned via manned EVAs have reaped great rewards for the scientific community.

     With these points in mind, the primary mission is designed to have two human crew members collecting samples directly on Phobos’ surface. The largest challenge to realizing this task is the near-zero gravity of Phobos. To operate on the surface, the crewed SEV will perform a rendezvous and docking procedure with the moon at each of the two pre-selected landing sites, where the vehicle will be anchored to the surface. Whether a conventional harpoon or drilltype anchor will be used, as opposed to an unconventional method such as microspines or netting47,48, will depend on site surface material characteristics (grain size, depth, density, cohesion). This composition information will be provided by data collected during the precursor mission.

     Once on the surface, the astronauts will have two modes available for EVA operations, depending on surface conditions. The first mode consists of one astronaut collecting samples and placing instruments with their feet fixed to the end of a robotic arm that extends from the SEV. In this configuration, one crew member must remain inside the vehicle in order to operate the robotic arm. This type of EVA has been shown to provide the most mobility of the methods investigated during the NEEMO under-water simulation program for manipulation of equipment in a microgravity environment. The robotic arm configuration would be especially advantageous if the surface regolith proves to be so thick and fine-grained that conventional maneuvering is unfeasible. The second mode, appropriate for sand- to boulder-sized regolith, anchors the astronauts to the Phobian surface using a tether and a scaled-down version of the regolith anchor used by the SEV. A secondary safety tether is connected to the SEV. For safety purposes, this second mode nominally involves only one crew member on EVA at a time, while the other performs monitoring and non-EVA activities inside the SEV. For both modes, use of an MMU-type device may aid astronaut maneuvers on the surface.

     Surface activities include collecting geologic samples and placing seismometers and retro-reflectors, as well as other experiments, such as ChipSat deployment. A full schedule of surface operations for the two SEV crew members over the two-week mobilization on Phobos is shown in Table 8. The Phobian surface exploration segment of the mission is designed for a nominal length of 14 days, as constrained by the planned mission trajectory. Using state-ofthe- art portable life support technology and relevant suit design, the safe duration for a single EVA has been estimated to be approximately four hours. Considering both rest periods and the constraints of the PLSS, the four-hour-EVA duration will allow for a maximum total of ten EVAs. Including contingency PLSS operational supply and potential for unanticipated required surface activity, a target number of eight EVAs has been planned, with four EVAs at each of the two predetermined landing sites.

     Surface samples will include rock and regolith scoop samples, as well as drilled core samples. Drill cores are planned to be 40-50 mm diameter x 3 m in length. The drill will be an electrically powered percussive hammer system, with a similar foot treadle contingency design for core removal as used on the Apollo 15-17 missions. Core samples will remain in their sleeves for direct placement into storage on the SEV. Thin samples of top-layer, fine-grained regolith will also be collected using adhesive pads. This will allow specific study of the regolith immediately exposed to the space environment. Loose geologic samples in collection bags and core samples will be stored in an exterior containment unit that will be placed in the SEV airlock using the robotic arm. The samples will be stored at appropriate cryogenic temperatures once on the SEV. Upon returning to the DSH, drill core will be stored in the modified MELFI freezer along with biologic samples to bring back to Earth.

     According to the surface operations schedule, on days three and ten passive seismic arrays will be placed on the Phobian surface to record seismic waves generated by internal strain in the moon. At each of the two landing sites, the seismometers must be placed with a spacing of 10s to 100s of meters apart and the exact location of each instrument recorded. EVA mode two would be preferable for this experiment, as it allows astronauts greater reach, and seismometers may be placed farther apart, which in turn, allows for deeper imaging into the crust of the moon. Inclusion of an active-source seismic array remains under consideration.

     On day two, one or more retro-reflector(s) will be placed on the Phobian surface. Retro-reflectors are mirrors that are used to reflect an electromagnetic signal back to its source. This instrument, once placed on Phobos, will have future application when a signal can be directed to the moon from rovers or stations on the Martian surface in order to determine the orbital distance of Phobos. The reflector measurements are regularly made over a long period of time (i.e., decades) to determine deviations in orbit. This method has yielded excellent results using our own moon, and we anticipate that it will answer such scientific questions as the rate at which Phobos is encroaching on Mars.

     After the full two week period on the surface, the SEV crew will return to the DSH and send the SEV (remotely) on a return path to Phobos, where it will re-anchor to the surface to prevent risk of drift-off and to comply with the requirements of planetary protection. The crew will be able to operate the SEV robotic arm remotely from the habitat, in order to continue to perform surface experiments after completion of the main mission and to demonstrate feasibility of performing telerobotic operations from orbit.

Atomic V-2 Rocket

Atomic V-2
ΔV8,120 m/s
Specific Power277 kW/kg
Thrust Power4.7 gigawatts
EngineSolid-core NTR
Specific Impulse915 s
Exhaust velocity8,980 m/s
Initial Thrust850,000 N
Maximum Thrust1,050,000 N
Wet Mass42,000 kg
Propellant Mass25,000 kg
Dry Mass17,000 kg
Payload3,600 kg
Inert Mass13,400 kg
Mass Ratio2.47
Turbopump Mass1,800 kg
Engine Mass
(including reactor)
4,200 kg
Reactor Mass1,600 kg
Height~60 m

The German V-2 rocket was an ultra-scientific weapon back in World War 2, in 1944. Unfortunately it only had a payload size of 1,000 kilograms. This is adequate for a small chemical warhead, but too small for a worth-while 1945 era nuclear warheads. If you want to invent an ICBM, the V-2 is just too weak.

Scott Lowther found an interesting 1947 report by North American Aviation (details in Aerospace Project Review vol 2, no.2, page 110). It had a simple yet audacious solution: take a V-2 design and swap out the chemical engine with a freaking nuclear engine! Atomic powered ICBMs, what a concept!

Anti-nuclear activists reading this are now howling with dismay over their narrow escape, but the NERVA will give the rocket a whopping 3600 kilograms worth of payload. That is large enough for a useful sized ICBM warhead.

But the US military managed to design two-stage chemical ICBMs, and the atomic V-2 became another forgotten footnote to history. But if you are an author writing an alternate history novel, you might consider how differently WW2 would have turned out if Germany had developed this monster.

Aurora CDF

Aurora Mars Mission
Num Crewx6
Crew Landedx3
Mass Schedule
Habitat Module
(THM)
66,700 kg (wet)
56,500 kg (dry)
Mars Lander
(MEV)
46,500 kg (wet)
29,000 kg (dry)
Earth Reentry
Capsule (ERC)
11,200 kg (wet)
10,200 (dry)
Consumables10,200 kg
Propellant1,083,000 kg
Propulsion130,000 kg
Structure19,700 kg
Wet Mass1,357,000 kg
Mars Samples65 kg
Trajectories
Trans-Mars
Insertion ΔV
(TMI)
3,639 m/s
Mars Orbit
Insertion ΔV
(MOI)
2,484 m/s
Trans-Earth
Insertion ΔV
(TEI)
2,245 m/s
Earth Atmo
Entry Vel
11,505 m/s
Earth Departure08 Apr 2033
Mars Arrival11 Nov 2033
Surface Stay30 days
Mars Departure28 Apr 2035
Earth Arrival27 Nov 2035
Engine
Cryogenic EngineVULCAIN 2
Cryogenic Isp450 sec
Cryogenic
Exhaust Vel
4,415 m/s
Storable EngineRD 0212
Storable Isp
(optimistic)
345 sec
Storable
Exhaust Vel
(optimistic)
3,385 m/s
Storable Isp
(realistic)
325 sec

This is from CDF Study Report Human Missions To Mars from the European Space Agency. The report is over 400 pages long, going into excruciating detail, so I'm only going to hit the high points.

The report cautiously states The main objective of the study was not to define an ESA “reference human mission to Mars” but rather to start an iteration cycle which should lead to the definition of the exploration strategy the associated missions and the set-up of requirements for further mission design and further feedback to the exploration plan. In other words it is not a Mars reference mission, it is the start of figuring out how to make a process that will eventually craft a reference mission.

The spacecraft is composed of four parts:

PROPULSION MODULE (PM)
This is a conglomeration of seventeen chemical rocket engines organized into six stages. Chemical engines have such a lousy exhaust velocity that they must use multi-staging. They are attached to a segmented cylindrical spine which acts as the thrust frame.
TRANSFER HABITATION MODULE (THM)
The habitat module. Where the crew lives during the mission.
MARS EXCURSION VEHICLE (MEV)
The payload: the Mars lander. It lands three crew on Mars to cram in all the exploring they can possibly do in thirty days while living in the cramped hab mod. At the end of the month it returns to the spacecraft in orbit along with a whole 65 kilograms of interesting Mars rocks.
EARTH REENTRY CAPSULE (ERC)
The way the crew returns to Terra's surface. They abandon what is left of the spacecraft to its fate, and ride in the gumdrop-shaped reentry capsule on a blazing 11.5 km/sec aerobraking. The surface of the capsule may be contaminated by Martian bugs from the MEV, but the high-temperature reentry should adequately sterilize it. It is basically a glorified Apollo Command Module, with an extra-thick ablative heat shield.

PROPULSION MODULE (PM)

A "stack" is a single chemical rocket engine with its fuel tanks. These are clustered into "stages". All the stacks in a stage burn simultaneously.

Nuclear Electric, Solar Electric, and Nuclear Thermal were ruled out because they are not mature technologies.

Storable chemical fuel does not need cryogenic cooling and does not boil off, it is also nicely dense so the fuel tanks are small. But it has a much lower specific impulse. Cryogenic fuel is the opposite. The designers studied what would be needed to keep cryogenic fuel for the months long mission, and concluded it was unworkable. They compromised by using cryogenic for the Trans-Mars Injection burn, since the fuel would not have enough time to apprecialy boil away. The other burns would have to make do with storable chemical fuel.


Trans-Mars Injection Stages

Trans-Mars Injection requires 3,639 m/s of ΔV. It uses three stages of 4 stacks each, for a total of 12 stacks. Since the fuel tanks have just been filled in Terra orbit, the stacks can use cryogenic fuel. So these stacks use Vulcain 2 engines.

The first two stages insert the spacecraft into eccentric orbits, the third and final stage into the hyperbolic escape. After each burn, the spent stages are jettisoned and perform a controlled reentry. The final burn does not aim the spacecraft into the transfer orbit, because the designers do not want the third stage crashing into Mars. Instead it aims the ship almost into the orbit, after jettison the ship uses its reaction control system to change course into the transfer.

Each of the three stages is a segment of ship spine with four rocket engines (stacks) attached. When the stage completes its burn, both the spine and engines are jettisoned.


Mars Orbit Insertion Stage

Mars Orbit Insertion requires 2,484 m/s of ΔV. It uses two stages of 2 stacks each, for a total of 4 stacks. Since MOI occurs almost seven months into the mission, cryogenic fuel cannot be used (by this time it would have all boiled away). Instead storable NTO/UDMH is used with a RD-0212 engine. Less exhaust velocity but no boiling.

The first stage has two stacks of 80 tonnes each, which performs the orbit insertion. The second stage has two smaller stacks of 50 tonnes each, which performes the final orbit acquisition.

Before the burn, the 4,900 kilograms of sewage (and other waste produced by the fact the life support system is not 100% closed) is jettisioned to increase the spacecraft's mass ratio.

When the first stage completes its burn, the two spent stacks are jettisioned. When the second stage completes its burn both the two spent stacks and the segment of ship's spine is jettisoned. This exposes the tiny Trans Earth Injection Stage, which had been hiding inside the spine segment.


Trans Earth Injection Stage

Trans Earth Injection requires 2,245 m/s of ΔV. It uses one stage containing one stack. This uses the same RD-0212 engine and has the same mass budget as the MOI stack. It has no spine segment to attach to. Instead it has the Propulsion Module Interface (PM I/F) on top, attached to the back node of the Transfer Habitation Module.

Before the burn, the 500 kilograms of sewage is jettisioned to increase the spacecraft's mass ratio. As well as the remaining parts of the Mars Excursion Vehicle.


TRANSFER HABITATION MODULE (THM)

The habitat module is a cylinder where the explorers live. It has two nodes, one at each end, to attach to the rest of the spacecraft. Each node has an interface (I/F) module, the propulsion module pluging into the PM I/F and the Mars excursion vehicle pluging into the MEV I/F.

The "back" node has an airlock (and spare docking port) and the Earth reentry capsule. It also has an EVA prep area (including three space suits), a toilet, and what passes for a shower (a "hygiene area"). For conceptual purposes the design is using an airlock straight off the International Space Station.

The "front" node has storage, a recreation area, a spare docking port, and the command area complete with a cupola. It also has the communication antennas. The cupola is kind of worthless but is included for psychological reasons (crew going bat-crap insane being cooped up in a tin can with no windows).

Each node has two solar power units, for a total of four. Each unit has a movable solar cell array and a storage battery.

The two nodes and the main cylinder can be sealed off from each other in the event one part springs a leak and depressurizes. If the main cylinder depressurises, the crew has to be evacuated to the front or back node for a couple of days until the leakage has been repaired.

The total habitable volume has a minimum of 450 m3; where 1/3 of the volume is used for storage, and the remaining 2/3 are the habitable volume. About 5% of the total volume has to be considered for the module structure.

The habitat module has 9 gm/cm2 of radiation shielding to stop enough galactic cosmic radiation to keep the astronauts under the yearly and career doses of radiation. The storm cellar has 25 gm/cm2 to protect the astronauts from solar proton storms.

The designers looked into adding a spinning habitat to help prevent the dire effects of prolonged free fall on the crew, but concluded it just had too much penalty mass. Instead the crew will just have to do daily exercise in a little one-person centrifuge.

The various areas inside the habitat are classifed by "zone":

  • PRIVIATE ZONE: Areas where the crew is always alone. Crew quarters
  • PERSONAL/UTILITY ZONE: Areas where the crew works/trains mostly on their own. Command, laboratory, exercise, toilet, hygiene, medical
  • SOCIAL/COMMUNAL ZONE: Areas where the crew is mostly with other crewmembers. Food preparation, eating, conferences, video

MARS EXCURSION VEHICLE (MEV)

The spacecraft will be orbiting Mars for 533 days. But the surface mission was limited to 30 days, because the mass and complexity of the MEV increases dramatically with surface stay time. Shorter than 30 days would not be worth the mission, since the crew will need about a week to get used to gravity and another week to prepare for lift off. The recommendations suggest seven EVAs as a minimum, which would take about two weeks.

The MEV has three parts: the Surface Habitation Module (SHM) where the Mars explorers live, the Descent Module (DM) which does it darnedest to get the MEV to the surface in one piece, and the Mars Ascent Vehicle (MAV) which gets the explorers back up to the orbiting spacecraft.

The descent module has four deorbit engines, an inflatable heat shield for aerobraking, and huge parachutes.


SURFACE HABITATION MODULE

The surface hab module is the Martian home-away-from-home for the three intrepid Mars explorers. It has enough life support for 30 days (i.e., 90 person-days). It has a total pressurized volume of 79 m3 and a habitable pressurized volume of 50 m3.

To recap, the various areas inside the habitat are classifed by "zone":

  • PRIVIATE ZONE: Areas where the crew is always alone.
  • PERSONAL/UTILITY ZONE: Areas where the crew works/trains mostly on their own.
  • SOCIAL/COMMUNAL ZONE: Areas where the crew is mostly with other crewmembers.

MARS ASCENT VEHICLE

This is the vehicle the explorers use to leave Mars and return to the orbiting space station. It is composed of a capsule, and two propulsion stages. The explorers ride in the capsule when the MEV lands, because it has the acceleration couches. The capsule has enough life support for five days (15 person-days). It has a habitable volume of 4 m3.

After leaving Mars and entering orbit, the capsule may take a few days to dock with the spacecraft.


MISSION

Upon arrival in Mars orbit, the crew spends one week doing a systems check of the entire spacecraft.

They do a more thorough two week check before the Trans Earth Injection kick.

Aurora CDF Project Troy

Project Troy
Engine
Mission
TypeChemical
(Cryo LOX/LH2)
Exhaust
Velocity
4,600 m/s
Isp469 sec
Uncrewed Precursor
Mission
TMI Kick ΔV3,620 m/s
MOI Kick ΔV2,397 m/s
Transfer Time264 days
TOTAL ΔV6,017 m/s
Crewed Principal
Mission
TMI Kick ΔV3,518 m/s
MOI Kick ΔV2,594 m/s
TEI Kick ΔV1,801 m/s
EOI Kick ΔV3,759 m/s
TOTAL ΔV11,672 m/s
Outbound
Transfer Time
251 days
Homebound
Transfer Time
282 days

This is from Project Troy: A Strategy for a Mission to Mars (2007).

This appears to be a study to promote Reaction Engine Limited's proposed SKYLON spaceplane.

It starts off by skimming over the highlights of NASA's Design Reference Mission (DRM) to Mars, and the ESA's response: the Aurora CDF mission. The report notes that the Aurora mission will work, but it unfortunately requires 25 main assembly launches to get all the components into orbit, plus two or three more to top up the propellant tanks. At a rate of one launch per two months it will take about 4.6 years to get the entire clanking mess up and assembled. Given the cost of boosting all that mass and the limited flight rate of expendable vehicles from existing facilities, realistically there is no way that Europe can afford to foot the bill for this mission.

Then the report brightly mentions that if the components are redesigned to work with REL's wonderful SKYLON, it becomes much more affordable.

For a fraction of the price of the Aurora CDF it could reproduce it. However this would be a dangerous mission with zero emergency contingencies that provides very little scientic return for its investment (little more than a "Flags & Footpring mission"). For a bit more money the program can send an uncrewed precursor mission full of supplies and scientific equipment, adding emergency back-up and increasing scientific return. If the crewed ship fails they could survive on Mars until relieved by a rescue mission. Scientifically it will allow a 14 month mission on the Martian surface by a distributed team of 18 explorers cover 90% of the planet's surface.

And for a bit more the program can send a fleet of three crewed spacecraft, enabling a full crew return even if one spacecraft fails.

The report points out that since SKYLON is reusable, this will not just be a Mars mission, it will be more of a Mars Transport system infrastructure. What the report only hints at is this would be a good reason to build SKYLON in the first place, which some cynics were wondering out loud if it was a bad idea. REL wanted some good PR full of reasons to invest in SKYLON. The way they put it: "The creation of a reusable transportation system which will go on to reduce the cost of space activity by over an order of magnitude long after the Mars missions are achieved would be a suitable legacy from such a laudable undertaking."


The propulsion section has three stages: the Earth Departure Stage (EDS), the Mars Transfer Stage (MTS) and the Earth Return Stage(s) (ERS). An automated uncrewed precursor mission delivers a habitat module and power supplies to the Martian surface and establishes orbital facilities two years before the crewed mission departs. Of course the second mission only departs after all the assets perform self-checkouts and report success to Terra. The assets are not just to assist the mission, they are emergency back-up in case the crewed ship malfunctions and the crew has to shelter in place on Mars until a rescue mission arrives.

The fuel is cryogenic liquid-oxygen / liquid hydrogen, along with the headache of cryogenic boil-off. The report looked at using methane instead of hydrogen because it does not boil-off, but the drastic increase in mission mass lead to rejecting that option.

The Earth Departure Stage is designed to be reusable, so it can send off both the precursor and the primary spacecraft. It boosts the spacecraft from LEO to just short of escape velocity. It separates and allows the spacecraft to continue to Mars. The EDS is now in a highly elliptical synchronous orbit with respect to the Troy Operation Base Orbit, it uses that orbit to return. Meanwhile the Mars Transfer Stage burns to complete spacecraft insertion into Mars transfer orbit.

On Mars, a small nuclear power supply is used to manufacture O2 and CO fuel out of carbon dioxide in the Martian atmosphere. This is used to fuel a single stage Ferry used to transfer from and to Martian orbit and between locations on the surface. The fuel can also be used in solid oxide fuel cells to power surface rover vehicles.

The report looked into using aerobraking for Earth capture instead of propulsive capture, but found it wasn't worth it. The payload mass would be reduced by half, which drastically reduced the value returned by the mission. Instead the report went with a more modest atmospheric assisted capture.

A three ship mission would not cost three times as much, due to the economy of scale. Two ships provides great redundancy, three ships allow up to 90% of the Martian surface to be explored. True, it would need three precursor missions instead of one, but it would be a cheaper than the Apollo missions. Apollo involved the launch of 30,000 metric tons to put 18 astronaut near Luna (12 who landed on the surface) over a period of four years.

Austin Mars Mission

This is from Proceeding of the Symposium on Manned Planetary Missions 1963/1964 Status, part 9: A STUDY OF MANNED MARS EXPLORATION IN THE UNFAVORABLE TIME PERIOD (1975-1985) by Dr. R. N. Austin of General Dynamics. Like all the other studies in the document, the landing craft was designed assuming that Mars' surface atmospheric pressure was 85 millibars so aerobraking could be used. Alas the Mariner 4 probe found it was closer to 7 millibars, aerobraking ain't gonna work.

Like the Boeing IMIS, the Mars mission was accomplished by using multi-staging. And with the same insane logic the design uses Nuclear Thermal Rocket stages. The only improvement is that Austin's design only ejects three nuclear reactors glowing with blue radioactive death for the next ten-thousand years into random orbits in the solar system, instead of five like in the Boeing design.

Staging was also mandated by the initial requirement that the nuclear engines were not to be restartable. This improves reliability by decreasing the operating time of any given engine.

Granted, the point of the study was to see how bad the design got if you purposely chose a launch date with an unfavorably high delta-V requirement (due to Mars' eccentric orbit) and during the solar proton storm maximum necessitating extra storm cellar mass. Producing the extra delta-V is a challenge. But still, discarding nuclear reactors like throwing a cigarette butt out the window would be frowned upon nowadays.

In the diagram at left:

  • RED: Terra Escape Stage
  • ORANGE: Mars Braking Stage
  • YELLOW: Mars Escape Stage
  • GREEN: Terra Braking Stage
  • LIGHT BLUE: Mission Module
  • DARK BLUE: Terra Reentry Module
  • VIOLET: Mars Excursion Module

The three habitable components are:

  • MISSION MODULE: provides living quarters for the six crew throughout the mission (LIGHT BLUE)

  • MARS EXCURSION MODULE: transports explorers between Mars orbit and surface (VIOLET)

  • TERRA REENTRY MODULE: provides a capability for atmospheric entry, landing, and safe return of crew to Terra (DARK BLUE)

The six propulsive stages are:

  • TERRA-ESCAPE: boosts spacecraft from LEO into trans-Mars trajectory (nuclear)

  • OUTBOUND MID-COURSE CORRECTION: tweeks ship's trajectory en-route to Mars (chemical)

  • MARS-BREAKING: moves spacecraft into circular Mars orbit (nuclear)

  • MARS-ESCAPE: boosts spacecraft from Mars orbit into trans-Terra trajectory (nuclear)

  • INBOUND MID-COURSE CORRECTION: tweeks ship's trajectory en-route to Terra (chemical)

  • TERRA-BRAKING: slows down the spacecraft, allowing the crew to bail out of the ship in the Terra reentry module and safely land on Terra. (chemical)

The three nuclear stages are stacked for ease of staging. The outbound mid-course correction (MCC) chemical engine is on the spacecraft's nose. The Terra-braking chemical engine is on the base of the mission module. The inbound MCC chemical engine will either be incorporated in the Terra-braking engine or mounted adjacent to the outbound MCC depending upon size.

If something catastrophic happens during the Terra-escape manuever, the crew can abort the mission via a thrust reverser mounted in the exhaust nozzle of the outbound MCC engine. It will detach the Terra reentry module and send it back to Terra. If something happens during other maneuvers, the crew is out of luck.

Electrical power is supplied by a Snap-8 reactor located aft of the Terra-braking engines, in the hope that the latter's fuel and oxidizer tanks will provide some of the required radiation shielding.

In the spin-gravity variant, the entire fore end of the ship rotates to provide artificial gravity. The mission module is part of the rotating section, except for the storm cellar. That is stationary, with the rotation bearing mounted on the fore end of the storm cellar. The rest of the mission module is divided into two cylindrical compartments on the end of long arms, each housing three crew. The arms have a folding parallelogram arrangement to move the mission modules to the center axis during thrust periods. In theory the crew can easily move to the storm celler with the arms in either position. Spin is created and removed by reaction jets mounted at the tips of the arms. Rotation bearing friction is counteracted by a synchronous electric motor.

The design would be much simplier if the there was no bearing and the entire ship rotated. However, the designers had doubts that accurate navigational observations could be made from a rotating platform.


The study looked at replacing the nuclear Mars-braking stage with an aerobraking heat shield. The thought of man-rating such a huge spacecraft carrying a nuclear engine on a fiery roller-coaster ride through the Martian atmosphere is rather daunting. The assessment board will take one look at the design and laugh in your face.

Mercifully Mariner 4's measurement of the tenuous Martian atmosphere made such aerobraking schemes impossible.

These Mars excursion modules won't work either because of aerobraking problems. Both have a maximum gross weight of 32,800 kg, crew of 3, and must be capable of being stored on the mother spacecraft with a 7.6 meter diameter space.

But the Terra reentry vehicle should work just fine. Mass of 4,000 kg, not including the 6 crew and the heat shield.


VEHICLE ASSUMPTIONS

  • Maximum allowable Terra-entry velocity of 15.24 km/sec
  • No aerobraking at Mars
  • Short missions have a stay-time on Mars of 40 days
  • Long missions have a stay-time determined by next launch window
  • 3% reserve delta-V
  • Mid-course correction delta-V 250 m/sec
  • Nuclear engine initial acceleration 0.3g
  • Chemical engine initial acceleration 0.5g
  • Crew size: 6
  • No spin gravity
  • All components have meteoroid protection, except Terra departure tanks
  • Cryogenic propellant is stored using insulation and boil-off margin, no refrigeration used
  • Maximum allowed crew radiation dose: 2 Grays
  • Nuclear engines: graphite-core, no restart, one for Terra-departure, Mars-arrival, and Mars-departure
  • Chemical engines: cryogenic chemicals, one for Terra-arrival, and each mid-course correction stage.

The study also looked at some variants that could improve performance if allowed. These included replacing the chemical engines with nuclear, allowing restartable nuclear engines, aerobraking in the Martian atmosphere, allowing a higher Terran reentry velocity, using Orion nuclear pulse propulsion, and filling the propellant tanks in LEO immediately prior to Terra departure. All of these reduced the vehicles mass, allowing more payload. Refer to the study for details.

Basic Solid Core NTR

Overview

RocketCat sez

Now this is design to pay attention to. Dr. Crouch did this one to a queen's taste, with plenty of delicious detail. Even if he did have some outrageous ideas, like detaching the freaking atomic reactor for splashdown and recovery in the Pacific Ocean!

This is from NUCLEAR SPACE PROPULSION by Holmes F. Crouch (1965).

Please note that this is a strict orbit-to-orbit ship. It cannot land on a planet.

The Command Capsule contains the payload, the habitat module for the crew, the ship controls, life-support, navigation equipment, and everything else that is not part of the propellant or propulsion system. It is designed to detach from the ship proper along the "Payload Separation Plane."

The Rocket Reactor is the actual nuclear thermal rocket propulsion system. It too is designed to detach from the ship proper along the "Reactor Separation Plane." This allows such abilities as to jettison the reactor if a criticality accident is immanent, to swap an engine for an undamange or newer model engine, or to return the engine Earth via splashdown.

The book had most of a chapter about returning an engine to various locations in the Pacific ocean where international condemnation was low enough and the problems of designing an ocean-going recovery vessel that can fish the reactor out of the water without exposing the crew to radiation. What an innocent age the 1960's were, that sort of thing would never be allowed nowadays. The illustrations above are provided for their entertainment value.

The propellant tank contains the liquid hydrogen propellant. The payload interstage and the propulsion interstage are integral parts of the propellant tank, and contains hardware items of lesser value than the payload and the reactor. The propulsion interstage also contains the attitude jets. As with all rockets, the propellant and its tank dominate the mass of the spacecraft. A larger propellant tank or smaller strap-on tanks can be added to increase the mass ratio. Note that the main propellant tank is load-bearing, it has to support the thrust from the engine. But the strap-on tanks are not load-bearing, they can be made lightweight and flimsy.

ItemMass (kg)Average Diameter (m)Overall Length (m)
Payload15,0004.579.14
Engine6,8001.52 to 3.056.10
Tank (empty)22,7007.3238.1
Tank (full)90,700--

Sample specifications : wet mass: 112,500 kg, maximum thrust 445 kN, specfic impulse 800 seconds. That implies a thrust-to-weight ratio of 0.4, which is its acceleration in gs when the propellant tank is full. The figures below imply a mass ratio of 1.5, and a ΔV capability of 3,200 meters per second. The spacecraft's specific power is 23 kilowatts per kilogram

The book implied that a solid core engine could be devloped up to a specific impulse of 1000 seconds, with a max of 12,000 seconds (but at max you'll be spewing molten reactor bits in your exhaust). A later design in the book had a specific impulse of 1000 seconds and a ΔV capability of 15,000 m/s (which implies a mass ratio of about 4.6, which is a bit over the rule-of-thumb maximum of 4.0). Please note that the dimensions below were originally in feet and pounds in the book, that's why they are such odd numbers (e.g., 1.52 meters is 5 feet).

Rescue Ship

This is a variant on the basic NTR rocket: the nuclear rescue ship. This is for use by the outer-space version of the Coast Guard.

Note the "Neutron isolation shield" between the two reactors. Nuclear reactors are throttled by carefully controlling the amount of available neutrons within the reactor. A second reactor randomly spraying extra neutrons into the first reactor is therefore a Bad Thing. "Neutronically isolated" is a fancy way of saying "preventing uninvited neutrons from crashing the party."

Reactor

The propulsion interstage is the non-nuclear part of the propulsion subsystem. It contains the propellant plumbing, the turbopump, and the attitude control system.

The nuclear part of the propulsion system is the rocket reactor. This is basically the reactor, the exhaust nozzle, and the radiation shadow shield.

The rocket reactor is designed to be detachable from the rest of the spacecraft.

Shadow Shield

The shadow shield casts a protective shadow free from deadly radiation. Care has to be taken or other objects can scatter radiation into the rest of the ship. Any side tanks will have to be truncated so they do not emerge from the shadow. Otherwise they will be subject to neutron embrittlement, and they will also scatter radiation. The reason the reactor does not have shielding all around it is because the shielding very dense and savagely cuts into payload mass allowance. The shadow shield typically casts a 10 degree half-angle shadow.

Note that shadow shields will more or less force the docking port on the ship to be in the nose, or the other ship will be outside of the shadow and exposed to reactor radiation.

When the reactor is idling, the shadow shield does not have to be as thick. In order to widen the area of shadow (for adding side tanks or whatever), the secondary shadow shield could extrude segments as extendable side shields.

Plug Nozzle

For nuclear thermal rockets, the exhaust bell tends to be about twice the size of a corresponding chemical rocket nozzle. A small concern is meteors. While very rare, the shape of the bell will funnel any meteors into a direct strike on the base of the reactor. This can be avoided by replacing the bell nozzle with a Plug Nozzle.

The basic design uses a bell nozzle, and powers the attitude jets from the reactor. This might not be the best solution. Compared to a chemical rocket, the moment of inertia of a nuclear rocket is about ten to thirty times as large (diagram omitted). This is due to the larger mass of the engine (because of the reactor) and due to the more elongated shape of the nuclear rocket (because of the shadow cast by the shadow shield, and designers taking advantage of radiation's inverse square law). Taking into account the relative moment arms, the attitude jets will have to be four to twelve times as powerful. Conventional attitude jets might not be adequate.

Also note that with this design, the attitude jets cannot be used during a main engine burn. Further: attitude jets are pulse reaction devices (maximum change in the minimum time). Also there is a mandatory delay time between reaction pulses to permit the nozzles to cool off and to allow propellant feed oscillations to dampen out. None of these limits work well with nuclear thermal rockets.

Mr. Crouch suggests that the basic problem is that bell nozzles are not the optimal solution for nuclear engines. He suggests that plug nozzles (aka "annual throat nozzle") can solve the problems. Plug nozzles have problems with chemical rockets, but have advantages with nuclear rockets. Mr. Crouch mentions that wide design flexibity arises from the fact that the outer boundary radius (rβ) and cowl lip angle (β) can be varied. Translation: you can design a hinge into the shroud that will allow the cowl lip to wiggle back and forth. This will allow thrust vectoring.

Mr. Crouch also likes how a plug nozzle can be structurally integrated into the reactor, unlike a conventional bell nozzle. It is also nice that the subsonic setion of the nozzle requires structural support in the very region where the core exit needs support. What a happy coincidence! The support grid, the plenum chamber, the plug body, and the plug supports could be integrated into one common structure. You will, however, have to ensure that the hot propellant passes through the plug body support, not across it.

Note the reversed curvature of the propellant flow. This allows placement of neutron reflection material to prevent neutrons going to waste out the tail pipe. The propellant can move in curves, but neutrons have to move in straight lines. This will create a vast improvement in the neutronics of the reactor.

Of course there are problems. The biggest one is burnout of the cowl lips. The lip is thin and the exhaust is very hot. The lip will be burnt away unless special cooling techiques are invented (Here Mr. Crouch waves his hands and states that such cooling will only be invented if there is a compelling need, and the desire for a nuclear plug nozzle is such a need. Which is almost a circular argument). Some form of regenerative cooling will probably be used, where liquid hydrogen propellant flows through pipes embedded in the lips as coolant.

Thrust Vectoring

The plug nozzle lends itself well to thrust vectoring, thrust throttling, and nozzle close-off. This is because of the short shroud and the configuration of the cowl lip. Unlike a conventional bell nozzle there is no fixed outer boundary. While the cowl lip defines the outer periphery of the annular throat, there isn't an outer boundary. So all you have to do is alter the cowl lip angle to adjust the throat area, which will vector the thrust (that's what Mr. Crouch meant when he was talking about varying rβ and β).

In the diagram at right, variable throat segments A, B, C, and D are sections of the cowl which are hinged (so as to allow one to alter the lip angle). This will allow Yaw and Pitch rotations.

If the pilot wanted to pitch the ship's nose up, they would decrease the mass flow through segment A while simultaneously increasing the mass flow through segment C. Segment A would have its lip angle increased which would choke off the throat along its edge, while Segment C's lip angle would be decreased to open up its throat section. The increased thrust in segment C would force the ship to pitch upwards.

It is important to alter the two segments such that the total thrust emitted remains the same (i.e., so that segment A's thrust lost is exactly balanced by segment C's gain). Otherwise some of the thrust will squirt out among the other segments and reduce the amount of yaw or pitch thrust. With this arrangement, it is also possible to do yaw and pitch simultaneously.

The moment arm of thrust vectoring via a plug nozzle is greater than that of thrust vectoring from a conventional bell nozzle. This is because the thrust on a bell nozzle acts like it is coming from the center, along the thrust axis. But with a plug nozzle, the thrust is coming from parts of the annular throat, which is at some distance from the center. This increases the leverage.

Nozzle close-off means when thrusting is over, you can shut the annular throat totally closed. This keeps meteors, solar proton storms, and hostile weapons fire out of your reactor.

Pivoting each section of cowl lips is a problem, because as you pivot inwards you are reducing the effective diameter of the circle that defines the edge of the lips. The trouble is that the lip is not made of rubber. The solution used in jet fighter design is called "turkey feathers" (see images above). It allows the engine exhaust to dialate open and close without exposing gaps in the metal petals.

Cascade Vanes

With chemical rockets, retrothrust is achieved by flipping the ship until the thrust axis is opposite to the direction of motion, then thrusting. This is problematic with a nuclear rocket, since it might move another object out of the shadow of the shadow shield and into the radiation zone. For example, the other object might be the space station you were approaching for docking. Ideally you'd want to be able to perform retrothrust without changing the ship's orientation. What you want to do is redirect the primary thrust stream.

Jet aircraft use "thrust reversers." These are of two type: clam shell and cascade vanes. For complicated reasons clam shell reversers are unsuited for nuclear thermal rockets so Mr. Crouch focused on cascade vanes reversers. The main thing is that the actuators for cascade vanes are simpler than clam shell, and unlike clam shells a cascade vane reverser surface is segmented. There are five to ten vanes in each surface.

Note that the maximum reverse thrust is about 50% of the forwards thrust.

Each vane is a miniature partial nozzle. It takes its portion of the propellant flow and bends it backwards almost 180°. In the "cascade reverser end view" in the right diagram above, there are eight reversers, the wedge shaped surfaces labeled A, A', B, B', C, C', D, and D'. Each reverser is normally retracted out of the propellant stream, so their rear-most edge is flush with the tip of the cowl lip. When reversal is desired, one or more reversers are slid into the propellant stream. At maxmimum extension, the rear-most edge makes contact with the plug body.

Vane segmentation of the reverser surface eases the problem of center-of-pressure changes as the reverser's position is varied in the propellant stream.

Inserting all eight reversers causes retrothrust (see "Full Reverse" in below left diagram). Inserting some but not all reversers causes thrust vectoring. You'd expect that there would be a total of four reversers instead of eight (due to the four rotations Yaw+, Yaw-, Pitch+, Pitch-), but each of the four were split in two for reasons of mechanical alignment and the desirablity of shorter arc lengths of the vanes. This means the reversers are moved in pairs: to pitch upward you'd insert reverser A and A' (see "Thrust Vectoring" in below left diagram).

I am unsure if using reversers means that it is unnecessary to use the variable throat segments for yaw and pitch rotations, Mr. Crouch is a little vague on that. And the engineering of reversers that can withstand being inserted into a nuclear rocket exhaust is left as an exercise for the reader. There will be temperature issues, supersonic vibration issues, and edge erosion issues for starters. These are desgined for a solid-core NTR, where the propellant temperatures are kept down so the reactor core remains solid. This is not the case in a gas-core NTR, where the propellant temperatures are so high that the "reactor core" is actually a ball of hot vapor. The point is that a gas core rocket might have exhaust so hot that no possible material cascade vane could survive. There is a possibility that MHD magnetic fields could be utilized instead.

But the most powerful feature of cascade vanes is their ability to perform "thrust neutralization". When all the reversers are totally out of the propellant stream, there is 100% ahead thrust. When all the reversers are totally in the propellant stream, there is 50% reverse thrust. But in the process of inserting the reversers fully in the propellant stream, the thrust smoothly varies from 100% ahead, to 75% ahead, to 50% ahead, to 25% ahead, to 25% reverse, and finally to 50% reverse.

The important point is that at a specific point, the thrust is 0%! The propellant is still blasting strong as ever, it is just spraying in all directions, creating a net thrust of zero.

Why is this important? Well, ordinarily one would vary the strength of the thrust while doing maneuvers. Including stopping thrust entirely. Trouble is, nuclear thermal rocket reactors and turbopumps don't like having their strength settings changed. They lag behind your setting changes, and the changes put stress on the components.

But with the magic of thrust neutralization, you don't have to change the settings. You put it at a convenient value, then leave it alone. The cascade vanes can throttle the thrust to any value from 100% rear, to zero, to 50% fore. And do thrust vectoring as well.

Mr. Crouch also notes that while using thrust vectoring for maneuver, the rocket will have to be designed to use special auxiliary propellant tanks. The standard tanks are optimized to feed propellant while acceleration is directed towards the nose of the ship. This will not be true while manuevering, so special "positive-expulsion" tanks will be needed. These small tanks will have a piston or bladder inside, with propellant on the output tube side of the piston and some neutral pressurized gas on the othe side of the piston.

I was having difficulty visualizing the cascade reversers from the diagrams. I used a 3D modeling program called Blender to try and visualize them.

Belcomm Manned Venus Flyby

This is from the report Manned Venus Flyby.

In 1966 NASA created the Apollo Applications Program, in a desperate attempt to convince Congress that the space program was not wasting money. The AAP tried to come up with new crewed spaceflight proposals that had modest budgets, by re-using as many of the gold-plated Apollo Program technologies. NASA wanted to tell Congress: "You see? We didn't spend a kajillion dollars on a one-off project, we can use Apollo-tech for an entire family of missions!" NASA management was also concerned about losing the 400,000 workers involved in Apollo after landing on the Moon in 1969.

Didn't work.

The Johnson Administration slashed the AAP's budget in favor of his "Great Society" initiative. He gave the AAP a measly $80 million US out of the $450 million NASA had asked for, hamstringing the program from the very start. About the only thing that the AAP actually produced was Skylab.

One of the AAP's programs that never got anywhere was the Manned Venus Flyby. Because if you have a solution that you are frantically trying to find a problem for, you eventually start hammering extremely square pegs into very round holes.


NASA was toying with the idea of turning the upper stage of a Saturn V into a "wet workshop" in order to make instant space station. I think the concept dates back to Krafft Ehricke's Atlas Space Station proposal, but I digress. The idea is that the fuel tanks of a given Saturn V stage are freaking ginormous, once they are empty you'll have a huge pre-constructed volume suitable to make a kick-ass space station. No costly and dangerous orbital construction needed: it arrives ready to go.

The drawback is that all the space station internal fittings and equipment has to be capable of withstanding being immersed in liquid hydrogen at a temperature near absolute zero. This is quite a challenge. Perhaps you've seen a video where liquid nitrogen is used to freeze a rose until it shatters? Well, liquid hydrogen is about four times as cold. Gets inside everything as well, since the blasted stuff can seep in between the molecules of any wall.

Skylab was planned to be a wet workshop, since it was going to be boosted on the weaker but cheaper Saturn IB. Luckily for Skylab (but bad for NASA), Apollo missions 18 through 20 were canceled. This meant that NASA has some surplus full-strength Saturn Vs on their hands. The extra power means that the rocket stack does not need the fuel inside the Skylab stage, so it can be launched full of relatively harmless room temperature air instead of destructive liquid hydrogen near absolute zero. This is called a "dry workshop."

The Manned Venus Flyby unfortunately did not have that luxury. It had to be a wet workshop, which would make the interior equipment much more expensive. It was planned to be built around the same stage as Skylab, the S-IVB stage.


The Manned Venus Flyby needed all the space inside a wet workshop. The typical Apollo mission only lasted a bit over a week, the Manned Venus Flyby was going to take a full year! Without the wet workshop you'd have three unwashed flatulating twitchy astronauts cooped up in a space no bigger than three telephone booths for a year (ask your grandparents, kids; it's a total of 6.2 m3). It would not be long before one or more of them goes postal and there will be blood and body parts floating round the comand module. The Space Safety and Human Performance equation says for a year-long mission the bare minimum habitable volume is 32 cubic meter per crew, the command module has only 2 m3 each. That is way below the "eventual psychotic break" level.


The study selected a thirty day launch window from October 31 to November 30, 1973. The mission duration will be about 400 days with the October 31 launch date, decreasing down to about 360 days at November 30th. The Terra entry velocity varies from 13,700 m/s to 13,800 m/s. After mulling over the implications, the study chose the October 31 launch date.

Mission Characteristics
Earth departureOctober 31, 1973
Injection velocity
(from 185 km orbit)
3,930 m/s
Outbound leg123 days
Venus encounterMarch 3, 1974
Periapsis altitude6,190 km
Inbound leg273 days
Earth return dateDecember 1, 1974
Entry velocity13,700 m/s
Launch azimuth72° — 108°
First launch window1305 — 1738 EST
Second launch window1855 — 2327 EST
BAD SF IDEAS IN REAL LIFE

Many readers may find the plots of some SF novels deeply implausible. “Who,” they ask, “would send astronauts off on an interstellar mission before verifying the Go Very Fast Now drive was faster than light and not merely as fast as light? Who would be silly enough to send colonists on a one-way mission to distant worlds on the basis of very limited data gathered by poorly programmed robots? Who would think threatening an alien race about whom little is known, save that they’ve been around for a million years, is a good idea?”

Some real people have bad ideas; we’re lucky that comparatively few of them become reality. Take, for example, a proposal to send humans to Venus. Not to land, but as a flyby.

After the Apollo program had landed humans on the Moon, the obvious question was, “What next?” Some proposals were carried out: Skylab space station1; U.S.-Soviet cooperation in orbit. Other proposals were binned because there was no money for such things or because they were obviously stupid.

The Manned Venus Flyby would have been both expensive and stupid.

The mission would have re-purposed Apollo-era equipment for a far more ambitious journey. Rather than a week or so in space, the astronauts would have spent more than a year on a slow cruise past Venus. Rather than expect the astronauts to spend this time in the cramped conditions of a Command Module and LEM, the Manned Venus mission would have converted a hydrogen tank into living quarters once it had served its original purpose and was no longer filled with liquid hydrogen. The interplanetary vehicle that resulted would have been quite impressive even by modern standards, let alone those of the Apollo era.

Of course, the mission was not intended to land on Venus. If you could get down to the surface (or what passes for a surface on Venus) you couldn’t get back up to the spacecraft. Venus is nearly as massive as Earth and its escape velocity is not much lower. Without in-situ resource utilization, the fuel demands for an Earth > Venus’ surface, Venus > Earth mission would have been intractable.

Not to mention the fact that Venus is a hell planet. The lower reaches of its dense poisonous atmosphere are hot enough to melt lead. Sending astronauts down to the surface would merely have tested how close to the surface they could get before the ambient conditions killed them.

Happily, that was not what was proposed.

Instead, the astronauts would have been sent on a flyby that would last from late October of 1973 to early December of 1974. The encounter with Venus would have occurred in early March 1974. While close to Venus, the astronauts would collect a wide variety of data about that world and its interplanetary neighborhood (which includes Mercury). They would also give the U.S. a reason to wave the flag and boast of achieving the first interplanetary manned mission. USA! USA!

If I sound unappreciative of this bold plan, you are right. I think it’s cockamamie. Because:

The mission does not do anything robotic missions could not do more cheaply. While humans are a lot more flexible than machines, they’re difficult and expensive to feed and protect. Not only do you need to pay for the fuel to toss humans across space, you need to pay for everything needed to keep them alive as well. Note that what we have actually done is send robots to explore Venus and Mars, as well as other worlds.

(But, you say, we would learn so much about how to feed and protect crew, which we cannot do without crewed missions. Hey, we’re still working on keeping humans alive on space stations safely below the Van Allen belt. That’s enough for now.)

An even more important reason why the Manned Venus Flyby would have been a bad idea (even if Congress had been inclined to fund it—which it was not) is that the interplanetary environment was more challenging than folks in the ‘70s understood. The Apollo-moon-mission-era solution to spacecraft radiation shielding was to hope very, very hard that no major solar storm would occur on the way to and from the Moon. As it turned out, this worked—which is good because a major storm would have definitely killed the Apollo astronauts. Hoping for good space weather would have been a no-go for a four-hundred- day mission, so a Manned Venus Flyby would have required a radiation shelter, yay. What the proposers could not have known, however, is that their mission would have run into a coronal mass ejection in July 1974, one major enough to overwhelm any currently implementable shelter2. This would have been fatal for the astronauts.

While this would at least have provided a distraction from Watergate, President Nixon probably wouldn’t have found it pleasant to explain to the press just how the U.S. lost a crew in deep space.

So the next time you set down a science fiction novel and think “nobody would be dumb enough to send people off on an obvious one-way trip to certain death”3, just remember that at one point in recent history, sending a collection of astronauts off to be crisped like KFC chicken seemed like a reasonable idea.

Footnotes

     1: As evidenced by the graphic linked above, Skylab’s launch marked the beginning of a multi-decade period in which there was always at least one space station (usually Russian) in low Earth orbit, in a chain that runs Skylab, Salyut 3 through 7, Mir, the ISS, Genesis 1 and 2, and Tiangong 1 and 2. Not all of these were actually crewed but still, for the lives of most people now living on Earth, there’s always been a space station in orbit.

     2: The initial proposal predated the OSO 7 spacecraft’s confirmation of CMEs (coronal mass ejections) in the early 1970s, so it’s not surprising the original vision didn’t have an effective contingency plan.

     3: A surprising amount of Canadian history was shaped by the British decision to appoint John Franklin, a man whose previous exploits included nearly drowning in a river, and losing eleven of twenty men in his Coppermine expedition, as leader of an expedition to search for the Northwest Passage. Poor planning and Victorian-era canning technology afflicted his mission with scurvy, TB, hypothermia, starvation, lead poisoning, and what gamers call “a total party kill.” Franklin’s final expedition didn’t directly provide any information to the Empire, what with the whole being-too-dead-to-report-their-findings thing. Silver lining: the people who tried to track the vanished expedition did provide useful info. Oh, and we learned that you should not put John Franklin in charge of arctic expeditions.

APOLLO’S HAIL MARY PASS

What it was: A proposed post-Moon landing manned mission using Apollo hardware. It would have launched during a good alignment of Earth and Venus in November 1973 and taken three astronauts on a flyby of the planet Venus, returning to the Earth 13 months after launch.

A later variation of the mission ambitiously suggested using a better conjunction in 1977 to visit Venus and Mars on an outbound leg and Venus again on the Earth-return leg, however most of the work done considered the shorter Venus flyby.

Details: By the mid-1960s NASA was well aware that if they successfully completed the Apollo moon landings they would probably face a severe decline in budget for the manned space program. In the hopes of proving their ongoing worth they developed a few different post-Apollo proposals using evolutionary versions of the Apollo hardware, including plans for a manned lunar base, space stations, and planetary exploration. The latter two of these goals were at first grouped under the name Apollo X, and then became the Apollo Applications Program (AAP).

By far the most ambitious of the AAP missions was a manned flyby of the planet Venus. After two preliminary missions in Earth orbit to test the technology, a Saturn V launch would lift an Apollo Command Module into orbit. As in a typical mission, the first two stages of the rocket would be jettisoned. However the uppermost stage, the Saturn IVB, would be kept and drained of any remaining propellant. Using gear stored where the Lunar Landing Module would have been placed in a Moon mission, the astronauts would then rig it as a habitation module.

The resulting 33-meter-long spacecraft would leave Earth orbit on October 31, 1973 and travel towards Venus for 123 days. There would be a flyby on March 3, 1974. The craft would have been aimed to pass Venus as close as 6200 kilometers above the surface (one planet radius) very quickly—orbital mechanics would have it moving relative to Venus at a clip of 16,500 kilometers per hour—crossing the lit side of the planet. A sidescan radar would map the portion of the planet they could see as they flew by, and the astronauts would perform spectroscopic and photographic studies.


After that burst of activity the MVF craft would then return home, taking 273 days more to loop out to 1.24 AU from the Sun on a hyperbolic trajectory and eventually swing back to Earth. The astronauts’ landing on Earth would happen on December 1, 1974—total mission time would be 396 days. The Triple Flyby variant would have taken more than 800 days starting in 1977.

When not at Venus, the MVF astronauts would have studied the Sun and solar wind as well as making observations of Mercury, which would be only 0.3 AU away two weeks after the Venus flyby. To keep them occupied otherwise their habitation capsule would have been outfitted with a small movie screen (to show 2 kilograms of movies allowed), and a “viscous damper exercycle/g-conditioner”. The crew would also be allowed 1.5 kilograms of recorded music, 1 kilogram of games, and 9 kilograms of reading material. Hopefully they would choose wisely.

What happened to make it fail: The MVF was part of the Apollo Applications Program, and the AAP was killed dead on August 16, 1968 when the House of Representatives voted to cut its funding from US$455 million to US$122 million. President Johnson accepted this as part of a larger budget deal that kept NASA’s near-term goals safe, though even at that the agency’s entire budget dropped by 18% between 1968 and 1969. The only AAP mission to survive was Skylab.

What was necessary for it to succeed: It’s tough to get this one to work as it’s difficult to see any advantage to sending people on this mission. Mariner 5 had already flown by Venus in 1967 and NASA was able to send a robotic orbiter as part of the Pioneer 12 mission in 1978, just a few years after MVF would have flown.

Even the many probes that the MVF would leave behind at Venus had no obvious connection to the manned part of the mission; it would have been easier to send an unmanned bus of similar size and drop the probes that way. There would be no need then for heavy food, water, or air, or the space for people to move around. And unlike the manned mission there would be no need to bring the bus back, greatly reducing the mission’s difficulty. About all the manned mission had going for it was an opportunity to see what kind of effect a year in microgravity would have on humans, and that could just as easily be determined using a space station in low Earth orbit.

On that basis we also need to be aware that Congress asked hard questions about the purpose of NASA’s manned Mars mission plans in the late 1960s and were hostile to all of them. If Mars wasn’t going to get any money, it’s hard to see what could influence them to fund a mission to Venus.

Finally it needs to be pointed out that no matter even if the MVF launched, nature itself probably had this mission’s number. We didn’t have a very good understanding of the Sun at that time, having only observed one solar cycle from above the atmosphere when the flyby was proposed in 1967. While the launch window was deliberately chosen to be near a solar minimum, and the flyby craft was to have a radiation lifeboat in the equipment module, the mission would have run into an unforeseen natural event on the way back to Earth.

On July 5-6, 1974 the Earth was hit by a big coronal mass ejection (CME), a storm of electrons and protons thrown off of the Sun. People down on Earth were protected by the planet’s magnetic field, as usual, but the astronauts coming back from Venus wouldn’t have been so lucky. Their line to the Sun was several degrees off from the Earth’s (at the time they would have actually looped out past Earth as their trajectory slowly took them back home), but CMEs can cover quite a bit of space. Had the mission actually flown, the astronauts on-board may well have died of radiation sickness after being hit with more (and more energetic) solar protons than their spacecraft was built to handle.

The saving grace here is that coronal mass ejections were discovered in 1971, so the initial plan probably would have been called off rather than risk casualties, or at least be reconfigured to give the astronauts the protection the 1967 plan failed to give them.

MANNED VENUS FLYBY

Manned Venus Flyby” was a 1967–1968 NASA proposal to send three astronauts on a flyby mission to Venus in an Apollo-derived spacecraft in 1973–1974, using a gravity assist to shorten the return journey to Earth.

Apollo Applications Program

NASA considered a manned flyby of Venus in the mid-1960s as part of the Apollo Applications Program, using hardware derived from the Apollo program. Launch would have taken place on October 31, 1973, with a Venus flyby on March 3, 1974 and return to Earth on December 1, 1974.

Background

The proposed mission would use a Saturn V to send three astronauts to fly past Venus in a flight which would last approximately one year. The S-IVB stage would be a 'wet workshop' similar to the original design of Skylab. In this concept, the interior of the fuel tank would be filled with living quarters and various equipment that did not take up a significant amount of volume. The S-IVB would then be filled with propellants as normal and used to accelerate the craft on its way to Venus. Once the burn was complete, any remaining propellant was vented to space, and then the larger fuel tank could be used as living space, while the smaller oxygen tank would be used for waste storage.

Only so much equipment could be carried in the hydrogen tank without taking up too much room, while other pieces could not be immersed in liquid hydrogen and survive. These sorts of systems would instead be placed in the interstage area between the S-IVB and the Apollo Command/Service Module (CSM), known as the Spacecraft-LM Adapter (SLA), which normally held the Apollo Lunar Module on lunar missions. To maximize the amount of space available in this area, the Service Propulsion System engine of the CSM would be replaced by two LM Descent Propulsion System engines. These had much smaller engine bells, and would lie within the Service Module instead of extending out the end into the SLA area. This also provided redundancy in the case of a single-engine failure. These engines were responsible for both course correction during the flight as well as the braking burn for Earth re-entry.

Unlike the Apollo lunar missions, the CSM would perform its transposition and docking maneuver with the S-IVB stage before the burn to leave Earth orbit rather than after. This meant the astronauts would fly 'eyeballs-out', the thrust of the engine pushing them out of their seats rather than into them. This was required because there was only a short window for an abort burn by the CSM to return to Earth after a failure in the S-IVB, so all spacecraft systems needed to be operational and checked out before leaving the parking orbit around Earth to fly to Venus.

Precursors to the Venus flyby would include an initial orbital test flight with an S-IVB 'wet workshop' and basic docking adapter, and a year-long test flight taking the S-IVB to a near-geostationary orbit around the Earth.

Scientific objectives

The mission would measure:

  • Atmospheric density, temperature and pressure as functions of altitude, latitude and time.
  • Definition of the planetary surface and its properties.
  • Chemical composition of the low atmosphere and the planetary surface.
  • Ionospheric data such as radio reflectivity and electron density and properties of cloud layers.
  • Optical astronomy - UV and IR measurements above the Earth's atmosphere to aid in the determination of the spatial distribution of hydrogen.
  • Solar astronomy - UV, X-ray and possible infrared measurements of the solar spectrum and space monitoring of solar events.
  • Radio and radar astronomy - radio observations to map the brightness of the radio sky and to investigate solar, stellar and planetary radio emissions; radar measurements of the surface of Venus and Mercury
  • X-ray astronomy - measurements to identify new X-ray sources in the galactic system and to obtain additional information on sources previously identified.
  • Data on the Earth-Venus interplanetary environment, including particulate radiation, magnetic fields and meteoroids.
  • Data on the planet Mercury, which will be in mutual planetary alignment with Venus approximately two weeks after the Venus flyby
From the Wikipedia entry for Manned Venus flyby

Benton Spaceship Discovery

This is from Spaceship Discovery – NTR Vehicle Architecture for Human Exploration of the Solar System by Mark G. Benton, Sr. (AIAA 2009-5309) 2009. Available here, paper labeled "MarkBentonSpaceship Discovery (SSD) Paper (AIAA-2009-5309)"

Mr. Benton also invokes the spacecraft Discovery from 2001 A Space Odyssey. The state-of-the-art has advanced to the point where the fictional movie spacecraft could be built in reality. This is a modular design built around multiple bi-modal nuclear thermal rockets. The design also includes for types of landers for a variety of missions. High-energy Mars and Jupiter missions are supported with dual strap-on NTR boosters.

The idea is that a modular design capable of being configured for a wide variety of missions would kickstart human exploration of the solar system.

Seven Design Reference Missions (DRM) were created in order to set the design requirements:

  • DRM 1: Shakedown mission to Luna
  • DRM 2: Mars Exploration Mission
  • DRM 3: Mars Colony Resupply Mission
  • DRM 4: Asteroid Ceres Exploration Mission (not designed yet)
  • DRM 5: Callisto Exploration Mission
  • DRM 6: Ganymede Exploration Mission
  • DRM 7: Ganymede Plus Callisto Exploration

Four types of landers were designed:

  • RM: Crew Reentry Module for Terra Return
  • LM1: Vacuum Exploration Lander for Luna, Callisto, and Ceres
  • LM2: Atmospheric Exploration Crew Lander for Mars
  • LM3: Atmospheric Cargo Lander for Mars

The nuclear engine has a specific impulse of about 950 seconds, as opposed to a pathetic 475 seconds for chemical. Nuclear can handle the 20 to 30 km/s delta V required for Ceres, Jupiter, and Saturn missions with a reasonable mass ratio. With chemical engines you might as well forget it.

Since it uses nuclear propulsion it does not have to use risky aero-capture maneuvers. Mars' aero-capture atmosphere can vary from 70% to 200% in a single day. Jupiter has such intense gravity that the transit velocity would be too high.

Design can use strap-on nuclear boosters for those high-energy sort-transit-time Mars and Jupiter missions. It has a backup abort propulsion system allows the crew to escape at multiple points in the mission. The cluster of NTR engines provides redundancy in case one of them fails. The hab module has galactic cosmic ray shielding composed of liquid hydrogen and water tanks. However additional radiation shielding would be required to visit Ganymede. The hab module even has spin gravity. The bi-modal NTR provides electrical power.

The basic Spaceship Discovery is a stack composed of an Engineer Module (EM), four Main Propellant Core Tanks (CT), Service Module (SM), and Crew Module (CM). It is customized for a mission by the addition of a Docking Module (DM), Terra Reentry Module (RM), Planetary Landers (LM1, LM2, LM3), and Propellant Drop Tanks (DT). A strap-on booster is composed of one EM, two CT, and up to 12 DT.


Crew Module (CM)

The standard configuration can accomodate a crew of six, with the strap-on boosters there can be only four. Assumes consumables of Oxygen 1.0 kg/person-day, Dried Food 1 kg/person-day, Food Water 2 kg/person-day, Drinking Water 1 kg/person-day, Hygiene/wash water 6 kg/person-day. Life support system is assumed capable of recovering 75% of oxygen from carbon dioxide and 90% of drinking and hygiene water. This includes roasting solid wastes to recover the water. The dry remains are then jettisoned prior to start of burns to reduce ship mass.

The food is stored in two locations: a zero-g aft compartment with 66 m3 volume and an artificial gravity compartment with 78 m3 volume. On high-energy 3.9-year four-crew missions this provides 9.3 m3 per person-year. On low-energy 2.67-year six-crew missions this provides 9.1 m3 per person-year.

The crew module inflates after launch into an oblate spheroid, the shell cures and hardeneds in the vacuum of space. The non-rotating corew is composed of graphite-epoxy composites and is the primary structural load path.

The radiation shielding is 4 gm/cm2 of liquid hydrogen (57.7 cm thick layer) to protect from galactic cosmic rays (GCR) and solar particle events (SPE). The part of the crew module which is the roof and floor of the uninflated module uses the hygiene water tanks and other assorted equipment for radiation shielding instead of liquid hydrogen. For a four year mission the cumulative radiation dose would be about 140 centi-Sieverts (1.4 Sieverts) which is below the lifetime limit for 45 year old males and 55 year old females. The liquid hydrogen mass is just enough for the final main propulsive burn.

The crew module centrifuge spins at 4.0 rpm to provide 1/6 g in the crew sleeping, exercise, and recreation spaces (centrifuge radius about 9.2 meters).

The forward end of the crew module has a docking port. The aft end has a personnel airlock, four docking ports, four deployable solar arrays (provides electrical power in abort mode), and high-gain mast antennas.

The crew module is a compromise between adequate habitat volume, minimum artificial gravity centrifuge radius, and mass of radiation shielding due to surface area.


Docking Module (DM)

On the crew module's forward docking port is installed the Docking Module (DM). This provides an airlock, personnel hatch, and five docking ports (up to four landers and one Terra reentry module). The Docking Module is jettisoned after the landers are deployed (to reduce ship mass) and the reentry module attached directly to the forward docking port.


Service Module (SM)

The Service Module structure is composed of graphite-epoxy composites. It houses liquid oxygen and liquid nitrogen tanks (consumables for life support system), gaseous helium tanks (propellant pressurization, centrifuge cavity purge, coolant for Very Low Boiloff System), forward RCS propellant tanks. It also has the 5.15 kN RCS thrusters and the 76.2 kN abort propulsion system (APS) engine. Both use storable hypergolic propellants (monomethyl hydrazine, MMH, and nitrogen tetroxide, N2O4) since they may have to be stored for years before abruptly needed.

The many cryogenic liquid hydrogen propellant tanks have to be kept cool or all the propellant will be lost to boil-off. The Very Low Boiloff coolant system including the heat radiators is also located in the service module. Liquid hydrogen tanks include the Propellant Core Tanks (CT), Drop Tanks (DT), the crew module radiation shield, and propellant tanks in all attached landers. Deployable shades do their best to shield the many propellant tanks from the thermal flux from the heat radiators.

Abort is performed in case of multiple nuclear engine failure. The APS has 76.2 kN of thrust and from 0.061 to 0.278 km/s of delta V. Additional delta V is available from attached landers. Before abort, everthing aft of the service module is jettisoned. Each lander module is fired in sequence then discarded. The docking module is discarded with the last lander. When the remainder of ship approaches Terra, the crew tries to do an unbraked reentry in the reentry module (rolling the dice to see which they run out of first: heat shield or velocity).


Engineering Module (EM)

Engineering module has a trio of bi-modal gimaled nuclear thermal engines, for redundancy. 178.4 kN of thrust each, for a total of 535 kN. Specific impulse of 950 seconds (exhaust gas temperature of 2,900 K). In electrical power generation mode they use closed Brayton cycle (CBC) turbo-alternator systems with recuperation. 76 kilowatts electrical each for a total of 200 kWE. After burn engines are cooled down with extra propellant. Excess heat is proportional to engine burn time and fission product buildup. Thrust from cooling has a specific impulse of 633 s (2/3 operating specific impulse). Once the cores cool enough the Brayton units can take over cooling duties, sending the heat to heat radiators instead of power generation gear. These heat radiators are located just forwards of the engines, along with the deployable shades that protect the cryogenic core tanks from engine and radiator heat.

As always the deadly radiation flux directed at the crew module is combated with a combination of distance and shadow shields. The crew module has a separation distance from the nuclear engines of 115 meters. The shadow shields are composed of layers of tungsten (gamma shielding) and lithium hydroxide (neutron shielding).

The engineering module also houses the aft RCS thrusters, MMH and N2O4 propellant and pressurization tanks. However the propellants absorbing all that radiation is a matter of concern.


Main Propellant Core Tanks (CT), Drop Tanks (DT)

Both of these types hold the liquid hydrogen propellant (LH2) for the nuclear engines. The main differences are:

  • Core tanks form the ship's backbone and thrust frame, so they are stronger. Drop tanks are more flimsy. Core tanks form the backbone with a skirt structure and tank-to-tank fittings. Drop tanks lack that.

  • Each drop tank has its own internal cryo-cooler, refrigeration unit, and heat radiator. Core tanks cannot have heat radiators because they might be incased in a clutch of drop tanks. So core tanks are refrigerated by the Very Low Boiloff System in the service module.

  • Core tanks have a donut-shaped (torus) LH2 propellant feed plenum at the base. The bottom of the tank has a propellant pipe connecting to the plenum. Any drop tanks attached to this core tank also has a pipe connecting to the plenum. The plenum connects to the three Main NTR LH2 feeds on the skin of the core tank. This core tank's feeds connect to the feeds of the core tank immediately above and below. The tank at the base connects each feed to one of the nuclear engines.

  • Core tanks are integral parts of the spaceship. Drop tanks are meant to be dropped when they run empty.

  • Core tanks have a length of 22.5 meters, drop tanks have a length of 21.7 meters

Both carry 43.05 metric tons of liquid hydrogen propellant each. Both are 7 meters in diameter. It is assumed that both suffer boil off losses of 0.05% of the LH2 per month. Both are built out of graphite-epoxy composites. Both have internal helium pressurization tanks.


Strap-On Boosters

This is a method of multistaging. A single strap on booster is composed of:

  • One engineering module (with three nuclear engines)
  • Two core tanks
  • Up to twelve drop tanks

A Discovery spaceship with no strap-on boosters has a maximum delta V of about 20 km/s, because other factors mandate the initial thrust to weight ratio cannot be less than 0.05. This delta V is adequate for DRM 1 and DRM 2, but not enough for any of the other design reference missions. Strap-on boosters give the extra delta V needed. Unfortunately due to other constraints, a spaceship using strap-on boosters can only carry 4 crew instead of 6.

For the higher DRMs a pair of strap-on boosters are required. The boosters are used during the Terra escape burns: Trans-Mars Injection or Trans-Jupiter Injection. During the burn the booster will cross feed so their tanks supply propellant to the core engines as well as the strap-on engines.


Crew Reentry Module (RM) for Terra Return


Airless Exploration Lander for Luna, Callisto, and Ceres (LM1)


Mars Exploration Lander Modules - Crew Lander (LM2) and Cargo Lander (LM3)


Design Reference Mission 2 (DRM 2) – Mars Exploration


Design Reference Mission 3 (DRM 3) – Mars Colony Resupply


Design Reference Mission 5 (DRM 5) – Callisto Exploration


Design Reference Mission 6 (DRM 6) – Ganymede Exploration


Design Reference Mission 7 (DRM 7) – Ganymede Plus Callisto Exploration


Comparison of Design Reference Missions

Boeing IMIS

RocketCat sez

Now this is audacious. Boeing sure thought big back in 1968.

Yes, there were quite a few proposed Mars missions back then. Many of them used multi-staging, discarding tanks and engines to increase the mass-ratio.

But none of them had stages with Freaking NERVA atomic engines, tossing five nuclear reactors glowing with radioactive death into eccentric solar orbits. They'll stop emitting dangerous radiation in only a few thousand years.

On the plus side the relatively huge specific impulse of the NERVAs means this monster spacecraft can boost more than one hundred metric tons of payload to Mars; including a huge habitat module, one of those workhorse North American Rockwell Mars landers, a pallet of scientific experiments, and re-entry vehicle to return the crew to Terra.

The Boeing Integrated Manned Interplanetary Spacecraft (IMIS) is a three stage spacecraft with nuclear thermal rockets.

Most of the diagrams here are from Integrated Manned Interplanetary Spacecraft Concept Definition. Volume 1 - Summary Final Report, with further data from Volume 4 - System Definition Final Report.

The False Steps blog calls this project NASA’s Waterloo, due to an utter disconnect between what NASA thought they should get in funding and what everyone else in the government was willing to give them.


Overview

  • Hot Pink: Primary engines - NERVA solid-core nuclear thermal rockets
  • Light Blue: Secondary engines - FLOX-methane chemical course correcting engines
  • Red: Propulsion Module 1 (PM-1). Three NERVA-propellant tank assemblies. Stage used for Terra Orbit Departure (~5,100 m/s)
  • Orange: Propulsion Module 2. One NERVA-propellant assembly. Stage used to brake into Mars orbit (~5,300 m/s)
  • Yellow: Propulsion Module 3. One NERVA-propellant assembly. Stage used for Mars Orbit Departure (~5,800 m/s)
  • Green: Payload. Mission Module (habitat module), Mars Excursion Module (Mars Lander), Experiments pallet, Earth Entry Module (reentry vehicle)

In the Boeing report they call the payload module the "spacecraft", the string of five engine modules is the "space acceleration system", and the entire thing is the "space vehicle"

It is oriented so that "down" is towards the nose, since the spacecraft is a Tumbling Pigeon.


Mission

Spacecraft is assembled in orbit. Just prior to trans-Martian injection, PM-1 jettisons its meteor shielding to reduce excess mass. PM-1 burns with all three NERVA engines to perform trans-Martian injection (ΔV 3,645 to 3,989 m/s) and is then jettisoned. The jettison path is designed so that PM-1 does not impact Mars nor does it stay too close to the spacecraft. PM-1 travels aimlessly in an eccentric solar orbit as a radiation hazard for several thousand years.

During the transit to Mars, PM-2 performs three midcourse corrections using its FLOX-methane secondary propulsion system. These are done at 5 days after leaving orbit, 20 days later, and 20 days before arrival at Mars.

On Mars approach the PM-2 meteor shielding and secondary propulsion system is jettisoned. The PM-2 NERVA engine burns for Mars capture (ΔV 2,568 to 2,947 m/s), placing spacecraft in a high Mars orbit. The PM-2 stage is jettisoned.

The PM-3 FLOX-methane secondary propulsion system puts the spacecraft into a lower 1,000 kilometer orbit, putting some distance between itself and the dangerously radioactive PM-2 stage in high orbit. The PM-2 stage will be a radiation hazard for a few thousand years.

The crew spends 2 to 5 days surveying Mars to locate a safe-but-interesting landing site. They also perform orbital experiments, deploy probes, and prep the Mars Excursion Module.

Three of the six crew board the Mars Excursion Module and lands on the pre-determined landing site (or close by if the site turns out to be full of giant dagger-like rocks or something).

The Mars team stays for 30 days planetside, exploring Mars. Meanwhile the orbital team continues the orbital experiments, monitors the planetary operations, and do maintenance on the spacecraft.

At the end of 30 days the MEM ascent vehicle delivers the Mars team and their collection of Mars rocks back to the orbiting spacecraft. After the ascent vehicles rendezvouses with the spacecraft and transfers the crew, it is jettisoned.

The PM-3 meteor shielding and secondary propulsion system is jettisoned. The PM-3 NERVA engine burns to put the spacecraft into trans Earth injection (ΔV 4,969 to 5,811 m/s). PM-1 is then jettisoned.

During the trip home, the FLOX-methane engine on the Mission Module perform three mid-course corrections.

One day before Earth arrival, the crew and the Mars samples move to the Earth Entry Module. They then leave the Mission Module, which does a final burn to move it out of the way. The Earth Entry Module aerobrakes to land on Earth (entry velocity 16,200 to 18,400 m/s).

Total mission duration is from 460 to 540 days. Total ΔV is 11,400 to 12,400 m/s


Propulsion Module

Propulsion Module
EngineNERVA
Engine mass
incl. thrust frame
w/o radiation shield
11,580 kg
Engine Length12.2 m
Engine Nozzle dia4.12 m
Thrust868,000 Newtons
Specific Impulse850 sec
Propellantliquid hydrogen
Tank Diameter10.6 m
Tank Length35 m
Propellant mass175,000 kg
Propellant Volume2,590 m3
Wet Mass227,000 kg

These look suspiciously like the NASA reusable nuclear shuttle.

The outer shell serves as a load-carrying structure during Earth-launch, and as meteoroid shielding during the mission. It is split into four segments secured by hoop straps. The straps are severed just prior to engine ignition, allowing the meteor shielding to drop off.

The engine is surrounded by a two-layer interstage shell. The outer interstage is a load-bearing structure for Earth launch. It is jettisoned after reaching Earth orbit. After that the inner interstage is the load bearing structure for mission flight loads. The interstage shell is 13.1 meters long, about 0.9 meters longer than the engine.

The module has a 20 cm fuel transfer line used to transfer propellant between modules during the mission.

The female and male docking modules allow the propulsion modules to be connected like Legos.

The propellant tank, thrust frame and engine support are constructed of aluminum (low mass and doesn't crack at liquid hydrogen temperatures). The tank dimensions were chosen so the diameter and the filled mass would not exceed the capability of a Saturn V launch vehicle.

The base of the tank rests on the thrust frame. This is a cross-beam structure with the propellant tank attached to the top and the NERVA engine gimbal attached to the bottom.

NERVA Engine
TypeSolid-core NTR
Engine mass
less rad shield
and thrust frame
25,540 lbs
Radiation shield mass1,940 lbs
Thrust Frame mass1,050 lbs
Specific Impulse850 sec
Reactor power4,000 MW
Engine Thrust868,000 Newtons
Propellant Mass Flow239 lb/sec

The study figured that the crew would be safe from the radiation emitted by the reactors in PM-1 and PM-2, mostly due to the shielding provided by the propellant in PM-3 (right below the habitat module). But radiation becomes a problem when PM-3 starts burning the PM-3 propellant, essentially removing the radiation shielding.

The study showed that there was a trade-off between the amount of mass in a beryllium oxide radiation shadow shield on top of the PM-3 NERVA engine and the amount of mass in a water shield around the biowell on deck 3. But it did not come to any firm conclusion. You can read the rambling argument in Volume 4 - System Definition Final Report on pages 194 through 199.


Mission Module

CREW COMPARTMENT

The crew compartment provides a pressurized shirt- sleeve environment for the crew and storage for equipment which needs a thermal or pressure environment or is expected to require maintenance. Atmosphere within the crew compartment is nominally 7 psia (48kPa) O2/N2, 70°F and 50% relative humidity. The crew compartment consists of a 17.8-foot (5.4m) cylinder, 22 feet (6.7m) in diameter (decks 2 &3), joined at both ends by hemispherical bulkheads (decks 1 & 4). A meteoroid bumper surrounds the cylindrical section of the crew compartment (decks 2 &3). Overall length of the crew compartment is 39.8 feet (12.1m) which provides a total volume of approximately 12,250 cubic feet (347m3). Total pressurized volume within the crew compartment is estimated to be 10,000 cubic feet (283m3) for 500-day class missions with the free volume (major areas unoccupied by equipment) 5400 cubic feet (153m3) or 900 cubic feet (25.5m3) per man (which is ample). A surface area of approximately 1200 square feet (112m2) is provided by the cylindrical portion of the crew compartment.

The internal arrangement of the crew compartment results from having to contain within the selected 22-foot (6.7m) diameter pressure compartment a floor area requirement of approximately 1400 square feet (130m2) and ceiling height of 7 feet (2.1m) in order to provide sufficient volume for equipment and men. As a result, the crew compartment consists of four separate levels of activity. Each level is designed to include those crew operations or equipment operations of a similar nature. The levels have also been located to minimize the interface or distance between levels of similar activities. An example is the above/below arrangement of the two levels which include all areas and equipment associated with spacecraft operations and crew living quarters. Equipment and cabinets within the crew compartment and located near the walls are attached in place and do not have provisions for removing or hinging the entire cabinet to expose walls for puncture repair caused by meteoroids. Previous inhouse studies such as Manned Orbital Laboratory have indicated a greater reliability benefit can be achieved by using a weight equal to the hinging mechanisms in the meteoroid shield itself.


DECK 1

Activities of a relatively quiet nature are located on Deck l. In general, this deck includes the sleeping quarters, dispensary, and personal care facilities. Each crewman is provided with a separate room to be used for sleeping and stowage of personal hygiene supplies such as clothes, cleaning pads, and personal care items. Cabinet space is also available for other equipment associated with the mission module. The rooms also provide solitude for crewmen if desired, and allow a crewman to be isolated should the need exist. Approximately 110 cubic feet (3.1m3) of free volume is provided per room. Included within the dispensary is the necessary equipment for crew psychological/physiological monitoring, medical/dental equipment and supplies, and physical conditioning equipment for the cardiovascular system and musculoskeletal system of the body. Personal care facilities include a zero-g shower and waste management system (toilet). Adjacent to the waste management system is the urine water recovery unit. After processing, this water is transferred to holding tanks on Deck 2. Located in the upper portion of Deck l is a pressure hatch leading to the EEM (Earth Entry Module, reentry vehicle) transfer tunnel. A centrally located, 36 inch (0.91m) diameter hatch leads to Deck 2.


DECK 2

Activities of a relative high intensity are located on Deck 2. In general, the activities include the command/control center, combination food storage/preparation area, and recreation area. The command/control center includes the necessary displays and controls to monitor and control all subsystem operation, environment parameters, and vehicle operations such as attitude changes, rendezvous, and dockings. The control center is occupied at all times. The food storage/preparation area includes freezer, hot water provisions, and food storage cabinets for missions greater than 500 days. Dining facilities are also included in the area. Another section of this area contains the remainder of the water management system consisting of the wash water/condensate water recovery unit and a 2-day water supply. Water for crew consumption comes to this supply from the makeup water supply located on the third deck. Storage for wash pads occupy the final bay in this area. The remainder of Deck 2 is used for recreation, conference room, and storage for spares (redundancy). Dividing the recreation area and food storage/preparation area is a bay for electronic equipment with the most significant being the control moment gyros (CMG) of the attitude control subsystem. Located in the center of the floor of this level is the pressure hatch leading to the radiation shelter on Deck 3. Also located in the floor are nonpressure hatches which allow access to the equipment bays of Deck 3.


DECK 3

The major features of the third deck are the combination radiation shelter/emergency pressure compartment and equipment bay. Height of this deck is approximately 10 feet (3.1m) rather than 7 feet (2.1m) as for the other decks due to the design feature of the radiation shelter. The radiation shelter consists of an inner compartment 10 feet (3.1m) in diameter and 7 feet (2.1m) high which also serves as the emergency pressure compartment should the remainder of the crew compartment become uninhabitable for short periods of time. A total volume of 600 cubic feet (17.0m3) is provided by the radiation shelter with approximately 60 cubic feet (1.7m3) of free volume available per crewman. The shelter also provides quarters for the crew during periods of high radiation. These periods include passing through the Van Allen belt anomaly while in Earth orbit; during the firing of each nuclear propulsion module, particularly during departure from Earth as the space vehicle may pass through the heart of the Van Allen belt, and the firing of PM-3 (the nuclear engine module directly adjacent to the crew quarters) when a minimum of hydrogen is between the crew and Nerva engine; and during major solar flares which may last up to 4 days. Because the shelter may be occupied for extended periods of time and during nuclear propulsion firings, it is necessarily provided with sufficient displays and controls to enable the crew to continue space vehicle operations. A 4-day emergency food, water, and personal hygiene supply is provided within the shelter as well as separate atmosphere supply and atmosphere control loops. Each crewman is provided with a storage compartment, which contains his pressure and emergency provisions. Should the crew compartment become uninhabitable, all crewmen transfer to the shelter and don pressure suits. A repair team can then be sent out to correct the malfunction. The final item housed in the shelter is the photographic film used in the experiment program. This location has been selected as it provides the maximum amount of radiation shielding at no additional weight penalty.

The bulk of the radiation protection for the shelter is provided by a 20 inch (0.5m) thick combination food/waste storage compartment. This storage compartment contains the initial 500-day supply of food and surrounds the entire shelter providing approximately 26 lb/ft2 (137kg/m2) of shielding. Further discussion of the radiation protection analysis is presented in Section 4.2.1.4. Food stored around the walls of the shelter is reached from the equipment bay. Floor panels are removed in the second deck to reach the food above the shelter, while ceiling panels of the fourth deck are removed to reach the food located beneath the shelter. As food is removed, the vacated area is filled with waste matter in order to maintain a nearly constant mass.

The equipment bay of this deck includes a storage area extending 2 feet (0.6m) inward from the outside wall and around the entire periphery. A passageway is provided between the equipment and the food storage compartment of the radiation shelter. The passageway is between 24 to 30 inches (0.6m to 0.8m) wide which should provide sufficient space for maintenance operations or removal of supplies even while operating in a pressure suit. Housed in the storage area are three 24 inch (0.6m) diameter water containers and positions for three other containers to be used for missions between 500 to 1000 days. Also included in the area is the major portion of the environmental control system equipment such as electrolysis unit, Bosch reactor and atmosphere control units, storage for spares and provisions for food, and spares storage for missions beyond 500 days.


DECK 4

The fourth deck of the crew compartment is comprised almost entirely of laboratories associated with the experiment program. These labs contain the necessary equipment to perform certain experiments, control the operation of all experiments, and process and store all experiment data. To accomplish these functions most effectively, the deck is divided into five separate labs. These include labs for optics, geophysics, electronics, bioscience, and science information center. Further discussion of these labs is presented in Section 4.2.2. Extending from the optics lab is a small 30-inch diameter airlock used to retrieve the mapping camera for film changing and maintenance.

Located centrally and in the ceiling is a pressure hatch leading to the combination radiation shelter/emergency pressure compartment. Also located centrally but in the floor is the pressure hatch leading to the airlock used for crew transfer to the MEM, logistics vehicles, or extra- vehicular activity operations. Beneath the floor of this deck and near the aft exit are located the automatic maneuvering units used for extra- vehicular activity (EVA) operations. Propellant for these units is replenished prior to entry into the crew compartment while oxygen and other expendables are replenished after entry.

Boeing STCAEM Mars NTR 2

STCAEM Mars NTR 2
(expendable version)
ENGINES
Typesolid-core NTR
Number of Enginesx2
T/W10
Engine Mass6,800 kg
Engine Thrust334,000 N
Total Thrust668,000 N
Initial Acceleration0.8 m/s
Specific Impulse925 sec
Exhaust Velocity9,070 m/sec
MASS SCHEDULE
(tonnes)
Crew Systems Payload
Habitat and internal subsystems34.5
Exterior power, TCS, communications4.4
Airlocks (2, 1 hyperbaric)6.0
Crew and consumables14.3
MMUs(2 + consumables)0.8
TOTAL60.5
Crew Return Vehicle (CRV)
TOTAL5.8
HMEV (cryo descent, 6 crew, 9 t cargo)
TOTAL72.5
Structures
TOTAL5.5
Attitude Control System
Accumulator tanks14.3
Plumbing0.3
TOTAL14.6
Hydrogen Plumbing
Main lines2.0
Crossfeeds, valves, fittings2.1
Pressurization lines0.5
TOTAL4.6
NTP Engines
TOTAL6.8
Shadow Shield
TOTAL6.8
Tankage and Propellant
Aft tank dry15.0
Drop tanks dry (3)77.1
Trans Mars Injection propellant325.9
Mars Orbit Capture propellant148.4
Trans Earth Injection propellant73.2
Earth Orbit Capture propellant0
TOTAL639.6
Initial Mass in LEO
TOTAL817
PERFORMANCE
Mass Ratio4.6
ΔV13,850 m/s

This is from Space transfer concepts and analysis for exploration missions (1991). It is the Phase 2 design, an earlier version of the Phase 3 design.

The values in the table are for the expendable version of the mission. That is why the Earth Orbit Capture propellant is zero, the spacecraft goes zooming by Terra into a random solar orbit while the crew abandons ship and bails out in a reentry vehicle. The RV aerobrakes to Terra's surface.

The reusable version has a initial mass in LEO of 1,028 metric tons, the extra 211 metric tons of propellant allow the ship to decellerate into LEO.

The values in the table are for a crewed mission, with the crew and ascent vehicle HMEV cargo package. The uncrewed mission would have a habitat module or mixed-cargo HMEV package, only one NTR engine, no crew systems payload, and perhaps no Trans Earth Injection propellant.

The spacecraft main sections are boosted by 150-tonne-capable heavy lift vehicles. Two launches puts the spacecraft into LEO where it is assembled. The propellant is boosted in expendable hydrogen tanks and attached.

On-board power is supplied by a solar cell array rated at 27 kWe average power, with batteries to supply power during propulsive maneuvers (the thrust of which would snap the array) or during solar occultations.


HIGH LIFT-TO-DRAG MARS EXCURSION VEHICLE (HMEV)

Three catagories of payload to be landed were planned:

  • Crew and ascent vehicle (MAV)
  • Bulky and heavy cargo, e.g., the surface habitat module
  • Mixed cargo: e.g., collections of rovers, science equipment, power systems, and supplies

The Crew and ascent vehicle mission can carry an additional 5.6 metric tons of cargo. The heavy and mixed cargo unmanned mission can carry 38 metric tons of cargo.

Boeing STCAEM Mars NTR 3

Boeing STCAEM
(Nuclear Piloted version)
EngineSolid-core
NTR
Thrust330,000 N
Engine Mass3,402 kg
T/W≥10:1
Specific
Impulse
925 sec
Exhaust
Velocity
9,070 m/s
PropellantLiquid
Hydrogen
Crewx6
Habitat Module47,000 kg
Payload
(Mars Lander)
5,700 kg
Dry Mass260,360 kg
Propellant Mass554,520
Wet Mass814,880 kg
Mass Ratio3.1
ΔV10,350 m/s

This is from Space transfer concepts and analyses for exploration missions (STCAEM), phase 3 (1993).

The report focuses on using a NTR rocket to bootstrap a lunar camp, but the latter part examines a Mars landing mission. It examines three mission options, I'm only going to give the details about the largest. The different missions hinge upon the capabilities of the Terra-To-Orbit heavy lift vehicles assumed to be available.

Boeing STCAEM Dress Rehearsal

STCAEM Dress Rehearsal
Characteristics
Crewx6
Mission Duration175 days
PayloadLEV w/ 30,000 kg cargo
Enginesx2 NTP
Thrust@334,000 N
(668,000 N total)
Isp925 sec
Exhaust velocity9,100 m/s
T/W10:1
Engine operational life10 hrs
Total mission burn time1.2 hrs
Number of burnsx7
Abort modeCrew Return Vehicle (CRV) w/chemical rocket
Length90 m
Width14 m
Mass Schedule
Habitat mod47,000 kg
Payload
(lander)
76,700 kg
Tanks32,600 kg
NTR engines7,500 kg
Shadow Shield6,800 kg
RCS3,700 kg
Structure5,200 kg
CRV8,700 kg
Dry Mass188,100 kg
Propellant202,100 kg
Wet Mass390,200 kg
Mass Ratio2.07
ΔV6,640 m/s
Lunar Excursion Vehicle (LEV)
EngineChemical
(LH2-LO2)
Number enginesx4
Engine mass@250 kg
Thrust134,000 N
Isp475 sec
T/W (w/x2 engines)1.6
Wet Mass76,700 kg
LEV cargo to surface30,000 kg
Crewx6
Hab Module4,250 kg
Propellant
(descent)
26,477 kg
Propellant
(ascent)
6,535 kg

This is from documents in the Boeing Space Transfer Concepts and Analyses for Exploration Missions (STCAEM) family: 19920013259, 19920014805, 19930021845, and 19940011358. Note that these documents are somewhat separated in time, and there are major differences between the various ship diagrams.

Yes, it does resemble the Discovery XD-1 from 2001 A Space Odyssey, but that is because form follows function. Both have a spherical habitat module because a sphere has the smallest surface area with the largest enclosed volume. And both are long and skinny because of the radioactive engines. Distance is great radiation shielding because it has no mass cost. Every gram counts, y'know.

What is different from the Discover is that the spine is off-center. The dress rehearsal uses a saddle truss.


Boeing figured that it would be prudent to perform a dress rehearsal for a Mars mission, but traveling to Luna instead. So if anything went wrong the astronauts could be rescued. Instead of being stranded 8.6 months and 1.66 astronomical units away from Terra. The spacecraft has lots of newly developed moving parts that needs to get the bugs worked out. But you don't want the shakedown cruise turning lethal. Sadly they cannot test the aerobraking system since Luna has no appreciable atmosphere.

Not that they were going to throw away the spacecraft after the dress rehearsal, that thing is expensive. The lunar mission would only use 1.3 hours of nuclear engine burn time, about 80% of the engine operational life is still available. The CRV Terra reentry capsule (a glorified Apollo command module) would not actually be ridden down to the surface, it would be examined in orbit in a post-flight checkout. So they probably could refurbish the spacecraft to the point where it could actually be used for the real Mars mission. Most of the refurbishing would be swapping out the LEV for the Mars lander, attaching fresh propellant tanks, and restocking consumables.

One of the most important checks will be on the reactor core elements. Each thermal cycle of powering-up/powering-down will stress the elements. The elements cores or the coating might delaminate, whereupon the hot hydrogen propellant would start eroding them. This will help determine reactor life expectancy. Which had better be longer than the total mission burn time or the mission is cancelled.

Another important test is checking that there is no neutron cross-talk between the two engines. The design does not include a neutron isolation shield. If neutrons from engine Alfa increase the fission rate in engine Bravo, that would be bad. As in "reactor melt-down" bad. Probably "kill all the crew" bad, since the radiation shield is only rated to stop irradiation from normal operation.

The habitat module is an aluminum composite-reinforced metal matrix pressure vessel with unreinforced interior secondary structures. It provides full-service crew systems with private quarters, galley/wardroom, command and control, health maintenance, exercise and recreational equipment, and science and observation posts. The crew on the dress rehearsal will probably have all sort of notes on how to improve placement of various systems and the internal geometry.


REAL MARS MISSION

Crewed vehicle is sized for a 465 day transfer trajectory. Instead of the LEV carred in the dress rehearsal, the spacecraft carres a 72,000 kg Mars Excursion Vehicle (MEV) to ferry the exploration team to and from the Martian surface.

As with many such missions the return to Terra is performed by abandoning the spacecraft and landing the crew on Terra in the CRV reentry capsule. The spacecraft goes sailing off into an eccentric helocentric orbit, spitting neutrons from the radioactive engines.

Boeing STCAEM Cryo/Aerobrake

Boeing STCAEM
(Chemical Piloted version)
Height50 m
Span30 m
Fuel/PropellantLOX/LH2
Crew4
Mars Surface
Payload
25,000 kg
Dry Mass301,000 kg
Propellant500,000 kg
Wet Mass801,000 kg
Propellant
Fraction
0.62
Mass Ratio1.6
Specific Impulse475 sec
Exhaust Velocity4,660 m/s
ΔV2,190 m/s
MissionOpposition
Outbound Time350 days
Mars Stay Time30 days
Return Time200 days
Total Mission580 days

This is from Space transfer concepts and analysis for exploration missions. Volume 2: Cryo/aerobrake vehicle It is a reference mission using cryogenic chemical fuel plus aerobraking. When you go chemical your delta-V budget become real tight, which explains the use of aerobraking.

The study assumed that the spacecraft will be boosted piecemeal into orbit with eight launches of a Shuttle Z carrying 140,000 kg per launch.

From LEO the Trans-Mars Injection Stage (TMIS) will use LOX/LH2 to inject the spacecraft into Trans-Mars trajectory. The TMIS is discarded after the burn. The crew breaks out a deck of cards to while away the next 350 days until Mars Capture.

The payload part of the spacecraft featured two aerobraking shells. One shell holds the unoccupied Mars Excursion Vehicle (MEV), the other holds the Mars Transfer Vehicle (MTV) containing the crew. As the vessel approaches Mars it will use aerobraking because it cannot afford to carry enough fuel for powered braking. 50 days prior to Mars capture the MEV and MTV will separate.

The unoccupied MEV will aerobrake one day in advance under robot control. This is so if the atmospheric composition of Mars presents any rude surprises, it will be the uncrewed MEV that will burn-up in reentry / ricochet off into a doomed orbit into the big dark.

The crewed MTV will aerobrake a day later, if need be altering the course using data from the MEV disaster. Assuming the MEV survived the MTV will rendezvous.

The crew enters the MEV and does a complete check out. Afterwards the MEV leaves the MTV in parking orbit and descends to the Martian surface. The MEV jettisons its aerobraking shell prior to landing.

The crew has 30 days to perform all the science they possibly can cram in.

Upon Mars departure, the crew uses the MEV's upper stage (the Mars Ascent Vehicle or MAV) to travel into Martian orbit to rendezvous with the MTV. The MAV is jettisoned and the MTV does a Trans Earth Insertion burn. The crew opens a fresh deck of cards to deal with the tedium of the next 200 days until Terra capture.

Depending upon the mission design the crew either abandoneds the MTV and lands on Terra in a Mars Crew Return Vehicle (MCRV), or uses the MTV's aeroshell to aerocapture into LEO parking orbit for refurbishment and reuse.


Spacecraft


Aerobraking Shield


Transfer Vehicle


Terra Reentry Vehicle


Mars Lander


Spin Gravity Configuration


Aerobrake Shield Booster Vehicle

Boeing STCAEM Mars NEP

STCAEM Mars NEP
COMMON
Payload
Descent Aerobrake7,000 kg
MEV Descent Stage18,700 kg
MEV Ascent Stage22,500 kg
Surface Equipment25,000 kg
Transhab (4 crew)44,300 kg
TOTAL117,500 kg
Engine
EngineIon
Isp10,000 sec
Exhaust Vel98,100 m/s
Propellantargon
Reactor 17,400 kg
Reactor 27,400 kg
Shadow Shield8,600 kg
Primary Cooling20,100 kg
Auxiliary Cooling2,200 kg
Boiler21,600 kg
Turboalternators16,300 kg
Alternator
Radiator
2,600 kg
Turbopumps400 kg
Rotary Fluid
Manage
3,100 kg
Main Cycle
Radiator
10,600 kg
Main Cycle
Condenser
1,300 kg
Main Cycle
Plumbing
5,000 kg
Aux Cycle
Radiator
3,300 kg
Aux Cycle
Condenser
1,300 kg
Aux Cycle
Plumbing
6,000 kg
Power Conditioning
Radiator
1,100 kg
Plumbing Insulation4,100 kg
Engine Assembly23,500 kg
Power Management
and Distr.
68,000 kg
TOTAL211,100 kg
MICROGRAV VERSION
Struture
5 m Bay
Graphite-Epoxy
Truss
4,500 kg
Berthing Adaptor6,600 kg
TOTAL11,100 kg
Utilities
Comm600 kg
RCS5,700 kg
Avionics2,500 kg
Housekeeping
Power Distr.
500 kg
PV/RFC Power2,300 kg
Robotics3,600 kg
TOTAL15,200 kg
Propellant System
Tanks3,300 kg
Feed Lines100 kg
Propellant167,200 kg
TOTAL170,600 kg
Micrograv Totals
RAW TOTAL525,500 kg
GROWTH15%
WET MASS561,100 kg
DRY MASS368,820 kg
Mass Ratio1.52
ΔV41,160 m/s
Trip Time490 days
SPINGRAV VERSION
Struture
5 m Bay
Graphite-Epoxy
Truss
8,100 kg
Berthing Adaptor6,600 kg
TOTAL14,700 kg
Utilities
Comm600 kg
RCS5,700 kg
Avionics2,500 kg
Housekeeping
Power Distr.
500 kg
PV/RFC Power2,300 kg
40 MWe Roll Ring12,000 kg
Robotics7,200 kg
TOTAL30,800 kg
Propellant Sys
Tanks3,400 kg
Feed Lines100 kg
Propellant171,700 kg
TOTAL175,200 kg
Micrograv Totals
RAW TOTAL549,300 kg
GROWTH15%
WET MASS587,800 kg
DRY MASS390,345 kg
Mass Ratio1.51
ΔV40,430 m/s
Trip Time520 days

This is from Space transfer concepts and analysis for exploration missions. Implementation plan and element description document (draft final). Volume 5: Nuclear electric propulsion vehicle

I apologize for the illustrations, saying they are of poor quality is putting it mildly.

This is part of the family of mission concepts developed by Boeing for their Space transfer concepts and analysis for exploration missions study (STCAEM). This is the one using nuclear powered ion drives.

In the diagram below:

This report focuses on the nuclear powered ion drive option. Fantastic specific impulse but the low thrust means it takes forever to spiral out of orbit. Another problem is the severe cathode and grid erosion, limiting the thruster lifetime to about 10,000 to 20,000 hours (about 833 days). Which is about the same lifetime of a modern day Hall Effect ion drive.

The advantages include resuability, incredibly good specific impulse, no need of aerobraking for the main vehicle, great mission flexibility (insensitive to mission start dates, capture dynamics, and/or changes in payload mass), and low resuppy mass (the argon propellant is a tiny 1/3rd of total vehicle mass, unlike the 3/4 typical of chemical rockets).

Disadvantages include the massive technological advancement needed to develop a complex high-performance power system and a large liquid-metal radiator system.


The power system uses twin uranium fasts reactors. They heat a working fluid which drives turboalternators, which produce electricity. The working fluid is then cool by heat radiators and sent back to the reactors. The electrical power is conditioned for transmission and sent to the thruster system by the distribution buss. The power plant has an expected lifespan of 10 years, allowing several trips to Mars. The report says that the disposal location of the reactors are yet To Be Determined.

To make fast trips and low Initial Mass in LEO (IMLEO) the design needs a reactor power level of 20 to 40 MWe and a low low specific mass (alpha of 4 to 7 kg/kW, that is, ship kilograms per kilowatt of electricity). Which is exactly the hardest thing to do, of course. This was the focus of the entire design, obviously because this uses nuclear ELECTRIC propulsion. No electricity = no propulsion.

Naturally the mission tried to use every gravity assist possible in a desperate attempt to reduce the required delta-V. During Terra escape the ship does a Lunar swing-by to get a sweet 1,000 m/s delta-V reduction.

Since ion drives have thrust measured in hummingbird-power and accelerations measured in snails, it is going to take a long time to slow down enough for Mars capture. In this case "long enough" means "one month." So as it goes speeding by Mars, the Mars Excursion Vehicle (MEV) jettisons and aerobrakes to land on Mars. Cleverly this allows the spacecraft an amount of braking time equal to the Mars surface stay time. When it finally finishes braking it enters an eccentric orbit. This allows the MEV multiple attempts to rendezvous.

The same trick is used for Terra capture. The crew bails out in an Earth Crew Capture Vehicle (ECCV) and aerobrakes to the surface and a ticker-tape parade. The spacecraft spends the next 200 days braking into orbit, which is really going to erode the heck out of the ion drive.

Refurbishing the ship for a new trip has a little problem. Due to the regrettable location of the deadly Van Allen radiation belt the refurbishment orbit can either be in LEO (400 km) or GEO (35,000 km). LEO is preferable but the NASA nuclear safety protocols frown on radioactive 40 megawatt nuclear reactors in such a low orbit. If it fell to Terra the disaster would make Kosmos 954 look as harmless as a glow-in-the-dark wristwatch. A research study finds the risk to be minimal, because the radiation from an operating reactor is within allowable limits at 400 km and the ion drive uses a circular spiral instead of a ballistic trajectory which eliminates the risk of accicental Terra atmospheric reentry. But fat chance of getting Congress to allow this.


Two versions of the NEP were studied, without and with spin-gravity. Or the Microgravity Version (μg NEP) and Artificial Gravity Version (Ga NEP).


Microgravity Version (μg NEP)

This design has no artificial gravity, so the crew need lots of medication and exercises or muscle atrophy will render the astronauts incapable of moving even under the relatively weak Martian gravity. On the plus side this design avoids the maintenance nightmare of rotating joints and a vast reduction of the number of points of failure.

The engine assembly has 40 ion thrusters (including 10 spares) in a 5 × 8 array. Each thruster is 1×5 meters with beam neutralizers located in between the thrusters.


Spin Gravity Version (Ga NEP)

This design does indeed have artificial gravity, so the crew will not suffer muscle atrophy. On the minus side this design has an increased number of points of failure, especially that accursed rotating joint.

Basically the ship is a Tumbling Pigeon. The entire spacecraft spins like a top, except for the ion engine arrays. These are de-spun by roll rings so the engines always point in the same direction. 1g of artificial gravity is provided using a rotation rate of no more than 4 rpm to avoid crew nausea.

The roll rings are a challenge since they have to transmit megawatt levels of electricity across a spinning joint. Not to mention transferring the propellant.

Other tumbling pigeon designs do not de-spin the engines, instead the engines are mounted on the spin axis to avoid the transfer problems. This design does de-spin the engines to avoid another problem: rotational angular momentum. You see, a tumbling pigeon's angular momemtum makes the ship act like a huge gyrostabilizer, resisting all attempts to change the spin axis. The trouble is that you have to change the spin axis for thrust vectoring. So when you want to turn the engine to point in the opposite direction for deceleration, the gyrostabilization effect fights you tooth and nail. This takes lots of RCS propellant to fight this, or lots of RCS propellant to de-spin then change engine orientation then re-spin. Either way you'll need significantly more RCS propellant, and every gram counts.



Bono Mars Glider

Bono Mars Glider
PropulsionChemical
LOX/LH2
Exhaust Velocity4,400 m/s
Specific Impulse449 s
Payload to
Surface
480,000 kg
Dry Mass300,000 kg
Propellant
Mass
500,000 kg
Wet Mass800,000 kg
Propellant
Fraction
0.62
Mass Ratio2.63
ΔV4,260 m/s
Glider Length38 m
Glider Wingspan29 m
Hab Module
Height w/engine
13.7 m
Hab Module
Dia
5.5
BoosterBono HLV
Booster
Mass
3,000,000 kg
Mass with
Payload
3,800,000 kg
Booster
Engine thrust
6,700,000 N
Rim Booster
Engine Dia
7.5 m
Core Booster
Engine Dia
9.5 m
Num Booster
Engines
x7
Total Booster
Thrust
46,900,000 N
Stack Height76 m
Stack Dia25 m
Crew8
Outbound time259 days
Mars stay time490 days
Return time248 days
Total mission
time
997 days

This is from "A Conceptual Design for a Manned Mars Vehicle" by Philip Bono, in Advances in the Astronautical Sciences, Vol. 7, pp. 25-42 (1960). Actually since I have yet to locate a copy of the paper, this is mostly from David Portree's article in his always worth reading Spaceflight History blog.

In 1960 the Boeing Airplane Company was working on the X-20A Dyna-Soar orbital glider for the US Air Force. This inspired Philip Bono to envision a huge version for a Mars mission. Just like the Widmer Mars Mission, it was optimistically scheduled to depart in 1971, to take advantage of the next Hohmann launch window. Oh, isn't it just precious how idealistic we were back in the 1970's?

Apparently this was the first design for a Mars mission that was single-launch. That is as opposed to multiple launches boosting components that are assembled in orbit to create the mission vehicle. It is an arrow design.

The Dyna-Soar was only 10.77 meters long and 6.34 meters wide at the tips of its delta wings, carrying a single person. Bono's glider was a monstrous 38 meters long and 29 meters along the wing, carrying a crew of eight. The glider is split into two stages, as part of the strategy to blast off from Mars. Pretty much all designs for Mars landers are two staged; but they look like two staged rockets, not two stage gliders.


Bono's Mars mission stack had the glider perched on a habitat module (with integral Centaur engine), which was in turn perched on the short third stage. This is the core. Six full sized booster rockets would be clustered around the core (this is what Kerbal Space Program calls asparagus staging). Four of the boosters are the first stage, two are the second stage. Stack would be 76 meters tall and have a wet mass of about 3,800 metric tons.

The cluster of six full-sized booster rockets and the short booster at the center compose the Bono heavy lift vehicle (HLV), that is, stages one through three. The stack of the glider, habitat module, and Centaur engine is the spacecraft proper. It has a wet mass of 800 metric tons.

The boosters use plug nozzles instead of conventional bell nozzles to reduce engine mass and cooling requirements. This is why the boosters in the pictures have pointed ends instead of the usual bell-shaped exhaust. The boosters would have a combined thrust of about 46,900 kiloNewtons.

The habitat module is 13.7 meters tall and 5.5 meters in diameter. Internal breathing mix is 40% oxygen + 60% helium, so it's going to be Donald Duck time for the next thirty months. Module has an inflatable 16 meter radio dish to communicate with Terra. It also has a Pratt & Whitney Centaur engine with 89 kiloNewtons of thrust.

Electricity is supplied by a small nuclear reactor located in the glider's nose. Which is why the crew will be spending most of the time living in the habitat module, as far away from the reactor as they can possibly get.

Through the use of cross-plumbing, all seven modules fired at lift-off, fed from four of the outlying tanks. These four were jettisoned at propellant exhaustion at 60 km altitude (first stage). The stack would continue with just the core and two outer boosters. At 107 kilometers the two remaining outer boosters would be jettisoned (second stage). The short core booster continues to burn until the stack enters the trans-Mars trajectory, then it is jettisoned (third stage). The habitat module's antenna is now inflated.

If at any point a booster fails, the upper stage of the glider will perform an emergency detachment and do its darnest to land the crew back on Terra.

The stack is oriented with the glider nose aimed at the Sun, to protect the habitat module and its rocket engine from solar heating. The eight crew members leave the glider, crawling through a tunnel to enter the habitat module.

Transit time from Terra to Mars is 259 days. I trust they brought along a poker deck.

Upon arrival at Mars, the habitat module would eject a 9 metric ton capsule containing 256 days worth of eight astronaut's sewage. This would eventually impact Mars' surface, prompting every exobiologist on Terra to howl for Philip Bono's head (now they will never ever be sure if a newly-discovered Martian bacterium is an alien life form or an e. coli fugitive from some astronaut poop).

The eight crew members exit the habitat module and enter the glider. The glider separates from the habitat module and heads for a Mars landing. Meanwhile the habitat moduel uses the Centaur engine for Mars orbit insertion, under automatic control. Note the Centaur engine does not do any braking for the glider. This means the glider is in for a hot time as it has to aerobrake not only the orbital velocity but also the transfer velocity. But it saves on Centaur fuel. Remember: every gram counts.

The glider enters the Martian atmosphere, slows with a drag parachute, and glides to the landing site. At an altitude of 600 meters it uses three landing engines to hover and gently set down. The glider sits on landing skids with its nose pointed 15° off vertical (angled for the future blast-off).

(Unfortunately for Bono's design, it was crafted with the assumption that Martian surface air pressure was 8% of Terra. We now know that it is less than 1%. Neither the parachute nor the glider wings would function at all in such a tenuous atmosphere. Oops.)

The crew would remove the reactor from the glider's nose and relocate it about a kilometer away, so the radiation doesn't kill them. It supplies electricity to the camp via cables that are, you guessed it, about a kilometer long. A six meter living dome is inflated, and a two metric ton Mars rover is unpacked.

The crew will live on Mars for the next 479 days, doing scientific research, until the next Mars-Terra Hohmann launch window arrives. Curse those long synodic periods.

On the eve of the launch window, the nuclear reactor is re-mounted on the glider's nose. The landing rockets are pivoted to point aft, so they can serve as ascent engines. Glider is angled up 15° from vertical for lift off.

The upper stage of the glider blasts off into orbit, using the lower stage as a launch rail.

(as a side note, I use the "blast-off" image as inspiration when I designed the scoutships for an illustration of the tabletop boardgame Stellar Conquest.)

In orbit, the glider rendezvouses with the habitat module. The crew perform an EVA to manually dock the glider to the habitat module, and to jettison the empty Centaur engine fuel tank. This torus shaped tank surrounds the fuel tank for the return trip. The empty was retained until now to protect the inner full tank from meteor strikes. But now it has to go because (chorus) every gram counts.

The Centaur engine does a burn to enter a Mars-Terra Hohmann trajectory, using fuel from its internal fuel tank. Transit time is about 120 days. Time to break out a fresh deck of poker cards.

It is unclear to me from the description if the stack does a further Centaur burn to enter Terra orbit, or if it uses aerobraking. Seeing the strategy of the rest of the mission, my money is on aerobraking. In any event, after the crew enter the glider, they jettison both the habitat module and nuclear reactor (and presumably 120 days worth of sewage). These burn up in the atmosphere, with the reactor causing screams of outrage from the anti-nuclear community.

The glider lands on its skids at a NASA landing site in the desert. The crew open the doors and can now stop talking like Donald Duck. The news reporters take lots of photos as the crew is stuffed into a quarantine unit. True if there were any lethal Martian plague germs the incubation period would probably be less than 120 days, but you can never be too careful with possible Martian versions of The Andromeda Strain.


I tried making some images of the Glider, using the horribly fuzzy blueprint above as a reference. I'd love to find a better blueprint, there are quite a few spots where it is not clear how the parts come together.

Borowski Inspired Designs

These designs are either by or share most of their features from those of Stanley K. Borowski. The characteristic features are:

  • A trio of solid core nuclear thermal rockets for the propulsion section (some designs have additional propulsion)
  • A TransHab for the crew section
  • Large liquid hydrogen propellant tanks covered in what looks like gold foil
  • Often includes a saddle truss

Most of them have large photovoltaic arrays for power, especially to cryogentically cool the liquid hydrogen. If a Borowski design does not have photovoltaic arrays, it uses a bimodal nuclear thermal rocket for power.

All use solid core nuclear thermal rocket engines:

Borowski NTRs
Class NameThrust
(Newtons)
Thrust
(klbf)
Engine Mass
(kg)
Notes
PEWEE111,200253,240A little too massive for Mars mission
BIMODAL-NEP111,200253,240?
(+power plant)
When not thrusting can produce 425 kW@
(enough electricity for ion drive + LS)
SNRE73,00016.52,400Small Nuclear Rocket Engine
Just right for a Mars mission
SNRE-LANTR73,000 to
253,000
16.5 to
56.8
2,400?Gearshift capable
BIMODAL67,000152,224
(+power plant)
When not thrusting can produce 25 kW@
(enough electricity for life support)
CRITICALITY-
LIMITED
33,0007.41,770Smallest possible engine due to fission critical mass.
Not enough thrust for Mars mission.

A. C. Clark

A. C. Clark
SNRE-class Engine
Thrust73,000 N
Specific
Impulse
900 s
T/W3.06
Engine
Length
4.46
Engine
Power
367 MWt
Fuel
Length
0.89 m
Pressure
Vessel
Diameter
0.98m
Num
Fuel
Elements
564
Num
Tie-tube
Elements
241
Fissle
Loading
0.6 g U
per cm3
Max
Enrichment
93%
U-235 wt
Max
Fuel
Temp
2,860 K
U-235
Mass
59.6 kg
Spacecraft
Crew Size5
Length89.4 m
Engine Arrayx3 engines
Mass
Engine Mass100 t
Shadow
Shield Mass
6 t
In-line
Tank Mass
95.8 t
Star Truss
& x4 drop tanks
197.5 t
Payload86.7 t
Inital Mass
LEO
480 t
Propellant
Engine
Propellant
62.4 t
In-line
Propellant
71.6 t
Drop Tank
Propellant
141.4 t
(35.4 @)
Payload
Hab Modules42.2 t
5 crew + suits1.0 t
Logistics Hub7.2 t
Tunnels
and braces
5.5 t
Consumables4.4 t
Contingency
Consumables
8.1 t
Orion MPCV13.5 t
RCS and
Propellant
4.8 t

The A. C. Clark (sic, Clarke not Clark) is a spacecraft built around the Small Nuclear Rocket Engine (SNRE) instead of the old Pewee-class. It is from Conventional and Bimodal Nuclear Thermal Rocket (NTR) Artificial Gravity Mars Transfer Vehicle Concepts (2014)

They originally tried designing a spacecraft (called Copernicus) capable of a Mars mission, for the Mars Design Reference Architecture (DRA) 5.0 study. Unfortunately they determined that exposure to freefall over the mission duration would cause unacceptable damage to the astronauts. So they created a variant using "tumbling pigeon" artificial gravity called the Copernicus-B, and a stretched tumbling pigeon called Discovery. Unfortunately again both Copernicus-B and Discovery require bimodal NTR, which the designers determined was not a mature technology and thus unsuitable for the DRA.

The designers went back to the drawing board and made the A. C. Clark. This was a spacecraft using the mature technology of photovoltaic arrays for auxiliary power. Such arrays work very poorly on tumbling pigeons, so the designers used a more conventional centrifuge, Martin Marietta's Concept 6.

This had two habitat modules whose long axes were oriented perpendicular to the longitudinal spin axis ("tangential" or "Dumbbell B" configuration). The hab modules are attached to an octagonal-shaped central operation hubs via two pressurized tunnel. The hub is 6.4 meters across the flats. It has the primary docking port on the front, and 2 contingency food containers port/starboard.

The tunnels have a length of 11.5 meters, any longer and the hab modules would not be protected by the engine shadow shields. The tunnels have an outside/inside diameter of 1.5 m/1.2 m, wide enough to pass two shirt-sleeve astronauts or one suited astronaut at a time. The tunnels contain ladders, electrical cables, and the ventilation system (fans, scrubbers, and ducts).

The spacecraft has one in-line liquid hydrogen (LH2) tank, and four LH2 on a "star truss."

The sun-facing side of the hab modules and pressurized tunnels is covered with the photovoltaic power array. 30 m2 of PVA over each tunnel, 75 m2 over each hab modules, for a total of 210 m2. The PVA is rated at 8.1 m2/kWe, so the total array produces 26 kWe.


Habitat modules

The habitat modules are Space Station Freedom type. Each module is a fully independent system. They have a diameter of 4.6 meters. Each module can support a five person crew. Ordinarily they support 3, but they have been uprated to handle the entire crew in case of emergency. Each module has a docking port at one end and a dish antennae at the other. To minimize habitat mass, the access tunnels enter directly into the “top” of each habitat module via pull-down ladders.

As with most centrifuges, the command/work station displays are oriented vertically to minimize left-right head rotations, crew at work station have the lateral axis through ears parallel to spin axis, and the sleeping bunks are oriented parallel to spin axis. This helps control spin nausea. Turning one's head or toss-turn in your bunk is just asking for the Coriolis effect to make your stomach heave.

Because each habitat is straight, not curved, artificial gravity will feel weaker at the center and stronger at the ends. If you stand in the center and place a marble on the floor, it will roll "downhill" to one of the ends. Walking from an end to the center will feel like walking uphill.

The rotational radius at the hab modules is 17 m. 3 rpm will produce 0.167 g (Lunar gravity). 4.5 rpm will produce 0.38 g (Mars gravity). Maximum nausea free spin rate of 6 rpm will produce 0.68 g. A nausea inducing spin rate of 7.25 rpm will produce 1.0 g. As previously mentioned the rotational radius is constrained in order to keep the hab modules inside the shadow cast by the engine shadow shields, protecting the crew from deadly atomic radiation. The radius can be increased if the star truss is lengthened (but this increases the structural mass at the expense of the payload mass). During the transit to Mars the spin rate will be set to Mars gravity to acclimate the crew.

Each hab module will have one crew quarter room outfitted as a storm cellar. The crew will shelter within them if a solar proton storm strikes (probably 6 storms will occur during the 900 day mission). The walls of each storm cellar will have a minimum of 20 g/cm2 of shielding, though if you really want to be safe it should be 500 g/cm2. The shielding will mostly be food, life support consumables, and/or sewage.

When spacecraft is assembled in orbit, each hab module will use its attached reaction control system to fly to its connecting tunnel and dock. The side struts on the star truss are then attached to the hab modules to keep them in place under spin, and to brace the tunnels so they do not collapse backward under thrust. The RCS has lots of propellant, because it is needed to spin-up and spin down the centrifuge.


Race to Mars

Race to Mars (2007) was a two-part miniseries about a fictional Mars mission that aired on Discovery Channel Canada. The mission is reasonably closely based on NASA's DRA 5.0 Mars mission, using Borowski's Bimodal NTR spacecraft designs.

You can watch it below on YouTube.

The producers of the show closely consulted with experts in the field of astronautics, which is how I justify including it here. To my untrained eye the show seemed pretty accurate. The only mistake I noticed was that they forgot that thrusting engines create thrust gravity. Anything floating will fall to the floor. This mistake happened about three times. But the rest was pretty good.

There are a few differences from Borowski. Since this was made in Canada, the ship was outfitted with a Canadarm (that's OK, Borowski's design probably should have used one as well). Borowski's had an single LH2 drop tank inside a saddle truss, RTM had a star truss with four drop tanks (just like the Arthur C. Clark). And instead of Borowski's inflated TransHab, RTM has a metal habitat module with three floors.

As is standard, the habitat module stores the consumable supplies on the walls as radiation shielding, and the sleeping compartment doubles as a storm cellar. Though I will mention that other ship designs put the flight deck inside the storm cellar as well, so you can control the ship without the pilot dying of radiation sickness.

In the movie this comes in handy. A control circuit board is damaged, the one responsible for the nuclear engines. It is located at about the midpoint of the ship, that is, half the distance of the habitat module from the radioactive engines. Any crew working on fixing the board will get a dangerous dose of radiation. As it turns out, the doors of the storm cellar are both radiation shields and easily removed. They take them out the airlock into space, attach them to the Canadarm, and position the doors so they will shadow the repair site. Two astronauts do the repairs, because of the buddy system. While one works at the hot site, the other waits in the relatively radiation free vicinity of the hab module. When the first repair worker gets their allowed radiation dose, the two astronauts swap places.

One of the repair astronauts quips that the doors are a great beach umbrella, to keep off the sun-burn. The no-nonsense flight surgeon coldly tells him if he out-stays his 90 minute alloted time, the radiation will give him a sun-burn on his internal organs.

For artificial gravity, the ship is a tumbling pigeon like Borowski's DRA, NOT a dependent centrifuge like the Arthur C. Clark. And for ships power it uses a bimodal nuclear reactor like Browoski's DRA, NOT a photovoltaic array like the Arthur C. Clark. Spin is 4.5 rpm giving an artificial gravity of 0.7 g. SpinCalc says that makes the distance from the habitat module to the center of gravity to be about 30.9 meters. CG might not be at ship's midpoint due to the fact that nuclear engines are very heavy.

They need artificial gravity. Otherwise the 330 day trip from Terra to Mars in free fall will atrophy their muscles to the point where they cannot stand and walk under gravity. Even under the 1/3rd g of Mars surface gravity. It will take about a month of reconditioning and exercise before they can function on Mars, and they only have two months of surface stay before planetary orbits decree Mars departure.

Naturally the direction of "down" while under acceleration is in the thrust direction. In tumbling pigeon mode it is the exact opposite, i.e., the floors become the ceiling (at least in the habitat module half of the ship, the other end is too radioactive to live in). They deal with this by having two flight decks, thrust flight deck and artificial gravity flight deck. Thrust flight deck is oriented for "down" while thrusting, and artificial gravity flight deck (and all the other decks) are oriented for tumbling pigeon. At beginning of mission the ship is neither thrusting nor tumbling, so the lack of gravity means the orientation doesn't matter.

The crew transport vehicle is named the Terra Nova. Its Initial Mass In LEO is about 325 metric tons.

As per DRA 5, three uncrewed cargo vehicles are sent ahead. Only if they successfully land their cargoes on the Martian surface will Terra Nova make the journey (otherwise what's the point?). Cargo 1 delivers the Shirase cargo lander with a payload of tools, supplies, the Surface Exploration Vehicles, the geological drill, and the power reactor. Cargo 2 delivers the Atlantis Mars Surface Habitat in its lander. Cargo 3 delivers into Mars orbit the Mars Ascent/Decent Vehicle (MADV) Gagarin. The MADV function is to transports the crew to the surface, then back to the Terra Nova at the end of the 60-day surface mission segment.

The four spacecraft are assembled in Terra orbit from components boosted by heavy lift vehicles. The cargo ships make the initial journey while Terra Nova waits in Terra orbit. The trip takes a bog-standard 260 day Hohmann transfer. Cargo 1 and Cargo 2 will deliver their payloads to the exploration site in Dao Vallis in the northeast corner of Hellas Basin. The Niger Valles has more interesting geography, but that place is a death-trap for landers. Dao Vallis is more dull, but more lander-friendly. Dao's floor is mostly smooth with eroded remnants. The canyon is 650 km long and averages 2.5 km deep. The long length gives the cargo landers an additional safety factor in case they drift slightly off course. Dao Vallis does border the Hadriaca Patera inactive volcano, which will give the geologist plenty of good stuff to sample and explore. They will particularly be looking for evidence of liquid water, and maybe even Martian life. Inactive volcanoes are also prime spots for lava tubes, which can be used for future Mars colonies.

Shortly before Terra Nova departs, the crew of six is boosted from Terra's surface into orbit using the Earth Return Capsule (ERC) Columbia. This will rendezvous with Terra Nova, delivering the crew.

Terra Nova uses an opposition class trajectory instead of the standard conjunction-class Hohmann. The advantage of opposition class is that the portion of the mission spent exposed to dangerous cosmic rays and solar proton storms is reduced. The disadvantage is that the duration of the surface stay is reduced as well, from 460 days to 60. The explorers will have to do their science about eight times faster. Mars opposition class missions typically include a sling-shot past Venus to reduce the mission delta-V and amount of propellant that has to be carried. The burn into Venus intercept trajectory takes about 25 km/sec of delta-V.

MISSION
Conjunction ClassRace to Mars
Opposition Class
Terra to Mars260 days330 days
Mars Surface Stay460 days56 days
Mars to Terra260 days195 days
Total Mission980 days581 days
Space Duration520 days525 days

Actually, the Race to Mars mission plan has a longer space exposure time than the conjunction class. Typical opposition class missions to Mars have Terra-Mars 290 days, surface stay 30, and Mars-Terra of 220 days. That would have reduced the space exposure by 10 days. I'm sure the experts had their reasons.

After 330 days and a sling-shot past Venus Terra Nova arrives at Mars. The Shirase and Atlantis will be on the surface while Gagarin is in Mars orbit. Also in orbit will be three dormant cargo spacecraft, each with three hideously radioactive nuclear engines. Some prudent planning will ensure that they are nudged into a graveyard orbit far away from anywhere the Terra Nova is scheduled to travel.

If bi-conic aeroshell heat-shield technology is quite advanced, there will be only one dormant spacecraft in orbit. Cargo 1 and Cargo 2 will go streaking past Mars as they jettison their payloads. Shirase and Atlantis landers will hit Mars at a much higher velocity that way, but at least it will reduce the number of nuclear reactors in Mars orbit by six. Cargo 1 and Cargo 2 will vanish into eccentric heliocentric orbits.

Terra Nova will dock with the Gagarin MADV, the crew transfers to it, then uses it to land on Mars. Note that, as with most nuclear rockets, the docking port is on the nose. Otherwise the MADV will not stay inside the anti-radiation shadow, and will backscatter deadly radiation from the Terra Nova's engines all over the crew and habitat module.

Terra Nova waits in Mars orbit unoccupied for the duration of the surface stay (60 days). Meanwhile the crew on the ground crams as much science as they can into two months.

The initial descent to the Martian surface will take about six minutes and will be rather exciting. "Exciting" being defined as "OH GOD! OH GOD! WE'RE ALL GONNA DIE!" The Gagarin is traveling at about 5,300 m/s, it has six minutes to bring the speed to zero or they will land a mite hard. The trouble with aerobraking on Mars is that the planet has very little air to brake with. The Martian atmosphere becomes thick enough to put up some resistance at a mere altitude of 125 kilometers. For the next thirty seconds the Gagarin's heat shield (on the bi-conic aeroshell) will have to cope with temperatures approaching 1,600° C.

At an altitude of 31 km the g-force will reach its maximum of 1.3 g, and maximum mach number of Mach 12.66 (using the Martian Mach scale, Mach 1 = 240 m/s). When the velocity drops to Mach 3 (in about five minutes) the parachutes are deployed. 20 seconds later at an altitude of 5 km and velocity of Mach 2, the parachutes and bi-conic aeroshell with heat shield are jettisoned and the Gagarin's retro rockets fire up. The retros pour on the thrust in a frantic effort to prevent the Gagarin from augering in.

20 seconds after thrust start the speed has dropped to 100 m/sec, slow enough for the pilot to take over manually. The altitude will be about one kilometer. An additional 20 seconds and the vertical speed will drop to zero, but it is still traveling horizontally. If everything has gone according to plan, the Gagarin will be between two and four kilometers away from the Atlantis surface habitat. 60 seconds of manually piloted horizontal flight should bring Gagarin to one km above the surface habitat and another 20 seconds of vertical descent will bring it down to a gentle landing.

Ideally landing at a point that is NOT [a] farther than walking distance to the Atlantis surface habitat and NOT [b] so close to the Atlantis that the Gagarin's retro rocket thrust sprays Martian regolith all over the surface habitat, destroying it.

After the crew does their Neal Armstrong moment and become the first humans to set foot on Mars on live television, it is time to do their job. The very first thing is to go 30 meters upwind of the rocket exhaust and get contingency samples of the Martian surface. So if some disaster makes it imperative that they have to lift off in the next 15 seconds, at least they won't go home empty handed.

The Atlantis Mars Surface Habitat is an octagonal cylindrical structure made of aluminum resting horizontally one meter off the ground. The main airlock has an electrostatic unit to remove Martian dust from returning astronauts, to prevent contaminating the habitat. The dust is abrasive and very bad for your lungs.

On the surface the nuclear reactor will provide about 50 kilowatts. The horribly radioactive little darling will be placed a long way away from the surface habitat, with long extension cords. Among other things this can be used to recharge the batteries of the two Surface Exploration Vehicles. Each can carry three astronauts. Regrettably they have no life support so the explorers will have to wear pressure suits.

After the duration of the surface stay segment (60 days) the astronauts will use the upper stage of the Gagarin to return to the orbiting Terra Nova. In the movie, the crew is annoyed to discover that the interior of the habitat module has become over-grown with penicillium chrysogenum mold. All the equipment and control panels have to be cleaned and disinfected, while the crew constantly coughs.

The crew departs for Terra in the Terra Nova, abandoning Gagarin in Mars orbit. 260 days later they approach Terra. The crew will enter the Columbia ERC and abandon Terra Nova. The crew returns to Terra's surface using aerobraking while Terra Nova sails off into an eccentric heliocentric orbit. Yes, Terra Nova still has three very radioactive nuclear engines containing valuable un-burnt nuclear fuel. But enough propellant to brake the Terra Nova into a parking orbit is precluded by the The Tyranny of the Rocket Equation. Maybe some future grubby asteroid miner will salvage Terra Nova.

Total mission duration is 581 days. All of this is pretty much standard.

Asteroid Survey Vehicle

Pewee-class Engine
Exhaust Velocity8,890 m/s
Specific Impulse906 s
Thrust111,200 N
(25 klbf)
Thrust Power512 MWt
Mass Flow12.5 kg/s
Engine Mass3,240 kg
T/W3.5
FuelUranium 235
Max Fuel Temp2940 K
Fuel Element
Length
1.32 m
U-235 Mass36.8 kg
ReactorSolid Core
RemassLH2
Specific Power6.3 kg/MW
Longest Single
Burn
44.5 min
Total Burn
Duration
79.2 min
Num Burns4

This is from Near Earth Asteroid Human Mission Possibilities Using Nuclear Thermal Rocket Propulsion (2012) by the indefatigable Stanley Borowski et al. It uses the small but potent Pewee solid-core nuclear thermal rocket. A cluster of three of these babies had more than enough thrust for a standard Mars mission. In fact, some later designes used three weaker SNRE engines to save mass and money.

This design was for an Asteroid Survey Vehicle (ASV) to explore a Near Earth Asteroid (NEA). The idea is to get some practical experience with technologies needed for a full-blown Mars mission but with a less ambitious mission. Baby-steps first. Technologies like reliable life-support systems, long-duration habitat modules, keeping blasted cryogenic hydrogen propellant from boiling away, and of course nuclear-powered rocket engines. None of these were needed for the Apollo lunar missions.

They started with the Copernicus, a three-Pewee ship designed for NASA's DRA 5.0 and described in “7-Launch” NTR Space Transportation System for NASA’s Mars Design Reference Architecture (DRA) 5.0. They created a family of options optimizing Copernicus for the Asteroid mission, each with slightly different tweeks.

Near Earth Asteroids (NEA) have a perihelion typically less than 1.3 astronomical units or 0.3 AU farther than Terra. Of course their minimum distance can be zero, if one of them crosses Terra's orbit at the wrong time. Mars never gets closer than 0.5 AU, a Hohmann trajectory is of course much longer. But the point is there are some missions to NEAs that are not much farther than the Terra-Luna distance, and much less than the Terra-Mars distance. Baby steps.

The report looks at missions to asteroids 2000 SG344, 1991 JW, and 99942 Apophis. The latter got its disturbing name when astronomers determined that the blasted thing is going to get closer to Terra than geosynchronous orbit on Friday, April 13, 2029.

Asteroid 2000 SG344 was chosen as a relatively small NEA with low delta-V mission requirements. Asteroid 99942 Apophis was chosen as a relatively large NEA with high delta-V mission requirements.


The crewed payload element includes TransHab module with four photovoltaic array power system, the short saddle truss, Multi-Mission Space Exploration Vehicle (MMSEV, basically a large space pod), transfer tunnel with secondary docking module, and the Orion Multi-Purpose Crew Vehicle (MPCV).

PAYLOAD MASS BUDGET
(metric tons)
Transhab Habitat Module
(less consumables)
22.7 (4 crew)
27.5 (6 crew)
Short Saddle Truss2.89 to 5.08
Transfer tunnel
w/2nd docking module
1.76
Crew0.4 (4 crew)
0.6 (6 crew)
Consumables
(1 year)
3.58 (4 crew)
5.37 (6 crew)
MMSEV6.7
MPCV10.0
Returned NEA samples0.1
TOTAL48.13 (4 crew)
57.11 (6 crew)

The report examined two types of missions: reusable and expendable.

In the former all the ship components and payload return to a 24-hour elliptical parking orbit (500 km × 71,136 km) around Terra for refurbishment and reuse on another mission.

In the latter the only thing that returns is the Orion reentry vehicle carrying the crew and asteroid samples, all the rest is abandoned in deep space. MMSEV and transfer tunnel are abandoned at the asteroid. Crew splashes down in Orion capsule. Abandoned spacecraft flies off into remote eccentric Solar orbit still carrying a trio of nuclear engines. This is called "disposal into heliocentric space", but in the far future there may be a mission to intercept and salvage the blasted thing and/or move it into a more permanent graveyard orbit. Those are live atomic engines after all.

The motive for expendable missions is to drastically reduced the required Initial Mass in LEO (IMLEO), reducing the hideously expensive surface to LEO boost costs.


The first three ASV options were designed for missions to the relatively small NEA 2000 SG344. Missions to that asteroid have a delta-V cost at the low end of the scale.

ASV OPTION 1

Note that Option 1 actually uses the smaller 15 klbf SNRE engines instead of the larger 25 klbf Pewee engines used by all the other options. They can get away with this by using a seven to 28 day stay at the asteroid instead of a longer stay. This reduces the delta V cost and the required propellant. On the minus side it forces the design to use a four person crew instead of six, so the designers can use the lower mass four crew Transhab module.

IMLEO is 178.7 metric tons, of which 67 is the wet mass of the propulsion stage (39.1 propellant), 60.7 is the saddle truss and wet drop tank (44.7 propellant), and 51 is crewed payload element (short saddle truss, MMSEV, transfer tunnel with secondary dock, Transhab with four photovolatic power panels, and the MPCV).

Pictured are four larger PVP panels, suitable for a Mars mission where the solar intensity decreases to 486 W/m2. Since the Near Earth Asteroid mission is not going to get much further from Sol that Terra already is, the solar intensity will stay at about 1,367 W/m2 This means the ship can get away with using two smaller PVP panels supplying about 30 kWe.

For this round trip reusable NEA mission, there are 5 primary burns (with 4 restarts) that expend a total of 79.6 metric tons of liquid hydrogen propellant. The three engines produce 200,170 Newtons of thrust at a specific impulse of 906 seconds. This translates into a total engine burn time of 58.9 minutes. The longest burn is the first perigee burn during the TNI maneuver which takes 29.8 minutes.

ASV OPTION 2

This option uses standard Pewee engines and standard tanks being developed for the SLS, in an effort to reduce development costs by using off-the-shelf equipment. But it still is force to use the smaller crew size of four.

IMLEO is 206.4 metric tons, of which 77 is the wet mass of the propulsion stage (39.5 propellant), 77.1 is the saddle truss and wet drop tank (56.7 propellant), and 52.3 is crewed payload element.

For this round trip reusable NEA mission, there are 5 primary burns (with 4 restarts) that expend a total of 91.4 metric tons of liquid hydrogen propellant. The three engines produce 333,620 Newtons of thrust at a specific impulse of 906 seconds. This translates into a total engine burn time of 40.6 minutes. The longest burn is the first perigee burn during the TNI maneuver which takes 15.7 minutes. This is about half the time required for Option 1, due to the larger thrust.

ASV OPTION 3

This is basically Option 2 upsized so it can carry a crew of six. The increase in Transhab and consumables mass means a drastic increase in propellant mass.

IMLEO is 222 metric tons, of which 81.4 is the wet mass of the propulsion stage (43.2 propellant), 81.4 is the saddle truss and wet drop tank (60.5 propellant), and 59.1 is crewed payload element.

For this round trip reusable NEA mission, there are 5 primary burns (with 4 restarts) that expend a total of 98.5 metric tons of liquid hydrogen propellant. The three engines produce 333,620 Newtons of thrust at a specific impulse of 906 seconds. This translates into a total engine burn time of 43.7 minutes. The longest burn is the first perigee burn during the TNI maneuver which takes 17 minutes.


The last three ASV options were designed for missions to the relatively large NEA 99942 Apophis. Missions to that asteroid have a delta-V cost at the high end of the scale.

ASV OPTION 4

The report calls this "Search Lite", and seems to think it has lots of advantages. Even if it is an expendable mission. Spacecraft is sized for a 344 day stay at Apophis with a crew of four.

Because of the larger delta V requirements compared to the 2000 SG344 mission, the drop tank is emptied and jettisoned during the first perigee burn. The propulsion stage tank holds the fuel for the other burns. It uses the smaller 8.5 meter diameter style of tank.

IMLEO is 221.3 metric tons, of which 94.1 is the wet mass of the propulsion stage (50.7 propellant), 74.9 is the saddle truss and wet drop tank (50.7 propellant), and 52.3 is crewed payload element.

For this expendable Apophis mission, there are 4 primary burns (with 3 restarts) that expend a total of 95.2 metric tons of liquid hydrogen propellant. The three engines produce 333,620 Newtons of thrust at a specific impulse of 906 seconds. This translates into a total engine burn time of 42.2 minutes. The longest burn is the first perigee burn during the TNI maneuver which takes 24.2 minutes, but only provides 66% of the delta V required for TNI.

Due to the lower delta V requirements for the 2000 SG344 mission, Option 4 can also go to 2000 SG344 with a reusable mission.

For a reusable 2000 SG344 mission, IMLEO is 217.6 metric tons, of which 92.3 is the wet mass of the propulsion stage (48.9 propellant), 72.7 is the saddle truss and wet drop tank (48.9 propellant), and 52.6 is crewed payload element.

For a reusable 2000 SG344 mission, there are 5 primary burns (with 4 restarts) that expend a total of 93 metric tons of liquid hydrogen propellant. The three engines produce 333,620 Newtons of thrust at a specific impulse of 906 seconds. This translates into a total engine burn time of 41.3 minutes. The longest burn is the first perigee burn during the TNI maneuver which takes 16 minutes.

ASV OPTION 5

This is a design to make a reusable Apophis mission. Which of course requires a huge increase in the amount of propellant. A third "in-line" tank is inserted between the two existing tanks. It still can only carry a crew of four.

IMLEO is 339.8 metric tons, of which 99.8 is the wet mass of the propulsion stage (57.4 propellant), 91.5 is the in-line tanks (64.8 propellant), 93.4 is the saddle truss and wet drop tank (64.8 propellant), and 55.1 is crewed payload element.

For a reusable Apophis mission, there are 5 primary burns (with 4 restarts) that expend a total of 176.1 metric tons of liquid hydrogen propellant. The three engines produce 333,620 Newtons of thrust at a specific impulse of 906 seconds. This translates into a total engine burn time of 78.2 minutes. The longest burn is the first perigee burn during the TNI maneuver which takes 38.6 minutes, but only provides 66% of the delta V required for TNI.

ASV OPTION 6

This is the second design to make a reusable Apophis mission. It avoids using a third in-line tank by outfitting the propulsion stage and drop section with tanks that are 10 meters in diameter instead of 8.5. Basically this is the full Copernicus spacecraft outfitted as an asteroid survey vehicle. It has enough extra propellant to host a crew of six.

IMLEO is 323.2 metric tons, of which 138.1 is the wet mass of the propulsion stage (87.2 propellant), 122.9 is the saddle truss and wet drop tank (93.9 propellant), and 62.2 is crewed payload element.

For a reusable Apophis mission, there are 5 primary burns (with 4 restarts) that expend a total of 171.7 metric tons of liquid hydrogen propellant. The three engines produce 333,620 Newtons of thrust at a specific impulse of 906 seconds. This translates into a total engine burn time of 76.2 minutes. The longest burn is the first perigee burn during the TNI maneuver which takes 38 minutes, but only provides 66% of the delta V required for TNI.

It can also perform a reusable mission to asteroid 1991 JW, since that only requires 7.188 km/s of delta V instead of the 7.378 km/s required for the reusable Apophis mission.


Bimodal NTR

Bimodal NTR
Engine
PropulsionSolid core NTR
Ternary Carbide
(15 klbf)
Number of engines3
Fuel Volume11.5 L
Core Power
Density
5 MW/Liters
Number Reactor
Elements
36
Number Safety
Rods
13
Reactor Vessel
Diameter
0.65 m
Reactor Fueled
Length
0.55 m
Engine Mass2,224 kg
Total Engine
Length
4.3 m
Nozzle Exit
Diameter
1.0
Engine (Thrust Mode)
Thrust per engine67,000 N
Total Thrust200,000 N
T/Wengine3.06
Exhaust Velocity9,370 m/s
Specific Impulse955 s
Propellant
Mass Flow
7.24 kg/s
Full Power
Engine Lifetime
4.5 hours
Reactor Power335 MWthermal
Engine (Power Mode)
Reactor Power110 kWthermal
Brayton Power
per reactor
25 kWelectricity
Total
Brayton Power
(2 reactors)
50 kWelectricit

This is from a NASA study Vehicle and Mission Design Options for the Human Exploration of Mars/Phobos Using "Bimodal" NTR and LANTR Propulsion (2002). The idea was to take NASA's Mars Design Reference Mission (DRM) and update it. Specifically a throwaway stage with a nuclear thermal rocket (NTR) was to be replaced with a reusable stage using an NTR with the bimodal option.

Three 200 kilonewton NTR can easily generate enough delta V to put the spacecraft through the Mars DRM. It's just that it consumes a measly 10 grams of Uranium-235 out of the 33,000 grams of 235U in each engine. It would be insane to throw away the remaining 32,990 grams of expensive 235U (per engine) as the rocket stages when leaving LEO, as per the DRM.

That's where the bimodal part comes it. Instead of using the rocket for about an hour total then either throwing it away or letting it sit idle for the rest of the 4.2 year long mission, put that sluggard to work! You throttle each engine from 335 megawatts down to 110 kilowatts and use it to run a Brayton electricity generator (about 25 kilowatts of electricity per reactor). A maximum of two reactors can be run simultaneously for generating electricity. The electricity will come in real handy to keep the fifty-odd tons of liquid hydrogen refrigerated instead of rupturing the propellant tanks. This will also remove the need for heavy fuel cells for power. And it will make the stage reusable.

Common Core Bimodal Stage
Structure2.5 mTon
Propellant Tank5.98 mTon
Propellant Tank7.4m I.D. × 19.0m
LH2 Refrigeration
System (@~75 Wt)
0.30 mTon
Thermal/
Micrometeor
protection
1.29 mTon
Avionics and Power1.47 mTon
Reaction Control
System (RCS)
0.45 to 0.48 mTon
NTR engines (x3)6.67 mTon
Shadow Shields (x3)0 or 2.82 mTon
Brayton Power
System (@ 50 kWe)
1.35 mTon
Propellant feed,
TVC, etc.
0.47 mTon
Contingency (15%)3.07 to 3.50 mTon
Total Dry mass23.55 to 26.83 mTon
LH2 Propellant51.0 mTon
RCS Propellant
max
1.62 to 2.19 mTon
Total Wet mass76.2 to 80.0 mTon

For this study they designed a common core stage, and made a family of designs by putting different payload modules on top of the core. The core has three bimodal NTR with power generation (50 kW total) and heat radiators, a propellant tank with a capacity of 50 or so tons of liquid hydrogen, and a propellant refrigeration system.

For manned missions each of the three NTR is fitted with an anti-radiation shadow shield to protect the crew. If there this is an unmanned mission the shadow shields are left off, which reduces the stage's dry mass by 3.2 metric tons. The unmanned cargo is relatively immune to radiation.

The integral liquid hydrogen tank is cylindrical with √2/2 ellipsoidal domes. It has a 7.4 meter internal diameter and a length of 19 meters. It has a maximum propellant capacity of 51 metric tons with a 3% ullage factor.

The forwards cylindrical adaptor contains avionics, storable RCS, docking systems, and a turbo-Brayton refrigeration system to prevent the liquid hydrogen propellant from boiling off over the 4.2 year mission. The highest level of solar heat for the Mars mission is when the spacecraft is in LEO, about 75 watts of solar heat penetrates the 5 centimeter Multi-layer insulation (MLI) blanketing the propellant tank (the stuff that looks like gold foil). The refrigeration system requires about 15 kWe to deal with the 75 watts of heat.

At the aft end, the conical extension of the thrust structure supports the heat radiator, about 71 square meters of radiator. Inside the cone is the closed Brayton cycle (CBC) power conversion system. It has three 25 kWe Brayton rotating units, one for each bimodal reactors. Only a maximum of two of the three units can be operated simultaneously. The CBC's specific mass is ~27 kg/kWe.

The payload is held on a "saddle truss" spine that is open on one side. This allows supplemental propellant tanks and contingency crew consumables to be carried and easily jettisoned when empty. The saddle truss would also be handy for a cargo carrying spacecraft who wants the ability to load and unload cargo in a hurry.

Bimodal Hybrid NTR NEP

NTR Engine
PropulsionSolid core NTR
FuelESCORT/
TRITON/
UO2-W cermet
PropellantLH2
Isp906 s
Thrust
per engine
111,000 N
(25 klbf)
Number
of engines
3
Total Thrust333,000 N
Expendable
ΔV
3,815 m/s
Reusable
ΔV
4,378 m/s
Electric Propulsion
PropulsionIon Drive
Power req.16 kWe
PropellantXenon
Isp3,000 s
Number
of engines
30
Total
Power req.
800 kWe
ΔV4,483 m/s

This is from A Crewed Mission to Apophis Using a Hybrid Bimodal Nuclear Thermal Electric Propulsion (BNTEP) System (2014). The same authors had an earlier version of this design.

A conventional Bimodal NTR (above) is a nuclear thermal rocket with the bimodal power option.

The spacecraft described here uses the bimodal Hybrid option (BNTR/EP), where the power output is also hooked up to an Ion Drive. This is a crude form of Shifting Gears: trading thrust for specific impulse/exhaust velocity.


Why bother with this contraption? Well, the short answer is that the BNTEP has 14.7 metric tons less wet mass than the equivalent conventional NTR. And every gram counts. Especially if you are boosting this thing from Terra's surface into LEO.

In addition, the conventional spacecraft has to be expendable. It does not have enough delta V to brake into LEO upon return, instead the crew abandons ship in a reentry vehicle while the expensive ship goes sailing off into the wild black yonder. This is because of a maximum of 110 metric tons on all spacecraft components due to booster rocket limitations.

But the hybrid BNTEP design can have the propellant tank expanded to the point where it is capable of braking into LEO and being reused, yet still keep all the components within the 100 metric ton limit.


Granted, the BNTEP has a higher dry mass because it needs more equipment (two separate propulsion systems for one). But since the ion drive has over six times the specific impulse of chemical thrusters, you need tons less propellant mass (the "wet" in "wet mass"). Both spacecraft need NTR drives for the main mission phases because you need high thrust. But for low thrust mission events, such as midcourse corrections (MCC), tank settling burns, some reaction control system (RCS) burns; the conventional NTR uses wasteful propellant-guzzling chemical thrusters (Advanced Material Bipropellant Rocket) while the hybrid BNEP uses the super-efficient propellant-sipping ion drive. Actually the ion drive can handle a small portion at the end of the departure burn as well.

A bimodal NTR requires extra power generating equipment (Brayton system) that adds dry mass (but it is insane to try and feed an 800 kWe ion drive by using a photovolatic array {PVA}). But on the other hand, this means the spacecraft does not need a photovolatic array for spacecraft life-support and cryogenic cooling power. But on the gripping hand a Brayton system has a mass of 2.87 metric tons as opposed to 0.57 metric ton for a minimal photovolatic array. Advantage goes to the conventional spacecraft.

Life-support and cryogenic cooling require 50 kWe. The ion drive array requires 800 kWe. So the conventional spacecraft has a power requirement of 50 kWe while the hybrid requires 850 kWe.

The conventional spacecraft uses a 0.57 metric ton photovolatic array that will produce 50 kWe at Apopis (practically the same distance from Sol as Terra). The hybrid spacecraft will have three Brayton units (one per engine, total 2.87 metric tons) rated for 425 kWe each but running at 2/3 maximum power (283 kWe each, total of 850 kWe). This means if one of the Brayton units malfunctions, the remaining two can be cranked up to maximum power and still supply the necessary 850 kWe.

Bimodal Hybrid NTR NEP 2

NTR Engine
PropulsionSolid core NTR
Fuel TypeUO2-W cermet
Fuel Mass200 kg
Isp906 s
Thrust
per engine
111,000 N
(25 klbf)
T/W4.5
Number
of engines
3
Total Thrust333,000 N
Total Thrust
Time
1.75 hrs
Max Thrust
Time
2.0 hrs
PropellantLH2
In-Line
Propellant
96,400 kg
Drop Tank
Propellant
39,200 kg@
Num
Drop Tank
4
Drop Tank
Propel Total
156,800 kg
Total
Propellant
253,200 kg
Reactor
Power
(NTR mode)
545 MWt
Max Power
Time
1.5 years
Operational
Time
(NTR mode)
40 MW-days
(3 engines)
Reactor
Power
(Ion mode)
1.76 MWt
Operational
Time
(Ion mode)
284 MW-days
(3 engines)
Operational
Time
(Total)
324 MW-days
(3 engines)
Operational
U-235 Fuel
burnt
0.389 kg/eng
(0.2% burn-up)
Generator
Type
Brayton
Generator
Output Max
500 kWe@
Total Power
Output Max
1.5 MWe
Generator
Output Norm
(2/3rd)
333.3 kWe@
Total Power
Output Norm
(2/3rd)
1.0 MWe
Brayton Heat
Radiator
970 m2
Electric Propulsion
PropulsionIon Drive
(Hall Thruster)
Isp3,000 s
Total
Power req.
1.0 MWe
Number
of engines
10
PropellantXenon
Propellant
Mass
20,400 kg
Power req.100 kWe@
(1 MW total)
Payload
Shakedown
time
50 days
Mission
Length
365 days
Total Crew
Time
415 days
Num Crew4
Crew800 kg
Habitat
Module
TransHab
Habitat
Module
22,700 kg
Consumables2.45 kg/d/crew
Consumables
Mass
4,080 kg
Multi-Purpose
Crew Vehicle
13,500 kg
This is from A One-year, Short-Stay Crewed Mars Mission Using Bimodal Nuclear Thermal Electric Propulsion (BNTEP) (2013), an earlier design from the same team that created the Bimodal Hybrid NTR NEP design for the Apopis mission.

A conventional Bimodal NTR is a nuclear thermal rocket with the bimodal power option.

The spacecraft described here uses the bimodal Hybrid option (BNTR/EP), where the power output is also hooked up to an Ion Drive. This is a crude form of Shifting Gears: trading thrust for specific impulse/exhaust velocity.

The nuclear engines have performance similar to standard PeWee class solid core nuclear thermal rockets.

The nuclear engines are used for the burns where a planet's gravity create troublesome gravity losses, while the more efficient ion drive is used for burns when there are no g-loss. Nuclear engines are less efficient but since gravity losses accrue on a second-by-second basis you want to get out of the g-loss zone fast while the meter is running. A low thrust propulsion like ion drive can take days to exit the zone.

The ion drive requires 1.0 megawatts of electricity. The 3 BNTRs can generate 1.5 MW total, but are throttled down 2/3rd so they generate 1.0 MW. The idea is that if one of the three BNTRs fail, as a fail-safe the remaing two can be throttled up to 100% and still generate teh 1.0 MW the ion drive needs.



THE MISSION

BNTR refers to the nuclear thermal engines and their burns. EP refers to Electric Propulsion (ion drive) and its burns. BNTR are used for burns where gravity-loss delta-V is a factor, and you want to use high thrust to get out of the G-Loss zone as quick as possible. Otherwise the more economical EP burns are used.

The initial Trans-Mars Injection burn (TMI) is divided into two burns: TMI-1 and TMI-2. This minimizes the gravity-loss of the TMI for reasons that I do not understand, and which the report is a little vague on. The spacecraft has four drop tanks and one in-line tank of liquid hydrogen propellant for the BNTR. Two drop tanks are jettisoned at the end of each TMI burn. The first two drop tanks have enough propellant for TMI-1, the second burn TMI-2 requires the remaining two drop tanks and some propellant from the in-line tank. Each burn is about 31 minutes long (0.52 hours).

After the TMI burns, the BNTRs throttle down from 545 megawatt NTR thrust mode to 1.76 megawatt electricity generation mode so it can feed the ion drive system. The ship coasts for 12 hours to let the BNTR engines cool off.

EP-1 burn uses the ion drive and lasts for 36.7 days. The ship then coasts for 44 days.

EP-2 burn lasts for 43.8 days and ends 12 hour prior to Mars Orbit Insertion (MOI).

The BNTRs then throttle up to 545 megawatts as they leave electricity generation mode and enter NTR thrust mode. The MOI requires 21.6 minutes of thrust (0.36 hours). The ship settles down into a 300 kilometer x 24 hour Mars orbit.

The mission has a disappointing objective of entering Mars orbit and cooling its heels there for a month. The mission is not equipped for landing crew on Mars. The crew has to stare longingly at the Martian surface through telescopes, so near yet so far. Frankly I do not see what a crew can do that an unmanned Mars orbiter cannot.

After 30 days, the spacecraft does a Trans-Earth Injection burn (TEI) of 21 minutes. The BNTRs then throttle down into generator mode. The ship coasts for 12 hours to let the BNTR engines cool off.

The EP-3 burn uses the ion drive and lasts for 80.8 days. The ship then coasts for 127.7 days.

The EP-4 burn is used if the velocity relative to Terra is greater than 11.5 km/s. That is the maximum velocity the reentry vehicle is rated for.

The spacecraft is not reusable. It does not brake into Terra orbit upon return. Instead, like other crude missions, the ship goes streaking by Terra (at 11.5 km/s) while the crew bails out in a reentry vehicle. The ship then vanishes into an eccentric Solar orbit, with most of its expensive U-235 fuel un-burnt.

Modular Mars Missions

This is from Nuclear Thermal Rocket/Vehicle Design Options for Future NASA Missions to the Moon and Mars (1995)

The main feature of the report is the NTR FIRST LUNAR OUTPOST. But the latter part of the report talks about how the spacecraft can be adapted to a Mars mission. The spacecraft are designed around Pewee-class solid core nuclear thermal rockets with a thrust of 111,200 N (25 klbf) and a specific impulse of 940 s.

At the time there were several schools of thought about how to design a Mars Mission.

"All-up" means the entire freaking Mars expedition is composed of one giant spacecraft.

"Split" is the new coolness (sometimes called split/sprint). The mission is split into an uncrewed cargo vehicle and a piloted vehicle holding all the human Mars explorers. The uncrewed cargo ship travels a slow leisurely Hohmann trajectory, since it contains no crew who would suffer from the long bombardment of space radiation and other unhealthy aspects of prolonged spaceflight. After the cargo ship reaches its destination around Mars and sends confirmation back to Terra, only then does the crewed craft depart to travel to Mars. Since this spacecraft does not have to lug along metric tons of Mars exploration equipment, it can manage a higher energy trajectory. This means the trip is shorter, reducing the time the crew suffers from space radiation et al.

The piloted mission waits for confirmation from the cargo mission because the crew needs the cargo stuff in order to return alive to Terra. If something had happened to the cargo mission, the crew would be stranded and would all die a lonely death around the Red Planet.

Glossary:

  • CTV: cargo transfer vehicle. The uncrewed cargo ship
  • PTV: piloted transfer vehicle. The crewed ship
  • MEV: Mars excursion vehicle. The Mars lander which transfers the crew to the planet's surface and later returns the crew to the orbiting spacecraft.
  • TEI: Trans-Earth Injection propellant. The propellent needed to return the crew home from Mars to Terra by a fast trajectory.

Three kinds of split missions were developed:

Split: 'All-Up' mode
     The PTV carries the crew, MEV, and TEI. The only thing the CTV carries is supplies to support the surface mission and an uncrewed supply lander. The PTV and CTV do not have to rendezvous, which reduces the number of failure modes the mission has to risk.
Split: 'no MEV' mode
     the PTV carries the crew and TEI. The CTV carries the MEV and surface supplies. Obviously the PTV and the CTV have to rendezvous in space so the crew can enter the MEV. This increases the number of failure modes.
Split: 'no MEV/no TEI' mode aka Minimum Piloted Mass mode
     the PTV carries the crew, period. There are two CTVs which carry everything else. CTV #1 has the MEV and surface supplies. CTV #1 is a propellant tanker carrying the TEI in a tank with a hose. The PTV has to rendezvous and refuel itself with all the TEI propellant, with all the extra failure modes that entails.
     In a variant, CTV #2 carries the Terra return spacecraft with the TEI propellant already inside. This avoids the headaches of zero-gravity refueling. The PTV makes its rendezvous, the crew abandons it and transfers to the return spacecraft (along with their Mars surface samples). They then use the return spacecraft for the voyage home, leaving the PTV in orbit around Mars.

The researchers decided to go with the Minimum Piloted Mass mode option. In an effort to further reduce the transportation mass requirements, they looked into making the MEV use aerobraking for landing and have some sort of in-situ resource utilization in the form of an automated surface factory manufacturing MEV propellant.

The piloted mission is preceded by three separate cargo missions which depart Earth orbit in September 2007 and arrive at Mars - 344 days later. Each cargo mission is launched on a single 200-240 t HLLV. The cargo missions use NTR propulSion for TMI and a "common" Mars aerobrake/aerodescent shell for either capture into Mars orbit or direct descent to the Mars surface. (The expendable NTR TMI stages are not shown in Figure 15.)

As envisioned by ExPO, the initial cargo mission would transport both surface and Mars orbit payload elements. The surface payload consists of a "dry" Mars ascent stage/crew cab combination along with the power system , LH2 propellant "feedstock," and propellant production plant necessary to convert Martian CO2 into LOX/CH4 propellant for the piloted MEV ascent stage. This aspect of the reference Mars mission was first proposed by Zubrin in his "Mars Direct" scenario.

The payload delivered to Mars orbit consists of a "fueled" trans-Earth injection stage and a "minimum mass" Earth return habitat. The later cargo missions deliver surface payload consisting of a habitat module, scientific laboratory, pressurized rover, consumables and miscellaneous supplies and spares needed to support a long-duration Mars exploration phase.

After the operational functions of the habitat and surface facilities are verified and the ascent stage is fully fueled, the piloted vehicle leaves Earth in November 2009. It arrives at Mars 180 days later using a "fast conjunction-class" trajectory which maximizes the exploration time at Mars while reducing the total in-space transit time to under a year.

After a 540-day stay at Mars, the crew returns in the ascent portion of the MEV to a waiting Earth-return stage and habitat module to begin its preparation for a 6-month journey back to Earth. The total duration for the piloted mission is 900 days. The crew returns to Earth in the Mars ascent vehicle crew cab which is retained and used as the Earth crew return vehicle.

After separation, the TEI stage and habitat continue along their interplanetary path for disposal into heliocentric space (because the blasted engine is radioactive).


Mars Mission/Transportation System Ground Rules and Assumptions

Table 5. Mars MissionfTransportation System Ground Rules and Assumptions
MissionSpacecraftPayload
Items
CargoTEIPiloted
Payload
Outbound
3 × (60-98.9 t)--MEV (w/41.5-64.4 t P/L)
-35.0 t-Crew Habitat
-5.5 t-ECRV
--(52.1-87.2 t)MEV (w/35-50 t Habitat
Payload
Return
-35.0 t-Crew Habitat
-5.5 t-ECRV
-0.5 t-Mars Return Samples
Maneuver
Type
SpacecraftManeuver
CargoTEIPiloted
Parking
Orbits
407 km407 km407 kmTerra Departure (circular)
-250 km × 1 sol-Mars Arrival/Departure
Perigee
Burns
222-3Earth Departure
Crew
Type
Spacecraft
CargoTEIPiloted
Crew Size-66
Propulsion
NTR System
PropellantCryogenic Hydrogen
Isp900 sec (NDR)
960 (CIS)
External Shield Mass≈ 60 kg/klbf thrust
Flight Performance Reserve1% of usable propellant
Cool down (effective)3% of usable propellant
Residual1.5% of total tank capacity
RCS System
PropellantN2O4/MMH
Isp320 sec
Structure
Tankage
Material2219-T87 Al
Diameter10m
GeometryCylindrical tank with √2/2 domes
Insulation
TMI application only2" MLI + micro shield
Cargo & Piloted Vehicles
w/NTR for TMI, MOC
and disposal
3" MLI + VCS ("core" tank)
2" MLI + micro shield ("in-line" tank)
Earth Return Vehicle
w/NTR for TMI, MOC, TEI
and Disposal
4" MLI + VCS/or
2" MLI + micro shield + refrigeration
Contingency
Engine & External Shield15%
All other dry masses10%

Expendable TMI Stage


Mars Cargo Vehicle - "All Propulsive" Option


Mars Piloted Vehicle - "All Propulsive" Option


Earth Return Vehicle - "All Propulsive" Option


NCPS Mars Mission

NCPS Mars Mission
Core stage (C)
Engine Isp, sec900
Inert Mass, mt44.99
x3 25 klbf NTP Engines12.32
x3 External
Radiation Shields
6.45
Tank inert
(w/ everything else)
26.22
Usable LH2 Mass, mt41.64
RCS Usable Prop Load, mt17.05
Boil-off to ullage, mt0.20
Stage Length, m
(engines, RCS, I/F)
~22.2
Approx. Effective LH2
PMF / λ
0.48
In-line Tank (I)
Inert Mass, mt
(w/ everything)
28.59
Usable LH2 Mass, mt66.40
RCS Usable Prop Load, mt5.51
Engine Isp, sec900
Stage Length, m
(incl. RCS & I/F)
~21.2
Approx. Effective LH2
PMF / λ
0.70
Saddle Truss & Drop Tanks, 1 ½ (D)
Inert Mass, mt38.35
Saddle Trusses
(w/ everything)
7.73
Drop Tanks
(w/ everything)
30.61
Usable LH2 Masses mt103.30
RCS Usable Prop Loads, mt8.58
Boil-off, mt1.54
Engine Isp,sec 900
Stage Length, m
(incl. RCS & I/F)
~33
Approx. Effective LH2
PMF / λ
0.73
Payload
Deep Space Hab (stocked)51.85
MPCV (CM+SM, no prop)14.49
Payload RCS/Truss/Canister14.14
Pre-TMI
Crew, mt0.79
Less mass exp.
prior to TMI, mt
(-25.95)
Mass Schedule
Core stage
wet mass total, mt
(on pad)
103.68
In-line Tank
wet mass total, mt
(on pad)
100.50
Saddle Truss & Drop Tanks
wet mass total, mt
(on pad)
151.76
Payload
wet mass total, mt
(on pad)
80.48
Mars stack interim total436.43
Pre-TMI, mt-25.16
Total TMI Stack Mass, mt411.26

This is from A Nuclear Cryogenic Propulsion Stage for Near-Term Space Missions (2013).

Abbreviations in table:

NASA experimented with nuclear thermal rockets with Project Rover, which ran from 1955 through 1972. It is really hard to work with spacecraft that use the "N-word" and which may spread the "R-word", but they are far too useful to leave on the shelf. Twice the specific impulse of the best chemical engines, and thrust values which make ion drives look like hummingbirds. So in 2011 NASA iniatied the Nuclear Cryogenic Propulsion Stage (NCPS) project.

This spacecraft design uses nuclear thermal rockets for a Mars mission.


THE MISSION

2037 Trajectory Constraints / Parameters
TMI ΔV1:1934 m/s (1813-1936)
TMI ΔV2:2084 m/s (1976-2172)
MOI ΔV:934 m/s (1029-1806)
TEI ΔV:1475 m/s ( 827-1524)
Total ΔV:5,645 to 7,438 m/s
Outbound time:212 days (158-225)
Stay time:489 days (448-569)
Return time:220 days (195-238)
TMI, MOI & TEI:1% ΔV Margin/FPR/other
TMI Gravity Losses:389 m/s total, f(T/W0)
MOI & TEI g-losses:Additional 1%
Post-TMI RCS ΔVs:180 m/s (>>7 burns)
Tank Masses (C, I, D):see table

Abbreviations in Trajectory Constraints table:

  • TMI = Trans-Mars Insertion
  • MOI = Mars Orbit Insertion
  • TEI = Trans-Earth Insertion
  • Tank C = Core Stage Tank
  • Tank I = In-line Tank
  • Tank D = Saddle and Drop Tanks

THE SPACECRAFT

Design Constraints / Parameters
# Engines / Type:3 / NERVA-derived
Engine Thrust:25 klbf (Pewee-class)
Propellant:LH2
Specific Impulse, Isp:900/nominal - TBD/max sec
Tank Material:Aluminum-Lithium
Truss Material:Composite
RCS Propellants:NTO / MMH
RCS Thruster Isp:328 sec (Fregat Isp)
Passive TPS:0.75” SOFI + 60 layer MLI
Active CFM:ZBO Brayton Cryo-cooler
I/F Structure:Stage / Truss Docking Adaptor w/ Fluid Transfer

Abbreviations in Design Constraints table:

  • ZBO = Zero boil off
  • CFM = Cryogenic Fluid Management for propellant tanks
  • TBD = To be determined
  • TPS = Thermal Protection System
  • SOFI = Spray-on foam insulation
  • MLI = Multilayer insulation

NTP system consists of 3 elements:

  1. core propulsion stage
  2. in-line tank
  3. integrated saddle truss and drop tank assembly that connects the propulsion stack to the crewed payload element for the Mars 2037 mission

Each element is delivered to LEO (407 km circular orbit) fully fueled on an SLS LV (178.35.01, 10-m O.D. / 9.1-m 25.2 m cylinder section). They are sized for an SLS capability of ~100 metric tons.

The stage uses three 25.1 klbf (111.2 kN) engines (Pewee-class) with either a NERVA-derived or ceramic-metallic (CerMet) reactor core. It also includes RCS, avionics, power, long-duration cryogenic fluid management hardware (e.g., COLDEST design, zero boil-off cryo-coolers) and automated rendezvous and docking capability. Saddle trusses use composite material and the LH2 drop tank employs a passive thermal protection system. I/F structure includes fluid transfer and electrical.


BONUS SPACECRAFT

This asteroid survey mission spacecraft from the same report uses lower-powered 15 klbf (67 kN) nuclear engines instead of 25 klbf engines. This is sort of midway between a Pewee class and a SNRE class engine.

NLTV

This is from Robust Exploration and Commercial Missions to the Moon Using NTR LANTR Propulsion and Lunar-Derived Propellants (2017) doc, slides. Supplemented with the expanded document Robust Exploration and Commercial Missions to the Moon Using Nuclear Thermal Rocket Propulsion and In Situ Propellants Derived from Lunar Polar Ice Deposits (2018)

NLTV stands for Nuclear Lunar Transport Vehicle. LTS stands for Lunar Transporation System.

The basic idea is if we set up in-situ resource utilization facilities on Luna which can produce Lunar-derived propellant (LDP) — specifically Lunar Liquid Oxygen (LLO2) and Lunar Liquid Hydrogen (LLH2) — what sort of spacecraft can this support? LLO2 can be obtained from lunar regolith or volcanic glass, both LLO2 and LLH2 can be obtained from lunar polar ice. The original 2003 study didn't know about polar ice, so it figured that hydrogen would have to be shipped from Terra while oxygen could be harvested from lunar volcanic glass. The discovery of lunar polar ice means nothing has to be shipped from Terra. The amount of lunar hydrogen and oxygen is estimated to be many billions of tons.

The availability of liquid oxygen makes the obvious choice of basing it around LOX-augmented Nuclear Thermal Rocket (LANTR) propulsion. This is a solid-core nuclear thermal rocket using liquid hydrogen propellant, but with a liquid oxygen afterburner which allows the engine to shift gears. So it can trade thrust for exhaust velocity (specific impulse) and vice versa. The gear shifting is due to the afterburner, the nuclear reactor operates at the same power level regardless of what gear is used. By judicious use of gear shifting, the total mission burn time of the engine can be cut in half. This doubles the number of missions the engine can perform before the engine comes to the end of its lifespan.

The report figures that the initial industrialization of Luna will be done by non-LANTR SNRE spacecraft, which will have to carry lunar landers along with the payload. This departs from LEO, but has to return to a 24-hr elliptical Earth orbit (EEO) because it just doesn't have the delta V to return to LEO. To give it that much delta V would require the ship's wet mass would have to almost double to 347.8 metric tons!

Once industrialization starts, small amounts of lunar liquid oxygen (LLO2 or LUNOX) will become available. This will allow lunar landers to be housed in the lunar base, so the SNRE spacecraft will not have to carry them. This will allow the spacecraft to carry lots more payload. They still will have to return to EEO instead of LEO, though.

When lunar industrialization becomes fully developed, larges amounts of LUNOX will become available and an orbital propellant depot will be established in lunar orbit. At that point the spacecraft's trio of SNRE engines will be swapped out for LANTR engines, and the in-line liquid hydrogen tank swapped for a liquid oxygen tank carrying 46.5 metric tons of LO2. Once the ship arrives in LLO, it will refill the liquid oxygen tank from the orbital propellant depot. The refueling and the LANTR gear shifting will allow the ship to return to LEO and reduces the engine burn time from 50 minutes to 25.3 minutes. This doubles the lifespan of the engine.

Bottom line is that the price to transport payload to and from Luna will drop dramatically.


These are two optimized LANTR designs: Conestoga Crewed Cargo Transport and the Commuter Shuttle. They share a common nuclear thermal propulsion system (NTPS), including the LO2 tank (though the size of the LO2 tank is different between the two). The one-way transit times to and from the Moon will be cut in half to ~36 hours. This will require the delta V budget to be increased by 25% from ~8,000 km/s to ~10,000 km/s.

NLTV Generic

In the above designs, all the LH2 tanks carry 39.7 metric tons of liquid hydrogen. The payload pallets are 2.5 metric tons each. One-way transit times to and from the Moon will be about ~72 hours.

NLTV Conestoga

ItemCargo Transport
MissionLEO⇒LLO⇒LEO
Duration36-hr “1-way” transit times
Habitat Module~11.2 t
Passenger
Transport
Module
n/a
Crew4
Passengersn/a
Star Truss
w/ 5 t payload
~8.6 t
In-line
LO2 tank
~86.6 t
LH2 NTPS~70.9 t
IMLEO
(wet mass)
~177.4 t
Refueled LLO2~71.6 t
Total
Burn Time
~25.3 min

This is the first of the two optimized LANTR designs: the Conestoga

The Conestoga carries a habitat module that supports a crew of four, weighing ~10 t. Two crewmembers operate the vehicle and manage the unloading of the payload (PL). The other two represent rotating crewmembers on assignment at the lunar base or the LPO propellant depot. Connecting the habitat module to the rest of the LANTR LTV is a “star truss” that has four concave sides to accommodate four 1.25 t payload pallets.

The forward circular truss ring also has a remote manipulator system with twin arms attached to it. Using the habitat module’s rear viewing window, the crew uses these arms to unload and attach the transport’s cargo to the depot or to a co-orbiting LLV that is transferring crew and awaiting cargo delivery.

The LO2 tank is smaller and customized for this particular application resulting in a lower Initial Mass In Low Earth Orbit (IMLEO or wet mass in 407 km altitude orbit) and LLO2 refueling requirement (~35 t).

All the missions start and end in LEO, with the mid-point being either Lunar equitoral orbit or Lunar polar orbit. The polar orbit requires more delta V. “1-way” transit times range from 72–24 hours are considered. Faster transit times are avoided, because they preclude Free-return Trajectories and thus are more unsafe. Meaning if the engine malfunctions the ship goes sailing off into the wild black yonder and the crew dies a lonely death.

Sampling of LANTR Vehicle Types
Case DescriptionObjectiveTrajectory/OrbitsIn-line LO2 TankResults
1cCrewed LANTR LTV
with MPCV
and 12 m saddle truss
carrying 5 t cargo to LLO
Determine LLO2 refueling needed
to deliver 5 t cargo to LLO
72 hour 1-way transit times
LEO–LLO–LEO
ΔV ~7.984 km/s
7.6 m OD x ~5.23 m L
(~163.5 t LO2)
IMLEO ~ 152.4 t
~48.8 t LO2 supplied in LEO
~46.9 t LLO2 refueling in LLO
2cCrewed space-based LANTR LTV
with 9.9 t habmodule
and 11 m star truss
carrying 5 t cargo to LLO
Determine LLO2 refueling needed
to deliver 5 t cargo to LLO
using alternative LTV configuration
72 hour 1-way transit times
LEO–LLO–LEO
ΔV ~7.996 km/s
4.6 m OD x ~3.4 m L
(~35.9 t LO2)
IMLEO ~ 131.1 t
~35.9 t LO2 supplied in LEO
~35.1 t LLO2 refueling in LLO
3cCrewed space-based LANTR LTV
with 9.9 t hab module
and 11 m star truss
carrying 5 t cargo to LLO
Determine LLO2 refueling needed
to deliver 5 t cargo to LLO
while also cutting transit times to 48 hrs
48 hour 1-way transit times
LEO–LLO–LEO
ΔV ~8.695 km/s
4.6 m OD x ~4.1 m L
(~48.0t LO2)
IMLEO ~ 143.4 t
~48.0 t LO2 supplied in LEO
~47.0 t LLO2 refueling in LLO
4cCrewed space-based LANTR LTV
with 9.9 t hab module
and 11 m star truss
carrying 5 t cargo to LLO
Determine LLO2 refueling needed
to deliver 5 t cargo to LLO
while also cutting transit times to 36 hrs
36 hour 1-way transit times
LEO–LLO–LEO
ΔV ~9.838 km/s
4.6 m OD x ~6.1 m L
(~81.2 t LO2)
IMLEO ~ 177.4 t
~81.2 t LO2 supplied in LEO
~71.6 t LLO2 refueling in LLO

Cases use a “Common NTPS” (carries ~39.7 t LH2). Propellant depots assumed in LEO, LLO and LPO. LANTR engines use optimized MRs. LEO=407 km Low Earth Orbit, LLO=300 km equatorial Low Lunar Orbit, LPO=300 km polar Lunar Polar Orbit. Total round trip mission ΔV values shown include g-losses

Case 1, a crewed LTV mission, carrying the Orion MPCV and 5 t of cargo (shown here, bottom ship), uses an oversized in-line LO2 tank consisting of two 7.6 m diameter ellipsoidal domes and requires ~47 t of LLO2 for Earth return.

Case 2 is a space-based crewed cargo transport (shown here, upper ship). It has its own dedicated habitat module weighing ~10 t, plus a 4-sided, concave star truss that has attached to it four 1.25 t PL pallets. The LO2 tank is smaller and customized for this particular application resulting in a lower IMLEO and LLO2 refueling requirement (~35 t).

Cases 3 and 4 show the impact on the crewed cargo transport mission of reducing the Earth-Moon transit times from 72 hours down to 48 and 36 hours, respectively. Because the LH2 propellant loading in the NTPS is fixed at ~39.7 t for these missions, the LANTR engines run “O2-rich” on the return leg (Mass Ratio = 5, Isp ~516 s) so the LLO2 refueling requirement for Case 4, with a 36-hour transit time, increases to ~71.6 t – more than double that needed for Case 2.

NLTV Commuter Shuttle

ItemCommuter Shuttle
MissionLEO⇒LLO⇒LEO
Duration36-hr “1-way” transit times
Habitat Modulen/a
Passenger
Transport
Module
15.2 t
Crew2
Passengers18
Star Truss
w/ 5 t payload
n/a
In-line
LO2 tank
~74.5 t
LH2 NTPS~70.9 t
IMLEO
(wet mass)
160.6 t
Refueled LLO2~67.9 t
Total
Burn Time
~25.3 min

This is the second of the two optimized LANTR designs: the Commuter Shuttle

The Commuter Shuttle carries a forward Passenger Transport Module (PTM) that contains its own life support, power, instrumentation and control, and reaction control system. It provides the “brains” for the LANTR-powered shuttle which is home to the 18 passengers and 2 crew members while on route to the Moon. Arriving in Low Lunar Orbit (LLO, 300 km altitude), the PTM detaches and docks with a waiting “Sikorsky-style” Lunar Landing Vehicle (LLV) that delivers it to the lunar surface. From here the PTM is lowered to a “flat-bed” surface vehicle for transport over to the lunar base and passenger unloading.

All the missions start and end in LEO, with the mid-point being either Lunar equitoral orbit or Lunar polar orbit. The polar orbit requires more delta V. “1-way” transit times range from 72–24 hours are considered. Faster transit times are avoided, because they preclude Free-return Trajectories and thus are more unsafe. Meaning if the engine malfunctions the ship goes sailing off into the wild black yonder and the crew dies a lonely death.

Sampling of LANTR Vehicle Types
Case DescriptionObjectiveTrajectory/OrbitsIn-line LO2 TankResults
5s LANTR commuter shuttle
carrying 15 t Passenger Transport Module (PTM)
to LLO then back to LEO
Determine LLO2 refueling needed
to deliver the PTM to and from LLO
with transit times of 36 hrs
36 hour 1-way transit times
LEO–LLO–LEO
ΔV ~9.835 km/s
4.6 m OD x ~5.4 m L
(~69.3 t LO2)
IMLEO ~ 160.6 t
~69.3 t LO2 supplied in LEO
~67.9 t LLO2 refueling in LLO
6sLANTR commuter shuttle
carrying 15 t PTM to LPO
then back to LEO
Determine LLO2 refueling needed
to deliver the PTM to and from LPO
with transit times of 36 hrs
36 hour 1-way transit times
LEO–LPO–LEO
ΔV ~10.006 km/s
4.6 m OD x ~6.0 m L
(~80.0 t LO2)
IMLEO ~ 172.5 t
~80.0 t LO2 supplied in LEO
~72.1 t LLO2refueling in LLO
7sLANTR commuter shuttle
carrying 15 t PTM to LPO
then back to LEO
Determine LLO2 refueling needed
to deliver the PTM to and from LPO
NTPS tops off with excess LLH2
36 hour 1-way transit times
LEO–LPO–LEO
ΔV ~10.047 km/s
4.6 m OD x ~4.6 m L
(~56.4 t LO2)
IMLEO ~148.2 t
LTV refuels with ~55.3 t LLO2
and NTPS tops off with ~6.9 t excess LLH2
8sRapid commuter shuttle
carrying 15 t PTM to LPO
then back to LEO
Determine feasibility of 24 hour transits
using twin LANTR engines
NTPS tops off with excess LLH2
24 hour 1-way transit times
LEO–LPO–LEO
ΔV ~13.225 km/s
4.6 m OD x ~8.3 m L
(~116.6 t LO2)
IMLEO ~204.3 t
LTV refuels with ~105.6 t LLO2
and NTPS tops off with ~13.2 t excess LLH2

Cases use a “Common NTPS” (carries ~39.7 t LH2). Propellant depots assumed in LEO, LLO and LPO. LANTR engines use optimized MRs. LEO=407 km Low Earth Orbit, LLO=300 km equatorial Low Lunar Orbit, LPO=300 km polar Lunar Polar Orbit. Total round trip mission ΔV values shown include g-losses

Case 5 is a commuter shuttle LTV that carries a 15 t PTM to LLO and back, has 36-hour 1-way trip times, and uses only Earth LH2. It has an IMLEO of ~161 t and refuels with ~68 t of LLO2.

Case 6 is similar to Case 5 but operates between LEO and Lunar Polar Orbit (LPO). Because of the higher DV budget needed to access LPO, the shuttle’s IMLEO and LLO2 refueling requirements are larger at ~173 t and ~72 t, respectively. The total burn time on the LANTR engines for the round trip mission is ~25.3 minutes. Also, with the engines running O2-rich and producing ~170.3 klbf of total thrust, the g-loading on the passengers during the final EOC burn varies from ~0.75 to ~1.5g.

Case 7 shows the benefit of utilizing the excess LLH2 produced from the depot’s H2O electrolysis process to top off the NTPS’ LH2 tank. By supplying the commuter shuttle with just under 7 t of LLH2, LLO2 refueling decreases by ~17 t and the shuttle’s IMLEO decreases by more than 24 t.

By switching to a “twin engine” NTPS, and again topping off with ~13 t of excess LLH2, 24-hour 1-way transit times are also possible as shown in Case 8. This rapid shuttle capability comes at the expense of increased mission DV (~13.2 km/s), IMLEO (~204 t) and LLO2 refueling (just under 106 t), but the passenger g-loading during the EOC burn is more manageable varying from ~0.5 to ~1g.

NTPS

This is from The Nuclear Thermal Propulsion Stage (NTPS): A Key Space Asset for Human Exploration and Commercial Missions to the Moon (2014).

Yet another nuclear rocket report with Dr. Borowski as lead author. He continues to patently point out the many advantages and uses of nuclear thermal rockets, especially the "right-sized" SNRE-class engines. If the powers that be would just get over their terror of things atomic.

The report outlines a standard nuclear thermal propulsion stage (NTPS) then gives several sample spacecraft for various applications. Each spacecraft is a classic example of fundamental spacecraft design: the NTPS is the propulsion bus and the payload section is optimized for the given function. The NTPS is basically the resurgence of NASA's 1970 Reusable Nuclear Shuttle project. Which was a promising project before it got axed in 1973.

Today NASA does all its rocket designs using relatively safe chemical propulsion, but the elephant in the room is chemical ain't ever gonna get a specific impulse much higher than a pathetic 450 seconds. Solid-core nuclear thermal designs can do twice that without even working up a sweat. That really gives the dreaded Tyranny of the Rocket Equation a solid kick in the gonads, and allows the design of much more useful spacecraft.


PREVIOUS DESIGNS


GROUND RULES AND ASSUMPTIONS FOR NTPS MISSION AND PAYLOADS

Payload elements for the Lunar missions:
  • CREW WITH EVA SUITS: Four to Seven. Mass 800 kg to 1,400 kg.

  • LUNAR HABITAT MODULE: An instant lunar base. On wheels. Can support a four man crew for up to 180 days. Mass 67,400 kg.

  • INFLATABLE HABITAT MODULE: A TransHab or Bigelow Aerospace BA-330 module. Both have 18 months life support for six crew. Mass 18,400 to 31,600 kg (minus consumables).

  • HL-20 LIFTING BODY: wingless lifting body spacecraft used to transport crew to and from low Earth orbit. A miniature version of the Space Shuttle, carrying seven passengers. The HL-20 is the parent design of the Dream Chaser and Prometheus. Mass 11,675 kg.

  • MULTI-MISSION SPACE EXCURSION VEHICLE (MMSEV): A NASA modular space exploration vehicle. Put wheels on and it becomes a mobile base, put attitude jets on and it becomes a space pod. Crew of two, life support for two weeks. Mass 6,700 kg.

  • ORION MULTI-PURPOSE CREW VEHICLE (MPCV): spacecraft that is an advanced version of the Apollo Command and Service Module. Crew of four to six, with up to 21 days active crew time plus 6 months standby while crew is absent at Lunar base. Mass 13,500 kg.

  • LUNAR DESCENT ASCENT VEHICLE (LDAV): advanced version of the Apollo Lunar Module. This has a wet mass of 35,300 kg, dry mass of 14,400, LOX/LH2 engine with Isp around 450. 4,100 m/s of delta V in actual use, since 5,000 kg of surface payload is not carried back up.

  • SADDLE TRUSS: spacecraft backbone with one side open to allow docking of auxiliary spacecraft or the jettisoning of spent propellant tanks. Mass 2,890 kg.

  • TRANSFER TUNNEL: used inside saddle truss to provide docking port for the MPCV, LDAV, MMSEV and/or inflatable habitat module; and a pressurized tunnel connecting the two. Crewed landing mission uses tunnel to connect MPCV and LDAV. Asteroid exploration mission uses tunnel to connect MMSEV and inflatable habitat module. Mass 600 kg.


GROUND RULES AND ASSUMPTIONS OF NUCLEAR THERMAL PROPULSION STAGE

The nuclear thermal propulsion stage has a three-engine cluster of SNRE-class engines. Each has a specific impulse of 900 s (exhaust velocity 8,829 m/s), thrust of 73,000 N (16.5 klbf), and a mass of 2,400 kg. Each contains 59.6 kg of uranium-234 fuel with 93% enrichment. The propellant mass flow is 8.40 kg/s and the engine thrust-to-weight ratio is 3.06. The over-all length is 6.1 meters including the nozzle skirt extension.

The basic spacecraft for the Lunar missions is built around a core nuclear thermal propulsion stage plus an in-line LH2 propellant tank. The Near Earth Asteroid (NEA) mission uses a saddle truss with a LH2 drop tank instead of an in-line tank. More delta-V is needed for the NEA mission, so excess weight has to be jettisoned.

The core stage tank is 15.7 meters long and has a propellant capacity of 39,800 kg LH2. The additional in-line tank size varies according to the mission from 15.7 meters (same as core tank) to 18.7 meters long, the longer tank's propellant capacity is 49,000 kg LH2. Note the 15.7 m tank is in a stage that is a total of 20.7 m, and the 18.7 m tank is in a stage that is a total of 23.7 m.

Not all the propellant is available. 3% of the usable LH2 is reserved for reactor cooldown, 2% of total tank capacity is the tank trapped residuals which are unavailable, and there is a 1% ΔV performance reserve for safety. So if my slide rule is not lying to me, the 39,800 kg tank has 39,000 kg useable (less trapped residuals) and 37,830 kg after reserving the reactor cooldown propellant. Then less the 1% ΔV performance reserve for the given mission.

The propellant tanks are constructed of aluminum, and are cladded in a combination foam/multilayer insulation (MLI) system for passive thermal protection (i.e., to shade the tanks from the awful heat from the sun). This gives the tank that characteristic "gold foil" look. It ain't really gold, it is actually a thin layer of aluminum sprayed on the inside of a sheet of thin yellowish-gold polyimide plastic.

The tank that is actually connected to the engines has a zero-boil-off (ZBO) "reverse turbo-Brayton" cryocooler system to keep the blasted liquid hydrogen from boiling away over the course of the mission. The heat radiator is the black band at the fore end of the tank. The additional in-line LH2 propellant tank has no ZBO cryocooler, since the tank is drained at the start of the mission during the Trans-Lunar Insertion maneuver. It won't have time for any of the LH2 to boil away.

Two circular solar photovoltaic arrays supply all the electrical power needed, mostly for the cryocoolers (5.3 kWe). The array provides 7 kWe at a distance of 1 AU from Sol. The array has a surface area of 25 m2 and a mass of 455 kg.

The Reaction Control System (RCS) Advanced Material Bipropellant Rocket (AMBR) attitude jets use a storable bipropellant fuel: NTO (Nitrogen Tetroxide) / N2H2 (Diimide). Jets have a thrust of 890 N and an Isp of 335 sec. Half of the jets are located on the fore end of the integral tank attached to the engines. The other half of the RCS jets are located just aft of the payload. On the Lunar mission ship this means on the fore end of the additional in-line propellant tank. On other ships this is on the fore end of the saddle truss just aft of the payload.


LUNAR CARGO AND CREWED LANDING MISSION

LUNAR CARGO DELIVERY
Mass Schedule
NTPS70,000 kg
Small In-Line LH2 Tank56,600 kg
Lunar Habitat Lander61,100 kg
Connection3,000 kg
IMLEO186,700 kg
Propellant79,400 kg
Height
NTPS26.8 m
Small In-Line LH2 Tank20.7 m
Lunar Habitat Lander12.9 m
TOTAL60.4 m
ΔV
Dry Mass107,300 kg
Propellant79,400 kg
Wet Mass186,700 kg
Mass Ratio1.74
Isp900 sec
Exhaust Velocity8,829 m/s
Max ΔV4,890 m/s
(doesn't take into account
habitat jettison)
MISSION
ManeuverBurn TimeΔV
Burn 1: Trans-Lunar Injection perigee 121.4 min
Burn 2: Trans-Lunar Injection perigee 215.5 min3,214 m/s
Burn 3: Lunar Orbit Capture8.0 min906 m/s
Burn 4: Trans-Earth Injection3.1 min857 m/s
Burn 5: Eccentric Earth Orbit Capture1.2 min366 m/s
TOTAL49.2 min5,343 m/s

This configuration uses the shorter 20.7m/39,800 kg LH2 in-line tank. This is because pretty much all the cargo remains on the Lunar surface, none of it gets lugged back to Terra.

The resuable Lunar cargo delivery mission departs from LEO (C3 or bare minimum escape velocity of -1,678 m2/s2) into Trans-Lunar Insertion requiring a delta-V (ΔVTLI) of 3,214 m/s (including a g-loss of 117 m/s).

About 72 hours later (3 days) arrives at Luna with an arrival Vinf (V) of 1,151 m2/s2. It captures into a 300 km circular Low Lunar Orbit (LLO) requiring a delta-V (ΔVLOC) of 906 m/s (including g-loss).

The key phases of the uncrewed Lunar cargo delivery mission are shown below:

The habitat landers use LOX/LH2 chemical engines to reach the Lunar surface. There they use the included wheels to move to optimal locations and link up with other habitats.

After the lander departs, the LNTR cargo transport spends a day in LLO. Then it departs from LLO (C3 945 m2/s2) into Trans-Earth Injection burn requiring a delta-V (ΔVTEI) of 857 m/s (including a g-loss).

72 hours later it arrives at Terra with an arrival V of 1,755 m2/s2. It captures into a 24-hour Eccentric Earth Orbit (EEO) requiring a delta-V (ΔVEOC) of 366 m/s. The post-burn engine cool-down thrust is used to lower the orbit a bit. A tanker vehicle operating from a LEO servicing node/orbital propellant depot does a rendevous with the cargo transport, and fills it up with enough LH2 so that the transport can circularize into LEO orbit.


CREWED LUNAR LANDING
Mass Schedule
NTPS70,000 kg
Large In-Line LH2 Tank63,300 kg
Saddle Truss6,400 kg
wet LDAV29,500 kg
LDAV payload5,000 kg
MPCV13,500 kg
Consumables100 kg
x4 crew w/Suits800 kg
IMLEO188,600 kg
Propellant88,700 kg
(39,700+
49,000)
Height
NTPS26.8 m
Large In-Line LH2 Tank23.7 m
Payload26.8 m
TOTAL77.3 m
ΔV
Dry Mass99,900 kg
Propellant88,700 kg
Wet Mass188,600 kg
Mass Ratio1.89
Isp900 sec
Exhaust Velocity8,829 m/s
Max ΔV5,610 m/s
(doesn't take into account
payload jettison)
MISSION
ManeuverBurn TimeΔV
Burn 1: Trans-Lunar Injection perigee 120.9 min
Burn 2: Trans-Lunar Injection perigee 216.2 min3,214 m/s
Burn 3: Lunar Orbit Capture8.2 min913 m/s
Burn 4: Trans-Earth Injection6.9 min856 m/s
Burn 5: Eccentric Earth Orbit Capture2.8 min366? m/s
TOTAL55 min5,349 m/s

The key phases of the crewed Lunar landing mission outbound mission leg are shown below:

The crewed vehicle does not just have the cargo stuck on the nose of the spacecraft. Additional equipment is required for the health and well-being of the crew. The unmanned ship does not need life support and other things important for squishy humans.

Besides the crew, the mission payload is the Lunar Descent Ascent Vehicle (LDAV). As previously mentioned this is a highly advanced version of the old Apollo Lunar Module. It delivers the crew from the orbiting spacecraft to the lunar surface, then back again at the end of the lunar stay.

In LEO, the crew is transported to and from the spacecraft in an Orion MPCV which is an advanced version of the Apollo Command and Service Module. It is boosted into orbit with the crew, docks with the spacecraft, acts as a habitat module for the trip, and at the end of the mission separates from the spacecraft then aerobrakes to land on Terra.

To accommodate the MPCV, a saddle truss is used. The truss provides a nook for the MPCV to inhabit, a docking port and transfer tunnel connecting the MPCV with the LDAV, and photovoltaic arrays to energize the MPCV's life support system. It also has additional RCS jets. The MPCV does have its own photovoltaic arrays but they are difficult to deploy when inside the nook.

Unlike the uncrewed mission, the crewed mission carries more mass back to Terra (saddle truss, MPCV and LDAV). It needs more propellant, so the longer 23.7 m/49,000 kg LH2 in-line tank is used.

The resuable Lunar crew transfer mission departs from LEO (C3 or bare minimum escape velocity of 1,516 m2/s2) using a 2-perigee burn into Trans-Lunar Insertion requiring a delta-V (ΔVTLI) of 3,214 m/s (including a g-loss of 117 m/s).

About 72 hours later (3 days) arrives at Luna with an arrival Vinf (V) of 1,217 m2/s2. It captures into a 300 km circular Low Lunar Orbit (LLO) requiring a delta-V (ΔVLOC) of 913 m/s (including g-loss).

The key phases of the uncrewed Lunar crew transfer mission are shown below:

LUNAR DESCENT ASCENT VEHICLE
LDAV
Mass Schedule
Inert Mass6,100 kg
Payload: Crew Cabin2,200 kg
Payload: to Surface5,000 kg
Payload: x4 Crew
with EVA Suits
800 kg
DRY MASS14,100
LOX/LH2 Fuel20,900 kg
WET MASS35,000 kg
ΔV
Mass Ratio2.48
Isp450 sec
Exhaust Vel4,420 m/s
ΔV4,014 m/s
(actually 4,100
surface payload is unloaded)
ΔV Budget
Descent Start Mass35,000 kg
Descent Fuel Burnt13,400 kg
Descent End Mass21,600
Descent Mass Ratio1.62
Descent ΔVdes2,115 m/s
Ascent Start Mass15,100 kg
(-5,000 surface payload
-1,500 consumables)
Ascent Fuel Burnt5,500 kg
Ascent End Mass9,600
Ascent Mass Ratio1.57
Ascent ΔVasc1,985
TOTAL ΔV4,100 m/s

In LLO, the crew enters the LDAV and undocks. The surface payload containers rotate 180° into their landing position. They are stored snug to the crew cabin so that the LDAV has a small enough diameter to fit into the booster payload faring. They are rotated so on the lunar surface the payload is next to the ground for easy access, instead of inconveniently one story up in the air out of reach.

On the lunar surface, the crew can operate for three to 14 days using life support consumables carried in the surface payload containers. Alternatively they can operate for up to 180 days using one of the habitat landers. Upon departure they can carry back to the orbiting spacecraft up to 100 kg of lunar samples. The surface payload containers are left behind.

The LDAV climbs into orbit and does a rendezvous with the ship. It docks with the transfer tunnel on the saddle truss and the crew transfers to the MPCV. Then the ship departs from LLO (C3 of 949 m2/s2) into Trans-Earth Injection burn requiring a delta-V (ΔVTEI) of 856 m/s (including a g-loss).

72 hours later it arrives at Terra and captures into a 24-hour Eccentric Earth Orbit (EEO). The MPCV separates from the ship and the crew returns to Terra in the command module via aerobraking.


ASTEROID EXPLORATION MISSION

This is a crewed mission to explore Near Earth Asteroids like 2000 SG344. This spacecraft could support a month-long mission to an asteroid at Earth-Moon Lagrange 2, a reusable 327-day mission to 2000 SG344, or a non-resuable 178-day mission to the same.

The E-ML2 mission is 9.8 days outbound, 5.8 day stay exploring the asteroid, and a 17.4 day inbound transit. Total mission delta-V is 5,150 m/s including gravity losses, plus lunar flyby impulsive burns on both outbound and inbound legs.

The spacecraft uses an inflatable TransHab as a habitat module, so the poor astronauts don't have to live for months to years inside a cramped MPCV with less living space than three phone booths. Kids, ask your grandparents what a phone booth was, and why Superman got arrested for indecent exposure.

The spacecraft carries a MMSEV space pod, which is a sort of space-going Alvin submarine that was born to explore asteroids.

The far asteroid mission carries a MPCV, while the near asteroid mission carries a HL-20 Lifting Body instead. Either is to deliver the astronauts from Terra to the spacecraft, and to aerobrake them back to Terra Firma at the end of the mission.

The total spacecraft has an IMLEO wet mass of 170,800 kilograms, and a length of 79.3 meters. With 222,400 Newtons total from the engines at an Isp of 900 s, the 5,150 m/s delta V can be performed with a total burn time of 48.3 minutes.


SPACE TOURISM MISSION


COMPARATIVE SIZE GALLERY

NTR FIRST LUNAR OUTPOST

NTR FIRST LUNAR OUTPOST
Payload96,000 kg
EngineSolid core NTR
Pewee-class
Specific Impulse900 s
Exhaust Velocity8,800 m/s
Thrust110,000 N
Engine Mass3,437 kg
Shadow Shield Mass1,500 kg
Num Enginesx3
Total Thrust330,000 N
Total Engine Mass14,810 kg
Inert Mass33,330 kg
Dry Mass129,330 kg
Propellant Mass67,570 kg
Wet Mass196,900 kg
Mass Ratio1.5
ΔV3,658 m/s
MASS SCHEDULE
Structural Mass13,360 kg
Avionics & Power1,000 kg
Reaction Control460 kg
x3 Engines10,310 kg
x3 Shadow Shields4,500 kg
Contingency3,700 kg
INERT MASS33,330 kg
Payload 1
(FLO Lander)
93,000 kg
Payload 2
(FLO Adapter)
3,000 kg
Total Payload96,000 kg
DRY MASS129,330 kg
LH2 Propellant66,540 kg
RCS Propellant1,030 kg
Total Propellant67,570 kg
WET MASS196,900 kg

This is from Nuclear Thermal Rocket/Vehicle Design Options for Future NASA Missions to the Moon and Mars (1995)

First Lunar Outpost (FLO) was one of NASA's "reference missions" studies. It was created in 1992. As with the other reference missions the mission parameters were nailed down, and researchers could design spacecraft capable of carrying out said missions. It got the ax shortly after 1992 for a variety of reasons.

The payload is 96 metric tons of lunar lander. This is 60 metric tons of lander stage which carries 36 tons of either: [a] cargo, [b] surface habitat, or [c] manned crew module with ascent/Terra-return stage.

The standard designs assumed that the lander would be transported to Lunar orbit by a conventional chemical propulsion module based around J-2S engines. But Stanley Borowski et al figured mission could be performed much more economically by using solid core nuclear rocket engines. The single J-2S chemical engine had a great thrust of 265 kilopounds force (klbf) (1,180,000 newtons) but a crummy specific impulse of 436 seconds (4,300 m/s exhaust velocity). A trio of NERVA derivative rocket (NDR) engines would only have a combined thrust of 75 klbf (25 klbf each) (330,000 newtons) but a much better Isp of 900 s (8,800 m/s Ve).

The nuclear stage carries 66.5 metric tons of liquid hydrogen propellant.

Bottom line is that the chemical stage had wet mass of 155 metric tons but an equivalent nuclear stage was only 101 metric tons. A savings of 54 metric tons is nothing to sneeze at. The nuclear stage is four meters longer than the chemical stage, but who cares?

After the lander detaches from the nuclear stage, the latter uses the RCS system to do a trailing edge lunar swingby. This provides enough of a gravity assist to put the spent stage into a disposal heliocentric orbit that will keep it away from Terra for at least one hundred thousand years.

"One Burn" Lunar Scenario
Trans-Lunar Injection (TLI) Payload
96 MT (pilot vehicle and TLI stage adaptor)
TLI Maneuver
ΔV3,200 m/s + gravity losses
Initial Orbit185 km circular LEO
NTR System
PropellantCryogenic hydrogen
Isp870 sec (graphite)
900 sec (composite)
960 sec (ternary carbide)
External shield mass60 kg/klbf thrust
(0.014 kg/Newton)
Burn Duration≤ 30 minutes
Flight Performance
Reserve
1% usable propellant
Cooldown (effective)3% usable propellant
Residual1.5% total tank capacity
RCS System
PropellantHydrazine
Isp237 seconds
TLI Burnout ΔV60 m/s
(30 m/s for trailing edge lunar flyby)
RCS System
Material2219-T87 Al
Geometry10 m diameter cylindrical tank
with √2/2 domes
Insulation2 inch MLI +
micrometeoroid shield (3.97 kg/m2)
Boiloff12.40 kg/day
Contingency
Engine &
external shields
15%
All other dry masses10%

Blue Max Studio Liberty Bell

Liberty Bell
ParamLaunchLunar Run
Structure
Mass
15075
Cargo
Capacity
250300
Dry
Mass
400375
Wet
Mass
820585
Thrust
(engine)
6.18 MN2.2 MN
Mass
Ratio
10.5 (4.2)*1.56
Accel10 m/s24.58 m/s2
Mass
Flow
1,132.4 kg/s1,000 kg/s
Exhaust
Velocity
5,457 m/s2,200 m/s
Specific
Impulse
556.2 s452.3 s
Burn
Duration
770 s105 s
Flight
Time
≈2 hrs5.56 days
Δv7,842 m/s4,947 m/s

The Liberty Bell is a tramp freighter created by Ray McVay for his Black Desert universe

The Liberty Bell proper is a command module with a dry mass of 50 tons, and 50 tons of propellant. It has a power plant, life support, and thrusters. It can carry a crew of five plus up to 20 passengers from the surface into LEO.

On the nose is an airlock with an androgynous docking port and a maneuvering unit.

On the tail there are four couplers, each of which can hold one cargo container. The containers are cylinders 9.5 meters long and 5 meters in diameter. They are rated to carry a maximum of 62.5 tons of cargo each.

There are four remote manipulator arms used to handle cargo containers. The arms are not permanently attached, they can move like a giant inchworm over the spacecraft's surface just like the Canadarm 2 on the International Space Station.


The Liberty Bell is boosted into orbit with an L-Drive assembly. This is a laser launch system. At the spaceport, the launch pad has a huge stationary laser built into it. The L-Drive assembly is attached to the bottom of the Liberty Bell. The L-Drive is an air-breather, it scoops up atmosphere and sprays it into the mirrored dish-with-a-spike. The laser beam from the launch pad heats the air, creating the thrust to boost the spacecraft into orbit. The laser beam tracks the L-Drive as it climbs into the sky. When the L-Drive reaches an altitude where the air is too thin, it switches to its internal propellant tanks.

Typically the L-Drives are owned and maintained by the spaceport, they cost $1,250,000 Black Desert dollars. The spaceport will rent an L-Drive, laser boost time, plus fees and taxes to the captain of the Liberty Bell. This will cost the captain $100,000 total to boost the Liberty Bell into LEO.


Upon reaching LEO, a Liberty Bell generally makes a rendezvous with an orbital transport nexus, unloads its four cargo containers (250 tons of cargo total) and 20 passengers, loads new cargo and new passengers to be delivered to Terra's surface, pays the spaceport for laser landing services (including fresh propellant for the L-Drive), and rides the laser beam back down to the spaceport.

However, our Liberty Bell is heading to Luna.

The Liberty Bell jettisons the L-Drive, delivering the rental vehicle back into the hands of spaceport personnel (the orbital representatives). The captain knows that when they make the return trip, the spaceport will be more than happy to reserve them an L-Drive for the trip down.

On this trip, instead of carrying four cargo containers, the Liberty Bell only has two containers (125 tons), a translunar rocket engine (20 tons, thrust equivalent to a SSME), and a small cobbled together weapons package (105 tons). The total payload tonnage is 250 tons, same as four cargo containers.

The weapons package contains two Kinetic Kill Vehicles (KKV) at 40 tons each, two Caltrop space mines at ten tons each, and a laser turret with power supply at five tons.

The Liberty Bell then moves into a higher orbit, to make a rendezvous with a transfer space station. In the Black Desert universe, the orbits are patrolled by the astromilitaries of various nations, all looking for trouble and whatever they can get away with. This is the main reason for the Liberty Bell's weapons package.


At the transfer station, the Liberty Bell outfits itself for the Lunar trip. It leases four propellant tanks to feed the translunar rocket engine. It also leases or purchases a cupola.

Using the remote manipulator arms, the translunar rocket engine and the airlock/docking ring swap positions. The rocket engine is mounted on the nose and the four propellant tanks are attached. The docking ring is mounted next to the other cargo, and a cupola installed on top. For the rest of the trip, the cupola will serve as the Liberty Bell's cockpit.

As it turns out, one of the captain's business partners had three cargo containers waiting at the transfer station to be delivered to Luna. The remote manipulator arms install these as well.


The Liberty Bell is ready for the trip to Luna. The command module now faces opposite the direction of thrust it had at launch, with the cupola and the weapons package aimed at the new forwards that used to be backwards. It is carrying three hundred tons of cargo.

It has enough life support and consumables to haul five crew and twenty passengers on the five and a half day trip to Luna or one of the La Grange stations.

Closed-Cycle MHD Nuclear-Electric

This is from Prospects for Nuclear Electric Propulsion Using Closed-Cycle Magnetohydrodynamic Energy Conversion (2001).

NASA’s Dr. Ernst Stuhlinger, a leading authority on electric (ion) propulsion, has often said that such a rocket system would be ideal for a manned journey to Mars.

“Yeah,” a wag once cracked, “if you can just find an extension cord long enough."

From A FUNNY THING HAPPENED ON THE WAY TO THE MOON by Bob Ward (1969)

What the joke is saying is that Electrostatic (ion drives) and Electromagnetic (VASIMR) rockets are power hogs. While they have outstanding exhaust velocity/specific impulse, they need huge solar photovoltaic arrays or nuclear reactors whose mass is measured in metric tons. Which really cuts into your payload mass.

Photovoltaics are an attempt to use Sol as the power plant and sunlight as the extension cord. Trouble is that sunpower is relatively dilute, and the inverse square law shortens the length of the extension cord to about the orbit of Mars. A "robust" mission at any rate.

This design is an attempt to reduce the mass of a nuclear power plant so it can be used in an ion-drive ship without reducing the payload mass to a couple of kilograms.


MHD Nuclear-Electric
General
Payload100,000 kg
MissionTerra to Mars
Mission Durations
  • 120 days
  • 150 days
  • 180 days
Power Plant
TypeClosed-loop nuclear MHD Brayton cycle
Reactor Power100 MWth
Enthalpy Extraction Ratio40%
Isentropic Efficiency70%
Power Output40 MWe
Propulsion System
Propulsion Alpha1.1 kg/kWe
Isp5,000 to 8,000 sec
Engines
TypeIsp
(sec)
ηtα
(kg/kWe)
Ion (Kr)≥5,0000.81.0
MPD (Li)4,000—8,0000.50.5
MPD (H2)≥8,0000.50.5
VASIMR (H2)3,000—30,0000.50.2—1.0

NASA and other space agencies tend to focus on technologies that can be realized in the near term. I mean, antimatter power would be nice but that ain't gonna be available anytime short of a century (i.e., 25 presidential election cycles, which is the average time between radical NASA policy shifts). However, the paper makes a case that expending some effort on a technology that is just a couple of steps over near term can give huge returns.

The paper makes the case that the penalty-mass problem with spacecraft nuclear power plants is due to the fact they are based on closed-cycle gas turbine technology. These are limited to low-heat rejection temperatures, which result in large and massive heat radiators. And just guess which nuclear power component is the major culprit affecting power plant mass? Yep, the heat radiators. The mass of the radiators is huge compared to the mass of the reactor and energy conversion equipment.

How do you reduce the mass of the heat radiators? You run them at a higher temperature, that's how. Why don't they do that? Well, gas turbines contain turbine wheels. If you run the system at a higher temperature the blasted turbine blades melt and the turbine is destroyed. Even if the turbine wheels have an active cooling system. Sort of like spitting on a blast furnace in order to cool it down. You don't want to run the turbine much hotter than 1,200 K or so.

This is why the authors of the paper say we should abandon closed-cycle gas turbine technology and make the jump to closed-cycle magnetohydrodynamic (MHD) technology. On the minus side this technology is not quite as mature. On the plus side you can run that sucker at 2,500 K with little or no problem (report says "minimum development risk") with a corresponding drastic reduction in heat radiator mass. If you did some work high-temperature fissile fuels for the reactor, you could push that to 3,000 K. And in the future if you developed gas core nuclear reactors, it is estimated that the theoretical limit of MHD generators is as high as 8,000 to 10,000 K. That will really shrink the heat radiators down to size.

You see, with a gas turbine, the turbine blades are bathing in the ultra-hot blast of gas. With an MHD on the other hand, none of it actually touches the gas, it just surrounds it. Which drastically reduces the "melting generator" problem. The MHD can be cooled with good old regenerative cooling, just like the nozzle of a chemical rocket. MHDs also have no moving mechanical parts, which improves reliability and reduces maintenance.

The main problem is that the gas has to conduct electricity, which generally means you have to seed it with cesium dust or something like that. Then it becomes electric charges moving through a magnetic field, which is the basis for all electrical generators. It is just that the electric charges are moving at hypervelocity so it generates lottsa current.

The fact that the gas is accelerated by a fission reactor opens up another seeding possibility. If you seed the gas with an isotope with a large neutron interaction cross section, as it passes through the intense neutron flux inside the reactor the isotope dust will create nuclear ionization events. Not just one or two as with chemical ionization, each nuclear interaction can produce hundreds to thousands of ionization events.


As a benchmark the report authors set up a sample space mission to demonstrate the performance of this propulsion system.

The mission was to deliver a 100 tonne payload from a 1,000 km circular Terran orbit (i.e., high enough so that the reactor radiation would not reach Terra) to a 500 km Mars orbit. Several 2018 mission opportunities were examined for trip times of 120, 150, and 180 days.

A 100 MWth nuclear reactor was assumed, driving an MHD generator with an enthalpy extraction ratio of 40% and an isentropic efficiency of 70%. This means it will generate 40 MWe for the ion drive. Using near-term technology assumptions for the subsystems, this implies an overal propulsion system specific mass of 1.1 kg/kWe.

Figure 14 shows the Initial Mass in Low Earth Orbit (IMLEO) for a nuclear MHD powered ship with the above specifications, over a range of engine Specific Impulse (Isp). The ranges of Isp for the four engines covered in the report are shown in color. The three black lines show the values for trip times of 120, 150, and 180 days. Example: if you had an engine with an Isp of 2,000 seconds, the 180 day transit would require an initial mass in LEO of about 270 metric tons.

The sweet spot seems to be with Isp between 5,000 and 8,000 seconds. Note that in that range the payload can account for as much as half the IMLEO.

The CHEBY-TOP software the writers used to figure the mission trajectories also had a function to determine the optimal power for a given configuration. So they gave it a try. Figure 16 shows the results for the 120 day mission, delivering a 100 metric ton payload. At a specific impulse fo 2,000 seconds the optimal power was 13.3 MWe. At 10,000 sec the optimal power was 30.6 MWe. Since the system was sized with a power level of 40 MWe, it turns out that the design is actually oversized for the mission. But that's OK, the extra power can be used.

The extra power can be used in two alternate ways: faster trip time and/or larger payloads.

Figure 17 shows the payload increase option. Here the trip time is still 120 days, the triangle line shows payload mass, the circle line shows IMLEO.

For instance, for an engine with an Isp of 2,000 sec, it could deliver 293 metric tons to Mars in 120 days. The drawback is that the IMLEO mushrooms to an unattractively monstrous 880 tonnes! An Isp of 10,000 sec can only deliver 132 metric tons of payload in 120 days, but the IMLEO is a much more reasonable 255 metric tons. Any Isp higher than 6,000 sec will have a payload mass fraction (payloadMass / IMLEO) greater than 0.56, which is pretty darn attractive for a 120 day mission.

Convair Electra

This is from Handbook of Astronautical Engineering, edited by Heinz Hermann Koelle, 1961. It describes Electra, a 1960 Convair/General Dynamics design for a manned ion-drive exploration spacecraft.

They proposed to use some sort of nuclear electric propulsion, but specifically they were looking closely at the ion drive designed for Ernst Stuhlinger's Mars Umbrella ship.

ELECTRA
Venus Mars♃ Jupiter
Departure orbit (Earth), km555
Capture orbit (target planet) km283.749
J-IV
Callisto
Total powered flight time (y)1.341.72.95
Initial acceleration (μg)10010060
Specific Impulse (sec)14,000
Specific jet power (kw/lbf)306
Initial weight (kg)127,000
Electric power for thrust (kwe)9,5009,5005,750
Total electrical power output (Mw)10106.5
Conversion efficiency0.25
Thermal reactor power (Mwt)404026
Electric-power-equipment weight (kg)†18,60018,60013,610
Radiator surface area (m2)1,9501,9501,300
Radiator planform area (m2)622622418
Radiator weight (kg)22,680
Reflector area (m2)‡1,2261,226836
Reflector weight (kg)2,9902,9902,270
Thrust device and equip (kg)§
(ion drive)
4,5404,5409,070
Propellant weight (kg)38,55548,53451,710
Payload weight (kg)◊39,46230,84427,670
Mass Ratio1.441.621.69
ΔV (m/s)50,29066,14071,930
Specific Power α (kwe/lb)0.08940.08790.091
μg = micro-g = 10-6 g
= Includes reactor, turbine, compressor, generator, and working fluid
= 5-ft-diameter inner cylinder and 4 reflector fins
§ Includes spare parts
Includes life-support systems

Since electric rockets need lots of electricity, and nuclear power generators have less than perfect efficiencies, this design will of course need extensive heat radiators to get rid of the waste heat. They were also overly concerned with radiator damage due to meteor strikes. If there is only one radiator and it gets holed by an errant meteor, all the radiator working fluid will escape into space and the crew will die in the cold and the dark. We now know that the danger of meteors is not quite as bad as the designers feared, so such an insane amount of redundancy is not needed.

In Fig. 24.39 the design has nine radiators, but in Fig. 24.40 this was latter reduced to four. The heat-shine penalty tells us that each panel only gets rid of the waste heat of 29% of a single unobstructed panel. So all four panes will get rid of 2.82 panels worth of waste heat instead of 4 panels worth. The entire heat radiator array has an efficiency of only 71%.

The Electra designers tried to avoid that by using heat reflectors (show in blue above) to shield each radiator (shown in red above) from the heat shine of the adjoining radiators. Basically each heat panel is placed in the focal plane of a cylindrical reflector which is just large enough to shadow the adjoining radiators. The reflectors also help shield the radiators from meteors. Like a Whipple shield, because the reflectors are going to be like metal foil, not armor plate. Every gram counts, and this is an exploration ship, not a warship.

The reflectors have to be carefully shaped and placed. If reflector A is preventing the heat shine of radiator B from hitting radiators C and D, it will be counterproductive if the reflector bounces the heat shine right back at radiator B. The report implies that this is impossible to avoid altogether. In addition, if the spacecraft travels near the Sun, the reflector is going to do its darnedest to focus all the Sun's heat right on the radiator. But there is always going to be some waste heat absorbed instead of radiated because there are no perfect infrared mirrors. Meaning the reflectors are going to soak up waste heat and transmit it into the spacecraft.

The report says that the four-fin reflector in Fig. 24.40 is a nice compromise between the monstrosity of Fig. 24.39 and having no protection from adjacent heat-shine at all.

The telescoping boom (B above) is mainly to get the habitat module as far from the nuclear reactor as possible. Unfortunately it cannot get it far enough away, a fault it shares with the A. C. Clark. The difference is that the Clark didn't let the artificial gravity centrifuge extend out of the shadow safe zone and into the deadly radiation zone. As you can see in the diagram above, the Electra was not so kind.

I think that part of the trouble is the Electra design team was so determined to have the ship easily perform the missions that they cut corners on the mass budget. Specifically they reduced the size of the heavy shadow shield which narrowed the angle of the shadow.

In almost every other spacecraft with artificial gravity (such as the A. C. Clark) they use a dependent centrifuge. The centrifuge's plane of rotation is normal to the ship's thrust axis.

In the Electra on the other hand, the centrifuge's plane of rotation is parallel to the ship's thrust axis. Why? If it was normal to the thrust axis, the centrifuge habitat modules would be in the radiation zone 100% of the time. If I am doing my math correct, having the rotation plane parallel means the centrifuge is in the radiation zone only 89% of the time. Not much of an improvement but they probably were happy to get any reduction at all.

So reduction of the radiation dose is the first reason to have the plane of rotation parallel to the thrust axis

The second reason to have the plane of rotation parallel to the thrust axis is to make the freaking ship easier to stablize under spin gravity. Now pay attention, this gets complicated. Most spin-grav ships are dependent centrifuges, that is, the entire ship spins, not just the centrifuge. For various reasons the Electra uses an independent centrifuge, where the ship is stationary while only the centrifuge spins.

Now, say the Electra had the centrifuge normal to the thrust axis. If the centrifuge bearings were sticking slightly, the transfer of angular momentum would make the spacecraft Roll (i.e., spin around the thrust axis). The pilot stops this by using the Roll attitude jets. The bigger the jet's mechanical advantage, the easier it is for the jets to stop the unwanted roll. And how do you give the roll jets a bigger mechanical advantage? The farther you can mount the blasted things from the thrust axis, the bigger the advantage.

If you look at the blueprints you can see the problem. Pretty much the maximum distance you can mount the roll attitude jets is 20' (6 meters) from the thrust axis. Not much of an advantage there. You will have to burn a huge amount of attitude jet fuel.

Now let's see what happens if the Electra had the centrifuge parallel to the thrust axis, as it does in the design. If the bearing stick, the transfer of angular momentum makes the spacecraft Pitch (i.e., entire ship spins around the centrifuge bearings like a US New Years Eve noisemaker). You put the pitch attitude jet as far away from the pitch axis as possible, which in this case is about 130'+179'=309' (94 meters). Which has almost sixteen times the mechanical advantage of the roll jets.

The third reason to have the plane of rotation parallel to the thrust axis is because if the centrifuge becomes unbalanced, it will cause the spacecraft to nutate. This is relatively easy for the navigation system to deal with. If the rotation plane was normal to the thrust axis the blasted spacecraft would precess instead, which is a nightmare for the nav system to cope with.

The fourth and final reason to have the plane of rotation parallel to the thrust axis is because it makes it easier to design the centrifuge so it can fold up like a telescope, and fit into the payload faring of the Helios boost vehicle in one piece. Otherwise you'll have to boost the components separately and attempt to assemble them in orbit. Which is a major headache in free fall.

Convair Helios

HELIOS Stage One
PropulsionChemical
Thrust12,000,000 newtons
Wet Mass700 metric tons
not including
Stage 2
Dry Mass32 metric tons
Body Diameter6 meters
Wingspan27 meters
HELIOS Stage Two
ΔV21,000 m/s
Specific Power57 MW/kg
(566,100 W/kg)
Thrust Power3.8 gigawatts
PropulsionSolid Core NTR
Thrust981,000 newtons
Exhaust Velocity7,800 m/s
Reactor Power2,600 MW
Wet Mass100 metric tons
Payload6.8 metric tons

HELIOS stands for Heteropowered Earth-Launched Inter-Orbital Spacecraft. Unfortunately "HELIOS" became a catch-all term for quite a few post-Saturn studies around 1963. This entry is about the 1959 version from Krafft Ehricke at Convair.

As you should recall, when dealing with a radioactive propulsion system the three anti-radiation protection methods are Time, Distance, and Shielding. A rocket cannot shorten the time, a burn for specific amount of delta V takes as long as it takes. Most designs use shielding, even though the regrettable density of shielding savagely cuts into payload mass.

But some designers wondered if distance could be substituted. The advantage is that distance has no mass. The disadvantage is it makes the spacecraft design quite unwieldy. You'd have to either put the propulsion system far behind the habitat module on a long boom, or more alarmingly have the propulsion system in front with the habitat module trailing on a cable. In theory the exhaust plume is not radioactive, so again in theory the habitat module can survive being hosed like that. The propulsion exhaust is poorly collimated so it is not like a spacecraft weapon is being directly aimed at the hab module.

There is no way this design would work as a warship. It would be like trying to run through a maze while carrying a ladder. If you made too tight a turn the tow cable will be subject to the "crack-the-whip" effect, the cable will snap, and the hab module will be shot into deep space like a stone from a shepherd's sling.

The break-even point is where the mass of the boom or cable is equal to the mass of the shadow shield. Past that point it is much less trouble just to use a standard shadow shield and deal with the mass.

This is the Waterskiing school of spacecraft design.


Dr. Ehricke design was two-staged. It has a liftoff mass of 800 metric tons, a diameter of 6 meters (omitting the delta wings) and a length of 60 meters.

The first stage was chemical powered since even in 1959 they knew nobody was going to allow a nuclear propulsion system to lift off from the ground. The lower stage has a delta wing, and will glide back to base after stage separation to be reused on future missions. The lower stage has a diameter of 6 meters, and a wingspan of 27 meters. Wet mass of 700 metric tons, dry mass of 32 metric tons, twin chemical engines with a combined thrust of 12,000,000 newtons. The first stage pilot rides in a little red break-away rocket in case the first stage has an accident. In which case it will just be too bad about the crew riding next to the nuclear reactor.

The first stage separates from the second at an altitude of about 50 kilometers when the velocity reaches 4.5 km/s. The corrugated coupler that held the two stages together falls away.

The second stage will use retrorockets to lower the habitat module on cables about 300 meters below the nuclear stage, then let'er rip. The second stage has a wet mass of 100 metric tons, the nuclear reactor has a power of 2,600 Megawatts, and a thrust of 981,000 newtons. Initial acceleration is 1 g.

When it comes to Lunar landing, the habitat module touches down, then the nuclear stage move down and sideways so it stays 300 meters away as it lands. HELIOS can deliver about 6.8 metric tons of payload to the Lunar surface, and stil carry enough propellant to make it back to LEO.

Dr. Ehricke does not give details above the return trip, but it would need to involve some sort of ferry rocket to retrieve the crew from Terra orbit. There is no way anybody would allow that radioactive doom rocket to actually land. Even if it could carry enough propellant. Dr. Ehricke Convair Space Shuttle would do nicely to retrieve the crew.


Nowadays most experts agree that a 300 meter separation from a 2,600 MW reactor is totally inadequate to protect the astronauts from a horrible radioactive death. I've heard estimates of a minimum 1,000 meter separation from a 1 MW reactor. For 2,600 MW you'd want a separation more like 14,000 meters, which probably has more mass than a conventional radiation shadow shield.

Convair Hyperion

Source [1]
(Booster + Sustainer)
Payload
(to orbit)
145,000 kg
Payload
(to Terra escape)
82,000 kg
Stage 1 enginechemical
(LOX/LH2)
Stage 1 thrust10,700,000 N
Stage 1 Isp
(sea level)
365 sec
Stage 1
exhaust vel sea
3,580 m/s
Stage 1 Isp
(vacc)
457 sec
Stage 1
exhaust vel vac
4,480 m/s
Wet Mass850,000 kg
Height85.4 m
Diameter8.54 m
Source [2]
(Booster only)
Stage 1 enginechemical
(LOX/LH2)
Stage 1
num engine
4
Stage 1 thrust13,700,000 N
Stage 1 Isp
(sea level)
365 sec
Stage 1
exhaust vel sea
3,580 m/s
Stage 1 Isp
(vacc)
457 sec
Stage 1
exhaust vel vac
4,480 m/s
Booster
Wet Mass
394,625 kg
Booster
Dry Mass
18,144 kg
Stage 1
Burn Time
70 sec
Height12 m
Diameter8.45 m
Span13 m
Source [3]
(Sustainer only)
Stage 2 engineNTR LH2
Stage 2 thrust5,782,680 N
Stage 2 Isp800 sec
Stage 2
exhaust vel
7,900 m/s
Sustainer
Wet Mass
453,592 kg
Sustainer
Dry Mass
110,000 kg
Height51 m
Diameter8.54 m
Span8.54 m
Source [4]
(Sustainer only)
Num Crew4
Stage 2
Height
39 m
Stage 3
Height
101 m
Total Height140 m
Stage 2 engineNTR LH2
Stage 2 thrust2,700,000 N
Stage 2 power10,000 MW
Stage 3 engineNTR LH2
Stage 3 thrust44,500 N
Stage 3 power170 MW
Leave Terra
propellant burnt
293,300 kg
Arrive Mars
propellant burnt
232,000 kg
Leave Mars
propellant burnt
100,000 kg
Arrive Terra
propellant burnt
21,000 kg
Source [5]
(Sustainer only)
Num Crew4
Trip Time347 days
Height100 m
Leave Terra
Orbital Altitude560 km
Payload26,400 kg
Propellant burnt293,000 kg
Initial Mass721,000 kg
Final Mass428,000 kg
Thrust2,900,000 N
Arrive Mars
Orbital Altitude1,900 km
Payload24,400 kg
Propellant burnt232,000 kg
Initial Mass414,000 kg
Final Mass181,700 kg
Thrust45,000 N
Leave Mars
Orbital Altitude1,900 km
Payload23,200 kg
Propellant burnt100,000 kg
Initial Mass161,000 kg
Final Mass61,010 kg
Thrust45,000 N
Arrive Terra
Orbital Altitude560 km
Payload9,100 kg
Propellant burnt21,000 kg
Initial Mass39,000 kg
Final Mass17,000 kg
Thrust45,000 N

This is from a 1959 study by Krafft Ehricke for Convair. Alas, details are sketchy, and some sources disagree with each other. Indeed some source disagree with themselves. In the table I separate the data as per the sources, so you can be as confused as I am.

The concept is a solid-core nuclear thermal rocket (the "Sustainer") that would do fast reconnaissance to Mars and Venus. A chemical booster lofts components to be assembled in orbit because even back then NASA was skittish about a nuclear-powered surface-to-orbit stage. In some versions the first stage is chemical, in others the first stage is a Helios nuclear engine.

The mission envisioned a fleet of three to four spacecraft, for mutual support.

24.184 Hyperion

      This is the prototype designation of the fast interplanetary reconnaissance vehicle, which is assembled in orbit from operational subunits, and uses the Helios engine in its hyperbolic escape stage. At a thrust of 650,000 lb (2,900,000 N) and an (adequate) initial acceleration of 0.3 to 0.4g, the required initial weight of the Hyperion ship lies in the range of 1.2 to 2 million lb (540,000 to 900,000 kg). Fast reconnaissance mission to Venus and Mars require ideal velocities (delta V) of the order of 60,000 to 90,000 fps (20,000 to 30,000 m/s). Tables 24.9 and 24.10 present typical missions.

     Table 24.11 presents characteristic data on the individual vehicles for Venus and Mars missions. The similarity in weight and size is expected to permit extensive standardization not only of propulsion systems, but of many other components as well.

     Figure 24.32 illustrates a Hyperion configuration.

     The vehicles start from a 300-n-mi orbit (600 km). In order to pass the Van Allen belt quickly, and in order to restrict the gravitational losses during the departure maneuver, a powerful booster thrust is required, which is provided by the Helios engine at a thrust of 650,000 lb or higher (2,900,000 N, thrust power 12 gigawatts). The booster engine is too powerful for the subsequent maneuvers and, therefore, is jettisoned, together with the booster tanks, after attainment of the required hyperbolic departure velocity. It appears possible to reduce the thrust of a nuclear engine to a small fraction of the design value, thus using the booster engine again during the later maneuver. However, it would be rather inefficient to carry the heavy high-thrust engine through the various propulsion maneuvers utilizing only a small fraction of its power. By jettisoning, the crew enjoys the advantage of a relatively "clean" vehicle during the outward bound leg of the flight (The booster tanks, representing the most irradiated, and hence, secondary-radiation-emitting part of the structure, will be dropped in any case. The gamma activation of the residual hydrogen will be small, since most radiation is absorbed by the booster hydrogen. Thus, only a small, if any, amount of radioactivity will remain after the jettisoning of the booster section.).

     The (cruise) engine used subsequently (after the huge radioactive Helios engine is jettisoned) is small enough (thrust 44,500 N, thrust power 170 megawatts) so that adequate protection can easily be afforded (meaning a relatively lightweight shadow shield that won't savagely cut into the payload allowance), and radiation danger to personnel and sensitive equipment is avoided. In jettisoning the booster tanks and the Helios engine, the large reactor is allowed to melt down, since attempts to retrieve this section would be impractical, the parts having high hyperbolic velocity (meaning the molten Helios engine goes off into an eccentric solar orbit to be a radioactive death trap for a few thousand years).

     The booster is mounted underneath the interplanetary vehicle. The Christmas-tree principle is applied to the design of the interplanetary vehicle proper. The tanks containing the hydrogen for the capture and escape maneuver near the target planet are arranged around a central stem, from which they can easily be separated after they are emptied. Since the tank material will be the source of activated radiation, and the major source of gamma and beta radiation after reactor shutdown, it will be jettisoned promptly (so the "Christmas tree" is the stem, and the hydrogen propellant tanks are the spherical tree ornaments hung on the tree branches for easy removal). The thrust load is taken by the stem, at the rear end of which a small nuclear engine is mounted (thrust 44,500 N, thrust power 170 megawatts). This arrangement permits the most effective shadow shielding of the stem during various power maneuvers (only the four main nuclear maneuvers near earth and the target planet are considered. For corrective maneuvers chemical thrust is provided). After the target-planet capture maneuver, the propellant tanks for the subsequent maneuvers are progressively farther removed from the reactor. In this manner, tank irradiation and heating following reactor shutdown are minimalized. The tank for the arrival-earth maneuver near earth is not jettisonable. It is used to provide the bulk of the space-radiation shield for the emergency life-support system (the storm cellar). The tank stays filled until the space vehicle has returned through the radiation belt close to the earth. If, before the final maneuver, emergencies should require that crew members work near the tank, the exposure effect is minimized by the distance of the tank, at which now most of the final connections, valves, and control instruments are located. The reactor itself and the structure will have "cooled down" considerably since the escape manuever from the target planet.

     The stem, being approximately 6 ft in diameter (1.8 meters), serves as a conduit for pipelines, cables, and instruments. Their concentration around the center line enhances the effectiveness of shadow shielding for a given weight of shielding material, and their location inside the conduit provides mechanical protection from meteoritic material, as well as allowing some temperature control. In view of the diameter side, the interior of the stem is accessible for repair work. The conduit is divided into sections which can be individually pressurized to allow extended repair work to be carried out more conveniently without a pressure suit. The stem is the spine of the spacecraft in a real sense.

     The head is the life-support system (LSS). The emergency LSS is located in the center of the earth-capture tank, properly insulated from it (because liquid hydrogen is dangerously cold). Cylindrical enclosure, serving as a general (jettionable) life-support system, are in front of the emergency LSS. Air locks connect all parts of the LSS. Before the recapture maneuver near earth, the LSS is reduced to the shielded emergency section (because every gram counts), as indicated in Fig. 24.32.

     The entire vehicle is expected to rotate slowly during most of the transfer period about an axis normal to the longitudinal axis, to provide centrifugal weight for the convenience of the crew and for a number of technical reasons (i.e., it is a Tumbling Pigeon. During transfer the ship has spin radius of 32.5 meters. If it has 1/6g lunar gravity, it rotates at about 2 rotations per minute).

From HANDBOOK OF ASTRONAUTICAL ENGINEERING, edited by Heinz Hermann Koelle (1961)

Convair Solar Ship

General Dynamics Solar Ship
EngineSolar moth
Crewx2
PropellantLH2
Thrust per collector355 N
Total Thrust710 N
Exhaust temperature2,300 K
Exhaust Velocity4,330 m/s
Specific Impulse440 s
Dry Mass2,440 kg
Propellant mass5,000 kg
Wet Mass7,440 kg
Mass Ratio3.03
ΔV4,800 m/s
Thrust to weight ratio
initial
0.976×10-2
Thrust to weight ratio
maximum final
2.963×10-2

This is a solar moth type ship designed by Krafft Ehricke.

KRAFFT EHRICKE'S PLASTIC SHIP

It would consist of a huge bubble of transparent polyester plastic. The bubble could be some 300 feet (90 m) in diameter with a skin only a thousandth of an inch (0.0254 mm) thick. It would be slightly ressurized to give it a spherical shape. Half the inside surface would he silvered to create a hemispherical mirror that would concentrate the sun's rays on a heating element. In this element the hydrogen would be vaporized.

Piped to directable nozzles, one at each side of the sphere, the gas would provide thrust for acceleration, braking and maneuvering. The crew's gondola and associated equipment including solar battery for auxiliary power would he supported by a framework in the center of the big sphere.

It should he remembered that a space ship uses power only during its initial acceleration. The vehicle coasts the rest of the trip. Nevertheless it should carry large reserves of propellant.

Here the solar drive has real advantage. Its heat-collecting device, the hemispherical mirror, weighs possibly 1000 pounds (450 kg) as compared to a much greater weight of oxidizer that would need to be carried in a comparable chemical rocket. This saving in weight permits additional hydrogen to be carried.

Solar drive provides low thrust as compared to the very high thrust of a chemical rocket. This is a good thing, for the fragile plastic bubble will tolerate only low accelerations. It will be necessary to remain under power for hours to achieve the acceleration obtained in minutes by a chemical power plant.

From POPULAR MECHANICS March 1957
THE SOLAR-POWERED SPACE SHIP
TABLE 4
Characteristic Data of the Solar Powered Spaceship Prototype
1. Weights
Radiation Collectors (2)1,000 lb.
     Polyester Spheres (2)740 lbs.
     Cold Tubes, Wires, Sprints190
     Connections, Reflector Drive, Misc.70
Hot Piping System (for 2 Collectors)700
     Heating Elements200 lbs.
     Hot Pining350
     Radiation Jacket for Hot Pipes150
Engine500
     Turbine and Alternator60 lbs.
     Low-Pressure Booster Pump
(0.37 lb/sec H2)
10
     Hi-Pressure Pump & Motor
(2 sets)
50
     Wiring
(2 sets, electrical)
20
     Motor50
     Connections & Shut-off Valves10
     Array of Solid Propellant
Starter Rockets
300
Liquid Hydrogen Tank
(17 ft dia., 900 ft2 surface)
800
Controls, Crew & Equipment2,400
DRY WEIGHT5,600 lb
Liquid Hydrogen11,000
GROSS WEIGHT16,400 lb
2. Miscellaneous Data
Collector
Sphere
Diameter12 ft
Intercepted Area12,870 ft2
Circumference402.1 ft
Surface Area51,46808 ft2
Volume1,098,000 ft3
Hydrogen gas
wt. in sphere
(0.01 psi)
(helium about
twice)
300°F2.74 lb
-150°F6.85 lb
Intercepted area theoretically required to
produce 80 lb of thrust at Isp=450 sec
9,300 ft2
Reflector efficiency 9,300/12,8700.725
Theoretically produced thrust per reflector111 lb
Energy theoretically collected by reflector1,287 kw
Theoretical specific energy consumption12.9 kw/lbf
Actual thrust assumed to be produced per collector80 lbf
Total thrust produced160 lbf
Thrust to weight
ratio
initial0.976×10-2
maximum final2.963×10-2
3. Ideal Performance
Loading Factor (H2/Gross Wt)0.67
Maximum Possible Mass Ratio
Based on Loading Factor
3.03
Operational Specific Impulse Assumed450 s
Delta-V Based on Above Data15,730 ft/sec
From THE SOLAR-POWERED SPACE SHIP by Krafft Ehricke (1959)

Convair Urania

This is from Handbook of Astronautical Engineering, edited by Heinz Hermann Koelle, 1961. It describes Urania, a 1960 Convair/General Dynamics design for a manned hybrid chemical/ion-drive exploration spacecraft.

It is basically a modified Electra. With Electra, the ion-drive second stage is boosted into orbit by a Helios C chemical stage. Once in orbit the Helios is discarded.

With Urania on the other hand, the Helios C boosts the Urania into orbit, coasts for a while, and boosts the Urania into a parabolic departure orbit. Only then is the Helios discarded. In order to accomplish this, the Electra wet mass of 127,000 kg had to be reduced to only 81,700 kg wet mass for the Urania.

Using the high-thrust Helios chemical stage to kick the Urania into parabolic departure has several advantages.

First off, since the ion engine is not responsible for achieving parabolic departure, the Urania design can get away with using a lower-thrust ion drive than is needed for the Electra. This is because it won't have to fight the intense gravitational pull near Terra, made worse by the fact that at this stage of the mission it will be lugging a full load of ion drive propellant. Instead the ion drive will be operating in the solar-g field at Terra's distance, which is a modest 6 × 10-4 g (600μg). The lower-thrust Urania drive will have a lower mass electrical power system and lower mass radiator array than the Electra. The radiator will also have a smaller surface area and be easier to shield from meteoric damage.

Secondly, the Urania ion drive has an initial acceleration of 60 μg and would to take eight months to boost into a parabolic departure. The Helios chemical stage can do the boost in a few minutes. This drastically reduces the time the ship spends in the deadly Van Allen radiation belt. Helios also reduced the total trip time, since obviously parabolic departure time counts towards the total.

Thirdly it allows better propulsion optimization. Remember that with ion drives the higher the thrust, the lower the specific impulse. This means that a spacecraft design can have more payload if the engine has a higher specific impulse, but at the price of reducing thrust and thus reducing acceleration and thus increasing the mission duration. However with Urania, the parabolic departure is handled by the Helios chemical stage, not the ion drive. So there is no mission duration penalty for increasing the ion drive specific impulse and increasing the payload. The Urania's payload weight for a Venus mission is 42,180 kg. The Electra's payload weight for a Venus mission is only 39,462 kg.

URANIA
First StageHelios C
(chemical)
Second StageUrania
(ion drive)
Parabolic delivery weight
(initial weight stage 2)
81,700 kg
Target planetVenus
Initial accel
(stage 2)
60 μg
(0.0006 m/s)
Total powered flight time
(stage 2)
1.32 years
Thrust49 N
Isp13,000 sec
Electric power output3.5 MWe
Electric power equipment8,160 kg
Radiator surface area836 m2
Radiator projected area269 m2
Radiator arrangementMulti-fin
Num fins in reflectorsx9
Length of radiator15.9 m
Weight of all radiator fins8,390 kg
Total reflector surface area688 m2
Total reflector weight2,268 kg
Ion drive4,536 kg
Propellant16,100 kg
Payload weight42,180 kg

The pictured multfin radiator has an inner diameter of 6 meters. During life-off, into this hollow can be stored part of the life-support system and the centrifuge, unlike the Electra whose radiator have no such hollow. Other than that, the telescoping boom, life-support system, and the centifuge of the two designs are much the same.

Douglas Mars

This is from Proceeding of the Symposium on Manned Planetary Missions 1963/1964 Status, part 10 "Manned Mars Exploration in the Unfavorable (1975-1985) Time Period". It is a report on a Mars mission study by the Douglas Aircraft Company.

The section about the spacecraft is interesting because they examine about 15 different options, and score them according to a variety of criteria. They went with option 5.

The missions was to be 460 days duration wih a 20-day Mars capture-orbit stay time. The unsurprising recommendations were to restrict crew selection to 20-percentile men (sexist!), have the crew cabins as close as possible to the drive-the-astronauts-to-psychotic-break mimimum size limit (31.15 m3 per crewperson), combine meteoroid and insulation with the load-carrying structure (oh, like any spacecraft design doesn't do that?), a crew of six, use fiberglas tanks, and gas core nuclear thermal rockets would be real nice if they could be man-rated (in your dreams...).

The spacecraft would have a wet mass of 979,000 kilograms, and a dry mass of 278,000 kilograms. It would have four stages, not counting the ROMBUS reusable chemical booster that lofts it into LEO. A separate ROMBUS flight lofts the propellant. After each burn the current stage is discarded along with the still-hot nuclear engines. This means the spacecraft does not have to carry along extra propellant to cool down the engines.

The report is a little vague on performance. If this was a single-stage rocket it would have a delta-V of about 11,000 m/s. Since it is a staged rocket it presumably has more than that.

The habitat module with consumables for the crew of six is 35,320 kg, which is the mass of the payload package less the mass of the Mars Excursion Module and the Earth Entry Module. The payload is packed around the fourth stage. The artificial gravity centrifuge is an enclosed ring containing two cable-driven carts riding on the inner surface of the cylindrical rails.

Stages one and two have 250K Phoebus nuclear engines, stages three and four have 30K metallic core nuclear engines (as opposed to graphite core). Each Phoebus engine has a thrust of 1.11×106N, each metallic engine has a thrust of 1.33×105N. Both have a specific impulse of 850 seconds (8,340 m/s exhaust velocity). It would be nice to use the metallic engines on the lower stages, but you'd need clusters of eight, and nuclear decoupling is a big challenge (neutrons from adjacent engines make the nuclear chain reaction in a given engine go out of control).

Engines
StageTypeNumber
1250K Phoebus2
2250K Phoebus1
330K Metallic2
430K Metallic1

After leaving Mars, when approaching Terra, the fourth stage nuclear engine will slow the vehicle down to 12.2 km/s relative to Terra. The remaining velocity will be eliminated by aerobraking with the astronauts inside the Earth Entry Module. The rest of stage four will go sailing off into the wild black yonder.

Dusty Plasma Fission Fragment

Dusty Plasma Fission Fragment
Engine
TypeDusty Plasma FF
Exhaust Velocity150,000,000 m/s
(0.05 c)
Isp15,000,000 sec
Mass Schedule
LiH Moderator6,000 kg
Heat Radiator2,000 kg
Magnets,
power recovery,
coils,
structure
1,000 kg
INERT MASS9,000 kg
Payload1,000 kg
DRY MASS10,000 kg
10 year / 500 AU mission
Fuel180 kg
WET MASS10,180 kg
Mass Ratio1.018
ΔV6×106 m/s
Power Level350 MW
30 year / 0.5 LY mission
Fuel? kg
Power Level5.6 GW
50 year / 4 LY mission
Fuel240,000 kg
WET MASS250,000 kg
Mass Ratio25.0
ΔV4.8×108 m/s
Power Level208 GW

This is from Dusty Plasma Based Fission Fragment Nuclear Reactor (2005). The basic design is for an uncrewed space probe, but it could be scaled up. This uses the basic Dusty Plasma Fission Fragment engine. Performance can be optimized by swapping this out for an Afterburner DPFF engine.

The report estimates a couple of mission delta-Vs assuming a one-way trip using a Brachistochrone trajectory (accelerates for half the distance, decelerates for the other half for a total specified time duration).

I tried to check their math, but my results were off by an order of magnitude so I must be doing something wrong. The report states that a ten-year mission (3.1×108 sec) to the gravitational lens point 550 AU (8.2×1012 m) from Sol would require a delta V of 2% the speed of light (6×106 m/s).

The exhaust velocity of the engine is 1.5×108 m/s, so the spacecraft's fuel mass fraction is 5%. For a dry mass of 10,000 kg this means 180 kg of nuclear fuel. They added the fuel fraction to the dry mass to get the wet mass. Then they multiplied the mass by the implied acceleration to get the thrust required. Assuming that 46% of the fission fragments provided thrust, this allowed the power level of the reactor could be calculated. They figure the power level is 350 megawatts. Which is good because they figured the basic rocket was only good up to a thermal limit of 1,000 megawatts.

A 30 year trip to the Oort cloud at 0.5 light-years would needd a 5,600 megawatt reactor (5.6 GW). And a 50 year trip to Alpha Centauri at 4 light-years would need a whopping 208 GW reactor and 240 tons of fissile fuel.

Endurance

The Endurance is a fictional spacecraft from the movie Interstellar (2014)

The ring has an outer radius of 32 meters and rotates at a rate of 5.6 rpm. This gives an artificial gravity of 1 g at the outer rim. The spin rate will require some training and several days of adaptation on the part of the crew to avoid spin nausea.

The primary engine is an array of twelve magnetoplasmadynamic (MPD) engines, divided into four main engine modules. Each engine module contains three engines, tokamak power plant which generates electricity via magnetically-confined fusion, a RCS cluster using standard hydrazine, and presumably the reaction mass since there doesn't seem to be any remass modules. Remass is usually helium, seeded with something like potassium. The tokamak power plants are needed since MPD engines are power hogs (figures I have say 4 gigawatts of electricity per engine). A bit of the electrical power is allocated for the rest of the shop systems.

Liquid helium is very non-dense, which leads me to think that it will be impossible to squeeze enough remass into the engine modules for the delta-V displayed in the movie. But MPD engines can use other elements as remass, lithium is sometimes used.

MPD engines typically have exhaust velocities from 30,000 to 120,000 meters per second (Isp 3,000 to 12,000 seconds) and thrust ranging from 11 to 20,000 Newtons per engine. The movie keeps stressing that these are "advanced" MPD engines while not getting too precise about how much thrust they actually produce.


In the movie the Endurance makes a run to Saturn then travels around in the Gargantua system. Terra-Saturn Hohmann is about 7,300 m/s delta-V. Add another 7,300 for returning to Terra, and a couple more to explore the Gargantua system and return. So as a SWAG the total delta-V would be on the ordr of 4 × 7,300 = 29,200 m/s of delta-V. If the MPDs produce 120,000 m/s exhaust velocity, the mass ratio would be a reasonable 1.28.

However, a Hohmann to Saturn takes about six years, while the movie has the trip taking only two. So they ain't doing no Hohmann, they are spending more delta-V to get there faster. I'm not sure but it is probably more like 28,000 m/s Terra-Saturn. Less whatever delta-V boost they got from their gravitational sling-shot around Mars. Worst-case is 112,000 m/s would mean an ugly mass ratio of 2.5 or so. But at this point we are piling flimsly assumptions upon more flimsly assumptions so who knows?


What is conspicuous by its absence is heat radiators, but almost all movies fail on that point. Sorry, no Heat Radiator Award for you.

Theoretical MPD engines require 4 GWe of electricity each and output 3.1 GW of thrust power. Which means you have to get rid of 0.9 GW of waste heat. Per engine. Total of 10.8 GWth. Plus the waste heat from the four tokamaks, whatever that comes to.

My slide rule says that if the radiator has an operating temperature of 1,600 K and an emissivity of 0.95, you'll need 30,600 square meters of radiator to get rid of 10.8 GWth (15,300 m2 if radiator is double sided). Equivalent to a square 124 meters on a side or about twice the diameter of the ring. Plus more to get rid of the tokamaks waste heat.

      So he looked away from the Earth and Brand, and focused his attention on the Endurance, as they approached her. His first impression was of a wedding ring, glittering in the twin lights of the Earth and the sun.

     The Rangers were sleek, winged, aerodynamic craft built for landing and taking off from planets that possessed atmospheres. Not so the Endurance—there was nothing aerodynamic about her, and any landing she made on any planet with an atmosphere would be pretty much the same sort of landing as a meteor would make: fast, fiery, and catastrophic.

     Yet floating in space—where she had been built—the vessel was a thing of beauty. She was, indeed, a ring—but only in the most basic sense, and as they drew nearer his original impression faded. He could distinguish that she was formed from a number of boxy, trapezoidal, prism-shaped modules jointed together by curved connectors. The “ring” wasn’t empty either. Access tubes led from the inner surface of the circular body to a central axis where the docking locks lay. Two ships—the landers—were already there. All she needed were the two Rangers. Feeling oddly calm, Cooper maneuvered his Ranger in, matching his velocity to that of the starship.

     As they boarded the Endurance, it became clear that it wasn’t as roomy as it looked from the outside. Part of this was because two-thirds of each of the modules was taken up by storage. The floors, the walls—almost every surface was composed of hatches of various sizes. On a deep-space vessel, there could be no wasted space—not even one the size of a matchbox.

     Flipping switches and adjusting settings, Amelia, Doyle, and Romilly began powering up what would be their home for—well, who knew how long? She watched Tars activate Case, an articulated machine like himself, who made up the final member of their crew.

     Doyle moved “up” to the cockpit and turned on the command console. Technically, there was no up or down at this point, but soon it would no longer be a technicality, as evinced by the ladder that led from the lower deck up to the command deck.

     She watched as Doyle finished linking the on-board systems to the Ranger.
     “Cooper, you should have control,” Doyle said.
     “Talking fine,” Cooper replied. “Ready to spin?”
     Doyle and Romilly strapped in. Amelia followed their lead and took a chair.
     “All set,” she replied.

     She felt nothing at first, but then the ship began to shake as Cooper fired the Ranger’s thrusters, angled perfectly to set the great wheel turning. As the spin picked up, weight began to return to Amelia’s body, pulling her feet toward the outer rim of the starship. It wasn’t gravity, exactly, but the manifestation of inertia often referred to as centrifugal force. Without it—without some semblance of weight— bad things happened to the human body over time, like bone loss and heart disease.

     We’re going to need our bones and our hearts when we reach our destination, she thought.

     Unfortunately, spin wasn’t a perfect substitute for gravity, because the inner ear wasn’t entirely fooled by it. It knew they were whirling around due to a little thing called the Coriolis effect.

     On Earth the Coriolis effect was a big deal. It drove the climate, creating huge cells of air moving in circles—clockwise in the northern hemisphere, counter-clockwise in the southern. But the Earth was so huge, the human body didn’t notice the spin on a personal level. Yet on a whirling carnival ride it was easy to feel, often with upsetting results.

     The Endurance lay somewhere in between those extremes, though leaning toward the carnival ride. Amelia felt it herself, especially when she moved toward the axis, but it didn’t really bother her.

     Romilly, on the other hand, already was looking a little green.
     “You okay, there?” she asked him.
     “Yup.” He practically gurgled as he replied. “Just need a little time—”
     “There should be a Dramamine in the hab pod,” she told him. He nodded gratefully, and moved gingerly in that direction.

     “Eight months to Mars,” Tars said, “just like the last time we talked about it. Then counter-orbital slingshot around—”

     “Things look good for your trajectory,” the professor continued. “We’re calculating two years to Saturn.”


The Endurance carries two Ranger exploration craft and two Lander cargo craft.


Rangers are single-stage-to-orbit (SSTO) reconnaissance spacecraft with enough delta-V for one trip from the surface into orbit and one trip from orbit to the surface. They can only manage this when the planet has an atmosphere since they rely upon air for propellant and aerobraking. This is not a problem since the mission is scouting for habitable planets.

High powered fans gulp air through intakes and and exhaust it through either rear jets for horizontal flight or through vents on the underbelly for vertical take-off and landing. The main propulsion system are twin linear aerospike hybrid plasma engines, which sound like utter technobabble to me. Chemical fueled aerospike engines have their thrust augmented by electromagnetic means, someway somehow. If the atmosphere contains oxygen the engines will cheerfully use it in order to conserve its oxidizer stores.

When landing, it aerobrakes away its orbital velocity, using a heat shield covering its belly much like the old NASA Space Shuttle.

Power is provided by twin miniaturized tokamak fusion reactors, a triple redundant fuel cell system, and high efficiency solar cells on its roof.

It also has four hibernation tanks that the crew can use for prolonged missions.


Landers are single-stage-to-orbit (SSTO) cargo spacecraft designed to deliver landing pods to a planet's surface in order to establish a colony.

In an odd design choice, the lander's heat shield is on what a layperson would call the roof. Let me explain.

The landers have a standard belly-lander configuration, it lands like an aircraft on its belly (not on its tail like a SpaceX Falcon) with the landing pod cargo slung underneath. This is a problem since the landing pod would be incinerated during aerobraking, and putting a heat shield on the landing pod would wastefully reduce its payload mass. So when the lander is aerobraking it flips over so the landing pod is uppermost and the lander's heat shielded back is lowermost. That way the heat shield takes the aerobrake heat and the landing pod stays cool and protected.

The lander is also heavily armored, just in case the surface enviroment proves to be hostile.

First Men to the Moon

This design is from a book called First Men to the Moon (1958) written by a certain Wernher von Braun, aka "The Father of Rocket Science" and the first director of NASA. The book came out shortly after the Sputnik Crisis.

Gary Johnson Mars Mission

COLONIZATION SHIP STUDY

I have gotten involved with some friends on the New Mars forums discussing what might be appropriate for very large colonization ships.  This kind of mission demands the delivery of very large payloads.  Doing this effectively requires a reusable ship.  That means you stage off (or jettison) nothing.

It is easy to run a rocket equation-based trade study that assumes a one-stage round trip, that jettisons nothing. Making it carry the same large payload on the return voyage simplifies the analysis, but very likely over-penalizes the design. But at this level of analysis, that really doesn’t matter.

This is basically just a bounding analysis for screening candidate propulsion approaches to a Mars colony ship design. I included nuclear explosion propulsion, nuclear thermal propulsion, ion propulsion, LOX-LH2 cryogenic chemical propulsion, and storable chemical propulsion.
Update 9-13-19:  there is more than one kind of nuclear thermal rocket.  I took a closer look at 6 different nucear thermal rocket approaches,  and in a more nuanced way,  in "A Closer Look At Nuclear Thermal"

Spreadsheet Inputs

The spreadsheet inputs are highlighted yellow. Payload delivered is common to all the designs, and actually arbitrary, but I thought 2000 metric tons might go a long way toward the beginning of a colony.

Inert fractions vary with the propulsion selection. I used data from Ref. 1 to set a realistic guess for the inert fraction, of the nuclear explosion drive. It is very high, reflecting the massive pusher plate, two-stage shock absorption system, and the armored hull.

The Hall effect ion drive is based on existing Busek satellite thrusters already in service, and modified to “burn” iodine, something plentiful, cheap, and storable at low pressure. Getting to an acceptable vehicle acceleration requires a very large thruster array and a nuclear power source in the multi-megawatt range. I just guessed the inert mass fraction that might cover this.

Because of the heavy reactor core and low engine thrust/weight achieved in the old NERVA nuclear thermal rocket development effort, I used twice the typical chemical stage inert fraction as a “good guess” for the nuclear thermal inert mass fraction. There is good data about this engine in Ref. 2.

Both the LOX-LH2 cryogenics chemical propulsion, and the NTO-MMH storable-propellant chemical propulsion, share the same “typical” stage inert mass fraction.

Delta-vees for the Mars trip are for departing and arriving in low Earth orbit to/from a min-energy Hohmann transfer ellipse, plus the corresponding delta-vees for arriving into and departing from low Mars orbit. The same applies to the Ceres transfer, except that the ship just matches Ceres orbital velocity about the sun instead of entering a “low orbit”. This would be typical of many small main belt asteroids.

For those types of propulsion in the order listed above (nuclear explosion, nuclear thermal, ion drive, LOX-LH2 chemical, and storable chemical), my assumed inputs for Isp were 10,000 sec, 1000 sec, 3000 sec, 470 sec, and 330 sec respectively. Vehicle inert mass fractions were 0.50, 0.25, 0.10, 0.05, and 0.05 respectively.

All these dV’s were summed, as required to do the entire mission single-stage. The total orbital delta-vee (dV) to and from Mars is 3.84+1.83+1.83+3.84 = 11.34 km/s. Impulsive-burn options need supply only that summed delta-vee with zero gravity and drag losses. Long-burn ion must supply a lot more than that, due to very large planetary and solar gravity losses.

All but the ion option were considered as "impulsive burn" and Hohmann min energy transfer, with vehicle acceleration exceeding 0.1 gee to enforce that. These used the unfactored sum of orbital dV's to and from Mars (orbit-to-orbit transport) as the mass ratio-effective dV for the rocket equation. The spreadsheet input is factor equal to one.

The ion option must spiral-out and spiral-in at the planetary orbits, and accelerates to midpoint then decelerates to arrival on the transfer trajectory (a patched spiral about the sun). Propulsion is sized for 0.001 gee to ensure that this kind of transfer is feasible. To account for the planetary and solar gravity losses of the resulting months-of-burn, I just doubled the orbital dV sum to 22.68 km/s. For the spreadsheet, this is factor equal to two.

For Ceres, Earth departure and arrival dV is 5.24 km/s. The orbit-matching dV at Ceres (arrival and departure) is just about 3.49 km/s. That round trip sum is 17.46 km/s for all but the ion drive option, unchanged by factor equal to one. Using factor equal to two for ion drive, that mass ratio-effective total is 34.92 km/s.

All 5 designs carried exactly the same 2000 metric tons of dead-head payload, an arbitrary selection perhaps appropriate for a colony-type mission. (I did not look at how to get that payload up to LEO, or down from LMO, that issue would be the same for all the candidates.) This was done for Mars in a spreadsheet worksheet, whose image is Figure 1. All figures are at the end of this article.

Analysis Equations

Sum the round trip delta-vees, and factor the sum for the mass ratio-effective delta-vee required of each propulsion type: required dV = (factor)(sum of all 4 orbital delta vees), where factor = 1 for impulsive propulsion (acceleration exceeding 0.10 gee), and factor = 2 for long-burn ion propulsion (0.001 gee required).

Estimate the effective exhaust velocity from the specific impulse: Vex, km/s = 9.8067 (Isp, s)/1000

Calculate the mass ratio required: MR = exp(dV/Vex), with both velocities in km/s

Calculate the propellant mass fraction: Wp/Wig = 1 – 1/MR

Input an inert mass fraction Win/Wig (must be justified in some way as “realistic”)

Calculate the available payload fraction Wpay/Wig = 1 – Win/Wig – Wp/Wig (must be positive to be even theoretically feasible)

Input the delivered dead-head payload Wpay, metric tons (arbitrary, but should be realistic)

Calculate the ignition mass Wig, metric tons: Wig = Wpay/(Wpay/Wg)

Calculate the inert mass Win, metric tons: Win = Wig*(Win/Wig)

Calculate the propellant mass Wp, metric tons: Wp = Wig*(Wp/Wig)

Calculate the ignition to payload mass ratio: Wig/Wpay = (Wig, m.ton)/(Wpay, m.ton)

Results Obtained

Results for Mars: nuclear explosion drive 5118 metric tons at ignition with ignition/payload 2.56:1 (see Figure 2). Nuclear thermal 30,945 metric tons at ignition with 15.47:1 ignition/payload (see Figure 3). Hall effect ion drive 5516 metric tons at ignition with ignition/payload 2.76 (see Figure 4). LOX-LH2 56,486 metric tons at ignition with ignition/payload 28.24 (see Figure 5). Storable chemical utterly infeasible with a negative payload fraction available (see Figure 6).

The nuclear explosion drive offers the lowest ignition/payload ratio going to Mars at 2.56:1, based on the old 1950's shaped-charge fission device technology. This would be a very tough ship design, probably usable for a century or more, and likely tough enough to aerobrake, reducing the load of bombs in favor of more payload. Its stout hull and huge pusher plate are effective radiation shields.

The ion propulsion offers the next best ignition/payload ratio going to Mars at a very comparable 2.76:1, which to be practical would require its thrusters operating on something cheap, plentiful, and storable-as-a-condensed-phase (at very low pressure), like iodine. This would be a relatively gossamer structure unable to survive aerobraking, and it would likely also have a limited service life. Radiation protection would have to be added.

Two of the others (nuclear thermal and LOX-LH2), while theoretically feasible, are nowhere close in ignition/payload ratio going to Mars. These are unaffordable “Battlestar Galacticas” for any reasonable payload delivery aimed at colonization. And the storable chemicals are just infeasible in any sense of the word for a Mars colonization ship, simply because there is a negative payload fraction available, once propellant fraction has been determined, and with a suitable inert fraction input. It simply cannot do the mission single stage.

I think you can look at the ignition/payload mass ratio to judge whether-or-not a given propulsion system might serve as a practical way to build a colony ship. This value needs to be no more than about 5 or thereabouts, in order not to build an unaffordable “Battlestar Galactica”. This is a “fuzzy” boundary, dependent upon how much you think you can afford.

The same sort of analysis applies to other destinations. You just need an appropriate list of orbit-to-orbit delta-vees, and the same list of realistic guesses for inert fractions.

Results for Ceres: I added a worksheet to the same spreadsheet for a colony-type ship to Ceres, as “typical” of the asteroid belt. Those spreadsheet results are shown in Figure 7. Figures 26 also show the Ceres results (as well as the Mars results).

The only feasible choices for Ceres colony ships were nuclear explosion propulsion and nuclear-powered electric propulsion. It’s the same basic calculation, just with somewhat bigger delta-vees. The nuclear thermal and both chemical options simply had fundamentally-infeasible negative payload fractions available. They simply cannot perform the mission single-stage.

The same general outcome choices obtain for Ceres as for Mars: your nuclear explosion drive ship is quite robust, promising a long service life, while the ion ship is rather flimsy. For this main belt asteroid application, the ignition to payload ratio is also substantially more favorable for the nuclear explosion ship (2.97), vs the ion ship (4.87).

Conclusions

The trend here is clear: the further out you go with a single-stage, round-trip colony ship, the more the ignition/payload ratio is going to favor nuclear explosion propulsion as the more affordable option. Radiation protection needs will also favor the shielding effect of the stout hull required of the nuclear explosion drive. Bigger also favors ease of incorporating spin “gravity”.

References

#1. George Dyson, “Project Orion – The True Story of the Atomic Spaceship”, Henry Holt, 2002.

#2. David Buden, “Nuclear Thermal Propulsion Systems”, Polaris Books, 2011.

A CLOSER LOOK AT NUCLEAR THERMAL

This article takes a closer, more nuanced look at nuclear thermal propulsion for large colonization ships. It still assumes fairly large dead-head payloads, but only carried on the outbound voyage! Propellant is sized to make the outbound and return voyages in one stage (no stage-off or jettisoning of anything along either way, just unload of the dead-head payload at destination). The journey baseline is low Earth orbit to low Mars orbit, and back.

How the ships or the payload get to low Earth orbit is unaddressed. How the payload gets delivered to Mars’s surface from low Mars orbit is unaddressed. How the ships are refueled and reloaded in low Earth orbit is unaddressed. What is addressed here, that is unlike the earlier simpler study, are the separate inert weights associated with the payload section, the propellant tankage section, and the engine-with-its-associated-subsystems. The minimum vehicle acceleration requirement is increased to 0.5 gee, except for one system deemed adequate at 0.33 gee.

The previous closely-related article was “Colonization Ship Study”, dated 9-9-19. It examined the simpler-to-analyze case of carrying the dead-head payload both ways (outbound and return), so that there was one mass ratio and one delta-vee (dV) to cover the round trip. That scope was multiple fundamentally-different forms of propulsion: nuclear explosion drive (or “pulse propulsion”) as it was envisioned in the late 1950’s, nuclear thermal propulsion (as a version of the solid core NERVA for which engine prototypes were tested), Hall effect ion propulsion based off of plentiful, cheap, and solid-phase-but-sublimable iodine, LOX-LH2 chemical rockets, and storable-propellant rockets.

Scope here is only nuclear thermal rocket propulsion, but with the highly-variable tested or envisioned characteristics of six different design approaches. It is these six that are compared in terms of the ratio of initial ignition to dead-head payload weight, using the same maximum-attractive criterion of 5 as in the earlier study. These six approaches and their relative states of technological readiness are:

  1. as-tested NERVA solid core
  2. the best-anticipated solid-core NERVA derivatives that never got built or tested
  3. the particle bed solid core reactor engine (one version of which was “Timberwind”, which got some exploratory testing revealing unresolved problems, but never reached the engine prototype stage)
  4. the so-called “nuclear light bulb” gas core concept (some insufficient feasibility tests)
  5. the open-cycle gas core concept restricted to regenerative cooling, meaning no radiator required (some insufficient feasibility testing)
  6. the open-cycle gas core concept with a large, heavy external waste heat radiator (some insufficient feasibility testing)

To accomplish this investigation, I added an additional worksheet to the colony ships.xlsx spreadsheet file that I used for the earlier study. Unlike the previous study, there are no closed-form ways to get from dead-head payload to a vehicle weight statement. The calculation uses iterative convergence of the propellant tank inert weight, and iterative convergence of engine thrust sizing in terms of the resulting vehicle acceleration gee capability.

For this investigation, the payload section is presumed to be some sort of enclosed hull, with adequate insulation, radiation shielding, and micrometeor protection for a crew built into it, in some unspecified way. The ratio of dead-head payload mass (contained inside) to the loaded payload section mass is a fraction denoted as fpay. The dead-head payload size drives everything in the end, as all results are directly proportional to the dead-head payload input. For this investigation, dead-head payload was arbitrarily set at 100 metric tons, and fpay = 0.8, the same for all six engine types. Thus:

Loaded payload section mass = dead-head payload mass/fpay

Payload section inert mass = loaded payload section mass – dead-head payload mass

The propellant tank section contains the common propellant for all nuclear thermal engine approaches: liquid hydrogen (LH2). This is a harsh cryogen, requiring solar heating control, significant insulation, and some sort of cryocooler to control evaporation. This is simply going to be heavier than the lightest-possible single-wall bare tank. The ratio of propellant mass to loaded tank mass is the fraction ftank. The single value ftank = 0.95 was used for all six engine types. Thus:

Loaded tank section mass = propellant mass/ftank

Tank inert mass = loaded tank mass – propellant mass

Propellant mass must be the sum for two burns at differing dead-head payload. One starts with a guess for tank inert, and iteratively converges it to the result

The engine “section” is the nuclear thermal rocket engine (or engines, for redundancy), complete with turbopumps and control equipment, a radiation shadow shield for the crew up forward, plus any waste heat radiator that may be required (if regenerative cooling alone cannot do the job). This radiator (if present) and the core-plus-engine hardware lead to a characteristic engine thrust/weight ratio T/We, which is dimensionless under the definition that both thrust T and engine weight We (on Earth) are measured in force units. This ratio is different for each engine type, as is the resulting specific impulse. The values I used follow:

TypeIsp, sT/Wedevelopment status
NERVA7253.6as-tested in engine prototypes
derNERVA10005derivative-of-NERVA, estimated on paper
PBR10007particle-bed reactor, based on “Timberwind”
Nuc.lt.blb130010“nuclear light bulb” gas core concept, some feasibility
Open GCR250020open-cycle gas core concept limited to regenerative cooling
GCR+rad60000.5open-cycle GCR with heavy waste heat radiator, concept

For this kind of data, the main results used to size the vehicle are the exhaust velocity Vex (km/s), and the engine system inert mass (metric tons). These are:

Vex, km/s = (Isp, s)*9.8067/1000

Engine system inert mass, metric tons = thrust level, KN/(9.8067 * T/We)

For the remaining vehicle characteristics, all the concepts except “GCR+rad” were required to size thrust level such that the vehicle acceleration at the initial ignition mass was at or just above 0.5 gee. This corresponds to about a 15 minute Earth departure burn, definitely short enough to qualify as “impulse”, and not have the orbital dV be factored-up for gravity loss to be mass ratio-effective.

With the data I used, the GCR+rad system could not reach half a gee, but converged fairly well at 0.33 gee. This is less than a 30 minute burn, still short enough to consider as “impulsive” for Earth departure.

Max gee at final burnout weight upon Earth return should be under about 5, but this proved not to be a problem.

It’s a two-level iteration: first set a thrust level, then converge your guess for propellant tank inert weight with the final result of the calculation for tank inert weight. Then check and adjust your thrust level for the right Earth departure gee level. Then converge the tank inert weight again. Repeat the process as needed to get however-close a convergence you deem tolerable (0.1-0.01 ton range).

The orbital dV’s that are required are those for getting from low orbit onto a min-energy Hohmann transfer ellipse. The values used are worst-cases that do not go together; the difference is a nice little “kitty” to cover midcourse corrections. Earth departure = Earth arrival = 3.84 km/s. Mars arrival = Mars departure = 1.83 km/s. These sum to 5.67 km/s outbound in a heavier ship carrying payload, and 5.67 km/s return in a lighter ship with no payload and already having burned off some propellant on the outbound voyage.

Factored for losses, these dV figures become the mass ratio-effective dV’s for design purposes. Those and the Vex for each engine type give you the mass ratio MR for each engine type, one for outbound, the other for return.

MR = exp(sum dV/Vex) with both velocities in km/s, and the sum dV for outbound or return

You start the calculation with the return voyage by summing up the inerts (payload section inert + tank inert + engine inert), plus zero dead-head payload, as the burnout mass at Earth arrival. This starts with a best guess for inert tank mass, as well as for installed engine thrust level. Apply the appropriate mass ratio to get Mars departure ignition mass. The difference in ignition vs burnout mass is the propellant expended for the two burns of the return voyage.

The next step is the outbound voyage. The Mars departure ignition mass, plus the dead-head payload mass, is the Mars arrival burnout mass. Apply the appropriate mass ratio to get the Earth departure ignition mass. The difference in ignition vs burnout mass is the propellant expended for the two burns of the outbound voyage to Mars.

The sum of the two propellant quantities is the total propellant for the round trip. Divide this total propellant by ftank to find the total loaded tank mass. The difference between loaded total tank mass and total propellant mass is the inert tank mass. This resulting inert tank mass is what your guess for tank inert mass must converge to! The best next guess is close to the last result.

Thrust divided by Earth weight is the vehicle acceleration gee estimate. This is done at each of the 4 vehicle masses: Earth departure ignition, Mars arrival burnout, Mars departure ignition, and Earth arrival burnout. Two of these are of real interest: Earth departure ignition (min gees), and Earth arrival burnout (max gees). The other two conditions fall in-between. You must adjust your installed thrust level to achieve min gees. Then iterate to convergence again on tank inert mass.

Max gees at Earth arrival burnout did not prove to be a problem, but should fall under 5 gees for the most tolerable results. Be sure you check for that outcome.

The last calculation sets up weight statements and estimated dV performance for the six propulsion types, using the data already calculated. The initial part of the weight statement is the vehicle buildup from payload and inert items to Earth departure ignition mass. Subtracting the total outbound propellant gives the Mars arrival burnout mass. Their ratio produces an outbound summed dV for both burns, to be calculated for each type (for comparison to the initial summed requirement).

That Mars arrival burnout mass, less the dead-head payload, is the Mars departure ignition mass. Subtracting the return voyage propellant produces the Earth arrival burnout mass. Their ratio produces a return summed dV for the two burns, done for for each propulsion type (for comparison to that summed requirement).

The deviations of these weight statement dV’s from the required values reflect just how closely you converged your tank inert weights. These should be only trivially off (by under 0.001 km/s = 1 m/s). If you see bigger errors, you didn’t converge your tank inert masses closely enough. The effect of being “off” on min gee (as set by installed thrust level) is small, when compared to the effect of being “off” on guessed tank inert mass.

At the very bottom of the weight statements are the vehicle payload fractions, in both definitions. One is the conventional definition: dead-head payload mass / Earth departure ignition mass. You probably should not consider anything under 0.2 for a practical colonization ship design. Its inverse is Earth departure ignition mass / dead-head payload mass. In that definition, you probably should not consider anything over 5 for a practical colonization ship design.

This limit (in either form) is inherently a very fuzzy judgement call. But, if dead-head payload mass is too small compared to Earth departure ignition mass, the resulting design will be inherently very expensive to build and to operate, just like with ocean-going transport when the cargo mass is small compared to the tonnage of the ship.

What I got for this study is given in Figures 1 and 2, a two-part image of the completed spreadsheet worksheet page. Of the six propulsion types, four look reasonably-to-very attractive. These are the derivative of NERVA, some form of PBR, and the two gas core concepts that do not require a huge waste heat radiator. The as-tested NERVA falls short because its engine thrust/weight is too low and the resulting large engine inerts drive the vehicle inerts, constrained by the large thrust level to achieve min acceleration gees. The gas core with radiator falls short because of the gigantic, heavy radiator.

Near-term, the higher Isp and engine thrust/weight of the derivative NERVA could be realized in a few short years, to an engine prototype ready for flight test. The PBR concept would take a few more years than that, since no prototype engines were ever ground tested, and some fundamental problems identified in testing of “Timberwind” components remain unresolved. The gas core concepts would require several-to-many years to reach a flight-testable prototype, since only very sparse lab-type feasibility demonstrations were ever done; plus, there is no guarantee of eventual success, either.

My own recommendation would be to base an initial design around the derivative NERVA as lowest-risk option of acceptable benefit, and plan on replacing it later with one of the non-radiator gas core designs, should that development prove successful.

Figure 3 sketches a ship design concept based on the derivative of NERVA, figured at 100 metric tons of dead-head payload delivered to Mars. Volume of LH2 and a guess for tank L/D set the tank dimensions. Everything else scales one way or another from that, as a first approximation. Everything about the weight statement and thrust level sizing is proportional to dead-head payload size. Dimensions would scale as the cube root of mass, provided that L/D ratios are preserved.

This vehicle rough-out delivers the same design dead-head payload to Mars as the proposed Spacex “Starship” design. The differences are several: this vehicle never lands on Mars (delivery to the surface is by unspecified other means), this vehicle must make a full Mars arrival burn into low orbit (“Starship” only makes a final touchdown burn after an aerobraking direct entry), and this vehicle returns all the way to low Earth orbit for reuse, unrefueled. It never needs to survive any sort of atmospheric entry.

This design makes the round trip single-stage unrefueled. The Spacex “Starship” is entirely one-way only, unless and until it can be refueled on the surface of Mars from local resources.

There is enough payload section volume to support a crew of up to 15, at about 300 cubic meters per person, in addition to the volume occupied by the dead-head payload, at a payload specific gravity averaging only 0.3.

This result says a Mars colonization ship able to carry 100 metric tons of dead-head payload one-way to Mars, and return to Earth with no payload, all one-stage, is not that large an item. It is not large enough to spin for artificial gravity like a rifle bullet, but it is large enough to spin end-over-end (like a baton) for artificial gravity. At about 3.24 rpm, there is about one full gee available in the payload section. That spin rate is tolerable to untrained, unacclimatized people, for long-term exposure.

The insulation and meteor shielding is about a meter thick on the payload section, meaning it can double as radiation protection. If those layers of fabric average 0.20 effective bulk specific gravity, that is some 20 g/sq.cm shielding mass, adequate for solar flare events, and offering some reduction of galactic cosmic radiation. The insulation and tank shell thickness of the propellant tank section was assumed to be 0.1 m. Engine section length was just a guess.

Key to this design as-sized is carriage of dead-head payload to Mars, but not from Mars. The return dead-head payload must be zero! If not, the propellant tank section must be significantly larger, to the detriment of the payload fraction criteria. Any crew and their life support must come out of that dead-head payload allowance (meaning near-zero crew on the return voyage).

These results look more favorable than the otherwise-comparable nuclear thermal option in the earlier study. That is precisely because dead-head payload is only carried one-way in this study, and it was carried both ways in the earlier study. That is one huge effect. But the trend from the earlier study applies here as well: if we design for a farther destination than Mars, the design won’t look so attractive in terms of the payload fraction criteria.

The restriction of zero dead-head payload on the return voyage is not as constraining as it first sounds, when one considers the goal is building a colony with these payloads. During that process there is little-or-nothing to ship home to Earth, except information, which is better sent electronically. Later, when an operating economy results in two-way trade, one will need commerce ships, not colonization ships. But, by the time that need arises, significantly-better propulsion technology should have become available.

Brief Result Summary: The best near-term option of the six nuclear thermal approaches, for a Mars colonization ship design, is the derivative-NERVA nuclear thermal propulsion approach (Isp ~ 1000 s and engine T/W ~ 5). For 100 metric tons dead head payload, the initial ignition mass is about 500 metric tons. That means for 1000 metric tons dead-head payload, the sized ship will initially mass about 5000 metric tons. For 2000 tons payload, the ship will be around 10,000 tons, etc. This is restricted to orbit-to-orbit operation, and to no dead-head payload on the return voyage. Even the small 100-ton payload size is large enough to spin end-over-end for artificial gravity at near 1 gee and an easily-tolerated spin rate. The payload section insulated design (if a meter of fabric layers) also inherently provides a fair amount of radiation protection.

MARS MISSION OUTLINE 2019

This year has been the 50th anniversary of the first man on the moon. That was the culmination of the space race between the US and Soviet Russia. That accomplishment was a whole lot more about “flags and footprints” and experimental flight test, than it was about science or real exploration.

This article builds upon some earlier articles posted upon this site.It presents the latest version of my Mars mission outline plan,with an enlarged manned transport,and the latest sizing of 1-stage 2-way reusable chemical landers.These earlier articles are as follows:

Why We Should Go Back (And Farther Still)

Is there anything worthwhile to accomplish out there?Yes,definitely!

In the longer term,there are those future off-world settlements and the associated future economies.I cannot tell you the details of how this might benefit us,because it has yet to be done.But it has always proven beneficial in prior centuries here on Earth.

In the shorter term,there are the possibilities of space resource businesses,and of planetary protection against rogue asteroid and comet impacts.That second item is the most important of all:there is simply no better reason for continuing both unmanned and manned space programs than finding ways to protect the folks back home!

It’s not about winning some race,and it’s not so very much about doing pure science just for the sake of knowledge.It’s about real exploration of the unknown,something hard-wired into humans.In centuries past,this was exploration of the unknown parts of the Earth.Now it is about space and the deep ocean floor.This article is concerned with the real exploration of space.

“Exploration” is a really an emotionally-loaded code word,something most people do not think about.What it truly means is you go there to find out “what all is there” (resources,including those you don’t at first recognize),and “where exactly it is” (how hard to obtain,as well as how much is there).Then you have to stay a while to figure out how to use what you found, in order to cope with living in the local environment. All of that is part of “real exploration”.

Unless you do that correctly,there is no real possibility of future settlements and the associated future economies,or any of the benefits that would ultimately derive therefrom!There is no way to accomplish much of anything else,except just the “flags and footprints” act of going there and returning (which is the bulk of what Apollo itself really accomplished at the moon).

Those who “get there first” do tend to do a little better in the long run,in terms of those benefits,provided that they do it “right” when they go.That is one crucial lesson from history.

My Suggestions for the Near Term

Establish a continuous human presence on the moon,the first item.Start small and expand it slowly over time. Do the lunar “exploration” thing right,this time.

Send humans to Mars as the fulfillment of a dream centuries old,probably the second item.When we go,do the “exploration” thing right,from the very first landing.Further,it starts long before the first item (going to the moon) is “done” in any sense of that word.

But,any vehicle capable of taking crews to Mars can also take a crew to near-Earth asteroids and comets.Visit those asteroids and comets and properly explore them,in order to learn how to defend against their impacting Earth,as well as “ground truth” for how to really do space mining.

That’s the third item,but it is just as easily done,and at least as important,as going to Mars. 

Maybe we do them at pretty much the same time.

Ethically and Responsibly Addressing Known Risks For Spaceflight

We are ethically-bound to address the known risks of manned spaceflight as best we can.There is a whole long list of safety risks associated with any sort of manned spaceflight.Three come to mind as the most truly credible risks:(1) reliability of,and escape from,spacecraft and booster rockets,(2) microgravity diseases,and (3) exposure to space radiation.

The first one has cost us three American crews totaling 17 people dead (Apollo 1,shuttle Challenger,and shuttle Columbia). Each caused a year-or-more stand-down,and very expensive investigations,plus very expensive changes.

The two shuttle losses were ultimately caused by bad management decisions valuing cost or schedule above safety.Apollo 1 was about a really-poor basic management attitude (“good enough for government work”) combined with technical ignorance,because we had never done this sort of thing before.

Those outcomes and their actual causes are why I claim “there is nothing as expensive as a dead crew,especially one dead from a bad management decision”.Bear in mind that those expenses are both economic and political (which includes public opinion as well as DC politics).

Making spaceflight more safe, from a reliability and escape standpoint, is now also something we already know how to address!This takes careful design allowing for failure modes,redundant systems, and copious verification testing.Mitigation efforts will never be perfect,but they can be quite good. Ethics requires that you treat this as a required constraint upon your designs.

It means you always provide “a way out” for your crew at every step of the mission.It really is as simple (and as hard to do) as that!This very seriously constrains your overall mission architecture,as well as your detailed space vehicle designs.

The other two have been long studied in low Earth orbit,where microgravity exposure is inherent in everything we have done there,and radiation exposure is somewhat more than on Earth’s surface,but less than outside the Van Allen radiation belts (and far less than inside the belts themselves).

Microgravity Diseases

Microgravity has proven to affect the human body in a variety of expected, and unexpected, ways.The longer one is exposed,the worse the various diseases become.Beyond the bone decalcification and muscle-weakening that we have long expected,there are also degradations of the heart and circulatory system,degradation of vision from eye geometry changes due to the fluid pressure redistribution,immune system degradations that we have yet to understand,and most recently genetic changes whose meanings are still a total mystery. No doubt more will be discovered,as that has been the trend.

The longer exposed,the longer it takes to recover upon returning home,with full recovery actually still in doubt for some of the effects,despite diet,drugs,and exercise. The practical time limit seems to be only a bit more than a year.For that very reason,usual practices on the International Space Station (ISS) call for 6 months to a year’s exposure at most,with 6 months the preferred limit.

We do know that something near one full Earth gravity (one “gee”) is therapeutic,precisely because that is what we evolved in.So,until we know better,any artificial spin gravity schemes need to supply very near one gee,in order to obtain the full Earthly benefits that we already know will work.

Destinations outside of Earth-moon space are very much further away than the moon:one-way travel times range from near 6 months to multiple years.This is pretty much outside the preferred limit of microgravity exposure that we have already established on ISS.

Mars is 6-to-9 months away one-way,and we do not know how therapeutic its lower gravity (38%) really is for the rigors of the return voyage.Other destinations are further away still,and all those we can actually land upon, are even lower gravity than Mars.That situation says quite clearly that we need to provide artificial gravity (no matter how inconvenient that might otherwise be !!!!) at something near one gee (until we actually know better !!) during these one-way transits to-and-from,in order to best preserve the health of the crews.

Ethically,you simply cannot argue with that conclusion,no matter how inconvenient for design purposes,or for total mission cost purposes.That is the only “box on thinking” applied here.

Supplying Artificial Gravity

There is as yet no such thing as “Star Trek”-type artificial gravity.The only physics we have to serve that purpose is “centrifugal force”.You must spin the vehicle,to generate “centrifugal force” as an equivalent to gravity.If the spin rate is low,then Coriolis forces (something everyone has experienced on a merry-go-round) become less important,and so fewer folks can tell the difference between this and real gravity,and there are fewer problems with disrupting the balance organs in the ear.

The physics of spin say that the acceleration you feel is proportional to the radius of spin and to the square of the spin rate.The actual physics equation says

a = R w2

where a is the acceleration,R the spin radius,and w the spin rate

Another form expressed in gees,and not absolute acceleration units, is

gees = 1.00 * [(R, m) / (55.89 m)] [(N, rpm) / (4 rpm)]2

Earthly experience with spin rates says that normal untrained and unacclimatized people can tolerate 3 to 4 rpm immediately,and for long-term exposures,without getting motion sick.People extensively trained might (or might not) tolerate higher spin rates in the 8-12 rpm class, without getting motion sick from long exposures.Still-higher spin rates (16+ rpm) induce blood pressure gradients head-to-toe in a standing individual, that are just unacceptable for long term exposures.Stand up,and you faint.

3-dimensional objects typically have 3 axes.About these axes these objects have properties called “mass moment of inertia” that relates to spin dynamics.Usually,higher moment of inertia correlates with a larger dimension along some axis perpendicular to the actual spin axis.These are typically proportional to mass,but proportional to the square of its distance from the center of gravity.

There are two (and only two !!) stable spin modes for most objects:  about the axis for highest moment of inertia (longest dimension),  and about the axis for lowest moment of inertia (shortest dimension).  The first case is exemplified by a baton twirler’s spinning baton,  and the second case is exemplified by a spinning bullet or artillery shell.  There are no other stable modes of spin.See Figure 1.

Clearly,  building a “spinning rifle bullet” 112 m in diameter at 4 rpm for one full gee at its outer girth is not so very feasible:  this is just too big to afford at this time in history.  But spinning a smaller-diameter “something” that is 112 m long,  end-over-end at 4 rpm,  for 1 gee at each end,  would indeed be a feasible thing to attempt!  That says select the baton-spin mode for practical designs.

We already know a lot about the transient dynamics of spinning rigid objects,  something important for spin-up and spin-down,  as well as for applying any thrust while spinning.  There would be no fundamental engineering development work to design a long,  narrow spacecraft that spins end-over-end for artificial gravity.  There would only be proving-out the specific design in tests before we use it.

The most-often-proposed alternative is a cable-connected structure,  because it is conceptually easy to reel-out long cables between two small objects.  Cable-connected transient dynamics for spin-up and spin-down,  and especially for applying thrust while spinning,  are incredibly complex and still not very well-known.  “You cannot push on a string”,  that is the complication!  So there is a huge fundamental engineering development effort needed,  beyond just proving-out the actual design to be used.

What this really says is that the preferred near-term spacecraft design is a long and rigid,  more-or-less cylindrical shape,  to be spun end-over-end,  baton-style.  This will generate varying artificial gee from a maximum near the ends,  to zero at the spin center. 

We know that microgravity vs gravity has no impact while prone sleeping,  or else Earthly bed rest studies would not be a decent surrogate for some of the in-space microgravity effects.  That means you can put the sleeping quarters in the low gravity section of the spacecraft near the spin center,  and just put the daily workstations in the full-gravity sections of the spacecraft near the ends.  See Figure 2.

Radiation Hazards

There are basically three radiation hazards to worry about:  galactic cosmic rays (GCR),  solar flare events (SFE),  and the Van Allen radiation belts about the Earth (or similar belts around some of the outer planets).  All three hazards are atomic or subatomic particles,  just at different speeds and quantities.  The threats they pose are location-dependent.

GCR is a very slow drizzle of really-high-speed particles,  moving at a large fraction of the speed of light.  Particles that energetic are very difficult to shield against,  because they penetrate deeply into shielding material,  and quite often create “secondary showers” of other harmful radiation when they strike the atoms in the shield material.  If the shielding atoms are low atomic weight,  the secondary shower effect is greatly reduced.

GCR comes from outside the solar system.  Its quantity is affected by the solar wind,  in turn affected by the sun’s sunspot cycle,  which is about 11 years long.  The solar wind is stronger when sunspots are active,  making GCR lower in the vicinity of the Earth-moon system at that time.

From NASA’s radiation effects website (ref. 1) I obtained these values that apply in the general vicinity of the Earth-moon system.  GCR maximizes at about 60 REM per year when the sun is quiet,  and minimizes at about 24 REM per year,   when sunspots are most active.  To “calibrate” the effects of what may be unfamiliar units of radiation,  the natural Earthly background radiation is about 0.3 REM per year (and up to 10 times higher in some locations),  and a lethal dose would be 300 to 500 REM accumulated in a “short time”,  meaning hours to a week or so.  (Just for information,  1 Sievert is 100 REM.)

The NASA astronaut exposure standards are set at about twice the levels allowed for Earthly nuclear workers.  Those NASA standards are no more than 50 REM per year,  no more than 25 REM in any one month,  and a career limit that varies with age and gender,  but maxes-out at no more than 400 REM accumulated over an entire career.  These career limits are predicated upon a single-handful percentage increase in the likelihood of late-in-life cancer.  

Clearly,  with a very modest shielding effect (to reduce worst-case 60 REM to an acceptable 50 REM annual),  GCR is not the “killer” it is often portrayed to be.

SFE (solar flare events) are different.  They are much lower-speed particles,  much easier to shield,  but there is an incredibly-huge flood of them,  when these events happen.  They come in very-directional bursts from the sun,  at rather erratic intervals.  There are usually more of them during times of active sunspots,  but they can indeed happen when the sun is quiet.  They come at irregular intervals measured in durations of “several months apart”.

The intensity of a burst can vary wildly from only tens of REM received over a few hours,  to tens of thousands of REM received over a few hoursThe median dose would be multiple thousands of REM over a few hours.  Obviously,  for unshielded persons,  the great bulk of events like this (those over about 300-500 REM) would be fatal doses,  and it is an ugly,  irreversible,  and miserable death.  There was a massively-fatal-level event in 1972 between the last two Apollo missions to the moon,  and a low-intensity (non-fatal) event during one Apollo mission to the moon.

We had chosen to ignore this SFE threat during Apollo because the short duration of the missions (at most 2 weeks) was small,  compared to the typical interval (several months) between events.  But,  had a large event hit during an Apollo mission,  the crew would have died in space in a matter of hours.  As it turns out,  this actual record shows that Apollo’s “ignoring-the-risk-as-low-probability”-assumption was not a good assumption to make!  That’s 20-20 hindsight,  but it is still a crucial lesson to learn!

For an extended or permanent return to the moon,  or going elsewhere,  radiation shielding is obviously imperative!  On Earth,  we are protected from these SFE’s (and the GCR) by both the Earth’s magnetic field and its atmosphere.  These are a very real threat anywhere outside the Earth’s magnetic field!  In low Earth orbit,  we are protected only by the magnetic field,  and the background exposure there is higher than down on Earth,  but still much less than beyond the magnetic field.

The Van Allen belts are concentrated regions of these same radiation particles trapped in the Earth’s magnetic field.  The intensity is lethal on a scale of days-to-weeks,  but tolerable on a scale of hours-to-a-day-or-so.  The inner boundary is not sharp,  but this is generally considered to become a problem at about 900 miles orbit altitude,  and extending many tens of thousands of miles out from the Earth. 

The exception is the “South Atlantic Anomaly”,  where the inner side of the Van Allen belt dips down locally to low Earth orbit altitude (100-300 miles).  Satellites and spacecraft in high-inclination orbits inherently pass through the South Atlantic Anomaly every several orbits.  The ISS does indeed encounter this threat,  it being short “flashes” of exposure that accumulate over time,  but these still fall well within the astronaut exposure standards (no more than 50 REM annually,  no more than 25 REM in any one month).  Their main effect is accumulation toward career limits.

Spacecraft traveling to the moon or elsewhere must transit the Van Allen belts.  Because of the potential for lethal exposure if you linger within them,  such transits must be made quickly!  Apollo did this correctly,  transiting within only several hours.  Given the state of today’s electric propulsion technology,  this rules out using electric propulsion for people to leave Earth orbit for the moon or elsewhere,  because the spiral-out time is measured in multiple months.  That would quickly accumulate to a lethal exposure,  even with some shielding.  

Passive Shielding

The same NASA radiation site has data regarding the shielding effects of typically-considered materials.  Those are hydrogen,  water,  and aluminum.  Mass of shielding above a unit exposed area turns out to be the “correlating variable”,  and 15-20 g/cm2 seems to be enough to generally address the worst SFE. 

Hydrogen has the lowest density,  requiring the thickest layering,  but also has the least secondary shower potential,  when used against GCR.  211 to 282 cm of liquid hydrogen suffices. 

15-20 cm of water is 15-20 gm/cm2,  same shielding effect as a really thick layer of hydrogen.  Water molecules are still light enough not to have much secondary shower risk. 

Aluminum would be the thinnest layer,  but with the greater secondary shower effect.  However,  of the practical metals,  its atoms are the lightest,  and this secondary shower effect is deemed tolerable with it.  6-8 cm thick aluminum plate would be required.  That is quite out-of-line with current spacecraft hull design practices:  something nearer a millimeter.

Other materials based on polymers,  and even most rocket propellants,  are light-enough atoms to be effective shielding with a low secondary shower risk,  yet with densities roughly in the same ballpark as water,  for a thinner layer thickness.  So,  any of these could be practical shielding materials!

Because weight is critical,  what you have to do is not simply add shielding weight to your design,  but instead rearrange the distribution of masses you already otherwise need,  so that they can also serve as radiation shielding.  You will need meteoroid shielding and thermal insulation,  and any manned craft will have water and wastewater on board,  as part of the life support system.  All spacecraft will need propellant for the next (and subsequent) burns.  You use a combination of these,  acting together.

The real suggestion here is to use water,  wastewater,  and next-burn propellant tankage as shadow shields,  in addition to the meteoroid protection and thermal insulation materials that the manned modules require anyway.  It doesn’t take much of this at all to cut the worst-case 60 REM/year GCR to under 50 REM/year.  It takes only a little more to cut worst-case SFE to safe short-term exposure levels. 

If you cannot protect the whole manned interior,  then the flight control station becomes first priority,  so that maneuvers can be flown,  regardless of the solar weather.  Second priority would be the sleeping quarters,  to reduce round-the-clock GCR exposure further.  These seriously constrain spacecraft design.

See Figure 3 for one possible way to do this,  in an orbit-to-orbit transport design concept.  This would also be a baton-spin vehicle for artificial gravity during the long transit.  Plus,  the habitation (“hab”) design requires a lot of interior space for the mental health of the crew,  something else we know is critical.  Somewhere between 100 and 200 cubic meters per person is needed as a minimum,  and at least some of it must be reconfigurable as desired by the crew. 

Spin-up is likely by electrically-powered flywheels in the center module.  The vehicle is spun-up after departure,  and de-spun before arrival.  If a mid-course correction is needed,  the vehicle could be de-spun for that,  and spun back up for remainder of the transit.  

Note in the figure how the arrival propellant and the water and wastewater tankage has been arranged around the manned core to provide extra shadow shielding,  for really effective radiation protection.  The manned core modules are presumed insulated by polymeric layers that also serve as meteor shielding (while adding to the radiation protection,  without being driven by that issue).  The pressure shell on the inside of this insulation should be unobstructed by mounted equipment,  so that easy and rapid access for patching of holes is possible. There is not time to move stuff when a compartment is depressurizing!  Ethics!

At departure,  the vehicle can be propelled by a different propellant and engine choice,  since departure is a short event.  The arrival propellant is likely a storable to prevent evaporation losses.  Storable return propellant tankage sets can be sent ahead unmanned,  for docking in orbit at the destination.  

There is an emergency return capsule (actually two capsules) mounted at the center module,  each one enough for the entire crew.  (“Bailout” at Mars presumes a rescue capability already exists there,  so we need redundant engines instead.) Emergency bailout,  upon a failed burn for returning to Earth orbit,  is the main function of this capsule.  Routinely,  it could return a crew to Earth from the spaceship,  once it is parked safely in Earth orbit.

This kind of orbit-to-orbit transport design could serve to take men to Mars or to the near-Earth asteroids and comets.  For Mars,  the lander craft could be sent ahead unmanned to Mars orbit,  and none are needed to visit asteroids.  But you cannot send return propellant ahead on an asteroid mission.

By refueling and re-supplying in Earth orbit,  such a manned hab design could easily be used for multiple missions,  once built.  Care must be taken in its design and material selection to support many thousands of cycles of use.  Thus the craft could safely serve for a century or more,  updated with better propellants and engines as the years go by.

There I went and wrote a basic “how-to” document for practical and ethical interplanetary spaceship design!


These first few sections so far have been reprised (with edits) from “Just Mooning Around”,posted 7-14-19.Everything that follows is new.

Mars Mission Outline 2019:Overall

The new version uses a larger orbit-to-orbit transport,and recovers the solar-electric tugs that preposition unmanned assets at Mars for the manned mission (2016 did not).It uses similar (but larger) landers as the 2016 version,and it still jettisons the Earth departure stage without recovery.

That last could be addressed by fitting the departure stage with a second propulsion system, possibly electric,and putting it into a 2-year-period orbit after stage-off.Then it could be captured into Earth orbit for reuse.That recovery possibility is beyond scope here in the 2019 version. Consider it to be a “future update”.

Main point here:if one does spin gravity in a baton-spin mode,the resulting transit vehicle is ill-adapted for a direct entry at Mars,or a direct entry at Earth.Such a design is far better-adapted as an orbit-to-orbit transport,with any Mars lander function relegated to a separate vehicle,sent separately.Long-life reusability also points toward an orbit-to-orbit transport design,free of entry heat shield requirements.It means we base our exploration forays onto the surface of Mars from low Mars orbit.

The resulting mission architecture requires that both the landers and the Earth return propellant get sent ahead unmanned to parking orbit about Mars,with the manned orbit-to-orbit transport arriving afterward,and rendezvousing in Mars orbit with those items.This powerful concept is not unlike the Lunar Orbit Rendezvous architecture that made it possible to mount each Apollo landing mission with only one Saturn 5 booster.See Figure 4 for the overall mission architecture.


The landers themselves are envisioned as one-stage reusable articles that make multiple flights,  based out of low Mars orbit.  Sending 3 landers ahead with their propellant supply allows one lander to make a landing with only part of the human crew,  with a second lander in reserve as a rescue craft.  Thus,  there is a “way out” even during the landings,  unlike with Apollo! 

Because of storability concerns,  the wisest choice is that the lander propellant and engine design be the same as the transport propellant and engine design.  This maximizes the interchangeability of engine hardware and propellant supplies,  in the event that there are mishaps from which to recover,  without aid from Earth.  It also simplifies the overall design and hardware development and prove-out.

The presence of a third lander allows one lander to become unserviceable,  while still maintaining the reserve rescue lander capability,  without which landings so far from Earth become too risky to ethically attempt.  This is shown in Figure 5,  including the velocity requirements for the lander design. 

The initially-sized version of the lander design concept was used in the 2016 posting,  and came from one of the options explored in another posting titled “Reusable Chemical Mars Landing Boats Are Feasible”,  dated 31 August 2013.  These landers are resized somewhat for this posting.

Note that for a rescue possibility to exist,  some of the crew must stay in the transport in low Mars orbit,  while others descend to the surface in a lander.  Because we do not know how therapeutic Mars’s 0.38 gee gravity might be for the surface crew,  I suggest we spin the transport for artificial gravity while it is in orbit,  de-spinning for lander departures and arrivals.  Thus everybody stays fully healthy no matter what,  while we alternate crews on the surface.

Now,  overall,  it is worst-case 9 months to and from Mars,  and in any case,  13 months at Mars waiting for the orbital “window” to open for the voyage home.  That last is simply inherent from the choice of min-energy Hohmann transfer orbits.  That leaves a long time for the crew to explore on Mars.  That plus the possibility that the initial landing site might not prove to be desirable,  makes it wise to plan for multiple landings,  at possibly-multiple sites. 

Basing exploration forays from low Mars orbit is what makes multiple landings at multiple sites possible at allNo other mission architecture can provide this capability.

It is that orbit-based architecture allowing for multiple landings which lets us alternate roles for the crew,  so that all of them get to spend time on the surface of Mars (unlike what was possible with Apollo).  With a mission crew of 6,  that means we could send down alternating crews of 3 in the lander,  while the other 3 do science from orbit and provide the critical watchdog rescue capability with the other two landers (two for the reliability of redundancy).  It is already known that odd numbered crews fare better in hazardous situations,  there being no possibility of the stalemate of ties,  in decision-making.

Given the existence of the rescue capability from low Mars orbit,  we can address lander reliability in two ways,  thus increasing the odds of success,  and also the odds of still saving the lander crew,  if things go seriously wrong.  (We are ethically bound to do this!)  First,  the lander must use redundant engines,  so that if one fails,  the remaining engine (or engines) can still perform the mission.  

Second,  the crew piloting cabin could be rigged as an abort-to-surface (or abort-to-orbit) capsule,  in the event that too many redundant engines fail,  or that there is some overall catastrophic failure of the lander.

The minimum number of landings is two,  one for each half of the crew.  Allowing some time for reconnaissance-from-orbit prior to the first attempt,  and for preparations for returning to earth,  we can plan on 12 months total for the landings,  splitting the remaining month between those other two needs in orbit about Mars.  That does cover up to two possible landing sites in the one voyage to Mars!

The surface crew will live inside the lander on the surface.  That means it must carry them,  their exploration gear,  and up to 6 months of life support supplies,  on each trip.  More exploration gear could be carried to the surface if we shorten the stay for each lander. 

If four trips will be made,  that’s 3 months each (not 6),  and one can trade away life support supplies for extra exploration gear carried down.  That could cover up to four possible landing sites in the one trip to Mars,  and each crew of 3 making 2 trips,  all with the same overall resources sent to Mars,  excepting the total lander propellant supply.

Continuing that logic,  if 6 trips are planned,  that’s 2 months each,  each crew of 3 making 3 trips,  and a higher weight of exploration gear relative to life support supplies.  That’s up to 6 separate sites that could be explored in the one voyage to Mars!  Or,  12 trips of 1 month each,  which is up to 12 sites explored.  Since the lander propellant is sent ahead by SEP,  it is rather easy to afford such a capability. 

The biggest mass ratio-effective burn for the lander is the ascent burn,  which can be at significantly-reduced payload,  since wastes can be left on the surface along with some exploration gear,  while the weight of a plethora of samples is far less than the weight of gear and supplies during the less-demanding descent.   That makes the overall 5.22 km/s delta vee far more affordable with an overall realistic mass ratio and storable propellant specific impulse (Isp).  

Those considerations very dramatically impact and constrain the design of the lander.

Sending Assets Ahead Unmanned

The unmanned transfers can be done more efficiently (lower total mass to be launched) with solar electric propulsion (SEP).  The manned transport uses short-burn chemical rocket propulsion to avoid long spiral-out/spiral-in times.  (An SEP-based transport would give the crew a lethal radiation dose spiraling-out through the Van Allen belts on departure from Earth,  and again spiraling-in through the belts on return to Earth.)  At least approximately 0.1 gee vehicle acceleration is required to qualify as a gravity loss-free “short burn”. 

This prepositioning of assets at Mars using SEP was also a part of my 2016 Mars mission posting.  The differences here are that I recover the SEP “tugs” for reuse on future missions,  and that I use a larger “hab” for the orbit-to-orbit transport. 

Earth Departure of Manned Transport

The Earth departure can be done with higher-performing LOX-LH2 tankage and engines on one end,  that are staged off after the burn.  To recover these,  a higher aphelion orbit with a 2 year period is required,  plus some sort of propulsion to return to Earth orbit.  This could be electric,  or some storable propellant rockets.  (Expecting LOX-LH2 cryogens not to evaporate over a 2 year period is just nonsense!)  I did not include that here,  but it is required for more reusability.  That’s a future growth item.

Velocity Requirements for the Mission

The orbital mechanics of min-energy Hohmann transfer determine the minimum velocity requirements for the manned (and unmanned) vehicles,  as well as the one-way travel time.  Shorter flights require more energy,  which is more propellant and tankage that must be sent to low Earth orbit and assembled. 

The basic velocity requirements for the manned orbital transport are shown in Figure 6.  These take the form of unfactored orbital mechanics values serving as the mass ratio-effective values for vehicle design.  This is allowable because all these chemical rocket propulsion burns are “short” and exoatmospheric.  The resulting mass-ratio-effective design values are given in Figure 7.

For only Mars arrival with the manned transport,  there is a need for a rendezvous propellant allowance.  It is necessary to adjust orbital position to coincide with the assets sent ahead.  As a wild guess,  add another 0.2 km/s delta vee to the value shown in Figure 7 as the mass ratio-effective value for design.  

For the assets sent ahead with SEP,  design velocity requirements are much more problematic.  There are no drag losses,  but the gravity losses are huge,  since the burns are months long!  For a rough rule-of-thumb estimate,  just use twice the values in Figure 7.  That is what I did here. 

Propulsion Estimates

No particular existing chemical rocket engine’s characteristics were used.  Ballistic estimates were made “from scratch” using shortcut methods.  For both the transport and Earth-departure engines,  it was assumed that no gas used to drive pumps was dumped overboard,  meaning 100% of the hot gas generated went through the propulsion nozzle.  This requires an efficient engine operating cycle. 

Estimates were made from 1000-psia data for chamber characteristic velocity and gas specific heat ratio,  using standard ideal-gas compressible flow methods to develop vacuum thrust coefficient (to include the effects of a nozzle kinetic energy efficiency reflecting streamline divergence).  The c* and r “constants” vary with chamber pressure in a way that conforms to empirical ballistic methods I have long used successfully.

This gets us to specific impulse (and thus effective exhaust velocity) for vehicle mass ratio determinations with the rocket equation dV = Vex ln(Wig/Wbo).  The actual design thrust level is driven by vehicle mass and the min 0.1 gee acceleration requirement,  which sizes throat (and exit areas) via the thrust/throat area/thrust coefficient equation F = CF Pc At.  That leads to real engine dimensions.  For not-quite-the-highest-tech in engine design technology,  a good “wild guess” for engine weight would be thrust/50,  both in force units,  figured at 1 gee Earth gravity for the weight. 

Assuming redundant engines for safety and reliability,  these rockets won’t be simultaneously run at full thrust.  For vacuum-only operation,  there is no need for really high chamber pressure,  and there is no need to worry about backpressure-induced separation effects,  because there isn’t any backpressure.  6-7 mbar on Mars is also effectively no backpressure at all,  so the lander engines can be the same vacuum design as the transport engines.

Reflecting those considerations,  I assumed 1000 psia at max thrust,  typical operation at 500 psia,  and min throttled-down pressure 200 psia.  Others may disagree,  but that is what I did.  The higher the Pc,  the higher the c*,  and thus the higher the Isp.  But so also the higher is the weight of the engine.

The data I got for the NTO-MMH storable transit engines are given in Figure 8.  The data I got for the LOX-LH2 Earth departure engines are given in Figure 9.  For both I assumed an expansion bell equivalent to a constant 15 degree half-angle conical bell,  leading to a kinetic energy efficiency of 0.983 for the nozzle efficiency.  Any real-world curved bell will have an average half angle not far at all from that value;  it will be slightly shorter than the equivalent conical bell,  and just about the same efficiency. 

The solar electric propulsion is more problematical in its characteristics,  it being currently available only in small sizes,  with scaleup efforts underway at both Ad Astra and NASA.  What is important for vehicle design purposes would be thrust/weight for the actual electric thruster equipment,  its operating specific impulse,  its electric power/thrust requirement,  and the type and phase of its propellant (liquids or solids are easier to store at lower total tankage weight than gases).   

Add to that the producible electric power/panel area,  the weight/panel area,  and miscellaneous equipment weight (if any),  for the solar power supply equipment,  and for autonomous robotic vehicle guidance.  The size of the thruster’s thrust relative to the full vehicle weight should probably fall near what the current small thrusters on satellites use:  something near or above 0.001 gee.

Here are the values for the putative system I “chose”,  it being something that does not yet exist,  but likely could be made to exist near-term.  Bear in mind the available solar power at Mars is half that at Earth (Mars actually sizes the panels).  The value shown for electric power/area of solar panel is for near-Earth space,  turned to face the sun directly. This data represents a Hall-effect device on iodine.

The solar photovoltaic power per unit area was estimated as the solar constant at Earth (in space 1353 W/m2) multiplied by a 20% conversion efficiency of sunlight power to electric power. That represents a high-tech space-industry type of solar cell.  The weight was estimated from reported data for the Alta Devices Alta 5x1 2J and Alta 5x1 1J satellite solar panel devices.  The miscellaneous equipment is not structure,  that is in the weight/area figure for the panels.  It is the mass of the autonomous guidance equipment,  including things like star trackers,  computers,  communications,  and accelerometers. 

Space Hab for the Crew:  Characteristics

I based these guesses off the Bigelow Aerospace B-330 space station module design as seen on the internet (ref. 2).  This is the big commercial product,  not the simple,  small BEAM unit attached to ISS for testing and evaluation by NASA.  These are nominally 15.7 m long and 20 metric tons.  They are somewhat inflatable,  and feature a core equipment and framing structure around which the inflated hull is unobstructed.   There is a meter of layers of micrometeoroid shield that also serve as thermal insulation,  and as low-molecular-weight radiation shielding.   Each module contains some 330 cubic meters of interior space. The hard core protrudes on one end,  providing a place for solar panels. 

The modules of the orbit-to-orbit transport cannot be exactly these B-330 modules,  but can be something rather similar!  Docking multiple modules end-to-end creates the baton-shaped vehicle this mission design needs.  The modules must have external features of some type that allow tankage to be attached around the outer periphery,  and internal fold-out decks as part of the core.  The center module must be very stout,  and contain big electrically-driven flywheels for vehicle spin-up and spin-down,  plus places to dock space capsules.

It would seem wiser to put big solar panels on the center module,  with the docked capsules,  and the flywheels inside,  where spin forces are zero-to-minimal.  It is likely to be hard shell,  not an inflatable,  for strength.  That module is also likely to be quite heavy.  As a wild guess,  call it 16 m long and 40 tons.  The others can be nominal 16 m long,  and nearer 20 tons,  reflecting inflatable pressure shell along almost the entire length,  plus the features for attaching external tankage.   Call the internal volume 350 m3 each as a best guess,  excluding what the hard core occupies.

Counting the center module,  some 7 modules each 16 m long docked end-to-end is 112 m long,  for 1 full gee at each end if spun at only 4 rpm.  That basic structure would total 160 metric tons,  using the guesses in the previous paragraph.  To that one must add masses for crew and 2 space suits each,  their personal effects,  and personal equipment (call it 0.5 metric ton per person as a guess),  and for fully-expendable supplies of food,  water,  and oxygen (call it 0.75 metric tons per person per month,  knowing that these are just “reasonable guesses”).  Crew and supplies must fit within the vehicle,  which has (for the 6 modules not filled with flywheels and heavy equipment) some 2100 m3 volume. 

If one assumes half the volume is packed supplies,  and also a crew of 6,  that leaves some 175 m3 per person as living space available.  That’s about like 3 large living rooms in a typical middle-class house.  That seems adequate at first glance,  if it is well distributed,  and some part of it is reconfigurable at some level.  

The crew weight allowance is 3 metric tons,  and the packed supplies mass is about 4.5 tons per mission month.  If the mission is 31 months long (9 months transit,  13 months at Mars,  9 months return),  that’s about 140 tons of supplies,  with no margin for error.  So call it a nominal 150 tons.  This presumes no recycling or growing-of-food in space or on Mars.  It’s a worst-case deal,  but we can do this “right now”.   

So,  the empty hab section is estimated at 160 tons.  It gets loaded with about 150 tons of supplies,  allowing for 7.5% safety factor on supply mass,  and loaded with about 3 tons of crew with their suits,  equipment,  and personal effects.  Fully loaded,  that’s 313 tons.  That would be crew of 6,  and supplies for a 31 month mission plus a small margin.  See Figure 10.  Figure 11 shows an image of the spreadsheet where these numbers were calculated.  Yellow highlighting denotes inputs.  Some selected outputs are highlighted blue. 

Assumed depleted at a constant rate,  the supplies total 150 tons at departure,  109.5 tons at Mars arrival,  51.0 tons at Mars departure,  and not-zero at 10.5 tons at Earth arrival,  assuming the safety margin is not consumed.  This presumes wastes are dumped overboard with no recycling at all!   This dumping reduces the effective mass of the hab section,  at each mission segment,  a benefit to propellant required. 

We can already do somewhat better than that with recycled water,  but this is a worst case estimate!  Yet this open-cycle assumption gets the smaller propellant supply for return to Earth.  “Efficiency” is not always beneficial:  that is too often presumed erroneously!  Jettisoned mass reduces next-burn propellant requirements.  That’s just physics you cannot fight!

Sizing the Manned Transport and Its Return Propellant

The fundamental notion for sizing propellant supplies for the four events (Earth departure,  Mars arrival,  Mars departure,  and Earth arrival) is that the mass of the loaded,  crewed hab,  plus the mass of all propellant tankage,  plus the mass of the engines,  is the ignition mass.  That minus the mass of propellant burned from that tankage is the burnout mass.  That produces a mass ratio for the burn,  and the delta-vee it will produce,  which must meet or exceed the requirement for that burn.  This is subject to the constraint that we want 0.1 gee or thereabouts as a min vehicle acceleration at each burn.

To do this,  one must estimate the ratio of propellant to loaded tank mass for the added tankage.  This has to reflect a long,  slim tank geometry for docking multiple tanks around the periphery of the hab,  and it must account for the mass of the docking structures needed to achieve that result.  As a guess,  I am assuming that the empty tank inert mass (with all those fittings) is 5% of the loaded tank mass,  so that the contained propellant is 95% of the loaded tank mass. 

To that end,  I used a series of calculation blocks in a spreadsheet worksheet to run the calculations.  Again,  inputs are highlighted yellow,  and significant outputs are highlighted blue.  Figures 12,  13,  and 14 show the results. 

Bear in mind that the loaded tank mass for the Mars and Earth arrival burns must be part of the “payload” for the Earth and Mars departure burns,  respectively.  They are unique in this way.  That means the dead-head payload is the appropriate hab mass plus the mass of the next burn’s loaded tanks.  The current burn’s tanks must push this (plus the added engine mass) to the required delta-vee for that burn. 

Added engine mass is handled by an iteratively-applied tankage scale-up factor just slightly over unity.

As it turns out,  finding the propellant tankage mass to push the hab to the required delta-vee is not an excruciating iterative process.  You first find the mass ratio MR that is required from the required mass ratio-effective delta-vee,  and the propulsion’s effective exhaust velocity,  by the rocket equation.  Ignoring the mass of the engines themselves,  it turns out to be closed-form to find the loaded tankage mass Wtf from that mass ratio,  and the total “dead head” mass to be pushed in that burn. 

For both departures,  the “dead head” mass is the appropriate loaded hab mass plus the loaded mass of the corresponding arrival tankage.  For both arrivals,  the “dead head” mass is just the loaded hab mass.  This can be corrected at the 1 or 2% level for total engine mass later,  to ensure fully meeting the delta-vee requirements,  simply by scaling up the loaded tank mass Wtf with a factor applied iteratively until delta-vee produced meets the requirement.   

Wtf = Wdead (MR – 1)/(1 – MR f)  where f = Wt/Wtf and Wt is dry tank mass

That’s the orbital transport rough-out design for Mars.  It can get there to low Mars orbit from low Earth orbit where it was assembled.  It can rendezvous with its Earth return propellant,  the Mars landers,  and the Mars lander propellant supply,  all three of which were sent ahead by electric propulsion.   The nonrecoverable items are the Earth departure stage and the empty Mars departure tanks.  The empty Mars arrival tanks are left in Mars orbit.  Everything else about this design is recovered in low Earth orbit.

Note that this ship is 1413 metric tons,  as assembled and loaded in low Earth orbit,  ready to go to low Mars orbit.  Its use requires that some 997.26 metric tons of loaded propellant tanks be sent ahead to Mars for the return propellant supply.  In order to actually make landings on Mars as staged out of low Mars orbit,  the landers and their propellant supply must also be sent ahead to low Mars orbit. 

With much bigger propellant tankage,  this same design could take men to a near-Earth asteroid.  For such missions,  landers are not needed,  and there is no practical opportunity to pre-position return propellant,  except many years ahead.  Those missions are far more difficult.  Analysis of one is not attempted in this posting.

Sizing the Lander and Lander Propellant Supply

The lander payload is its crew,  their suits and personal equipment,  plus an amount of life support supplies that depends upon how long the crew will live in the lander on the surface,  each landing.  The de-orbit burn for a surface-grazing ellipse is a trivial 50 m/s delta-vee.  Most of the deceleration is aerodynamic drag,  effectively terminating at end-of-hypersonics at Mach 3,  just about 1 km/s velocity,  but at a low altitude because of the high ballistic coefficient.  That altitude is only about 5 km

From there,  deceleration is by retropropulsion alone,  with a large “kitty” to cover hover and divert requirements.  Assuming 1 km/s velocity at 5 km altitude,  along a straight slant trajectory at 45 degrees,  the average deceleration level required is 70 m/s2,  or 7.211 gees,  which with the lander mass,  sets the required engine thrust level for landing.  That is a rough ride,  about twice the rigors of return from low Earth orbit,  and justifying all by itself the maintenance of full crew health with artificial spin gravity!

The lander is a one-stage reusable “landing boat” intended to make multiple flights,  each fueled from a propellant supply sent with it to low Mars orbit.  Factored,  the mass ratio-effective descent delta-vee is just about 1.5 km/s.  Propellant is storable NTO-MMH,  to preclude evaporation losses and massive energy requirements to prevent freezing or boiling.  The ascent must account for small but non-zero gravity and drag losses (about 2% of velocity),  and a “kitty” for rendezvous maneuvers.  That mass-ratio-effective delta vee is just about 3.62 km/s. 

The payload requirements for crew,  equipment,  and supplies as a function of surface duration are given in Figure 15,  along with a crude estimate of the “larger-than-minimum” vehicle inert weight fraction that is appropriate to the necessary structural robustness,  and to the equipment required to function as a reusable entry-capable vehicle,  and as a surface habitat.  Conceptually,  the lander is sketched in Figure 16.  Some of its backshell panels double as cargo load/unload ramps.  Most of the cargo space can be isolated and pressurized as living space,  once unloaded.  The piloting cabin is the abort capsule,  something somewhat similar to a crew Dragon from Spacex.  This thing is NOT a minimalist lander the way the Apollo LM was.  

The ascent payload is smaller,  since most (but not all) the supplies are used up (and wastes left behind) at ascent liftoff.  There is a generous allowance for Mars samples to be returned to the orbital transport. This has to be taken into account in calculating the actual vehicle masses,  since the two delta-vees are handled at two different payload fractions,  in the one vehicle design.  That process is inherently iterative,  as shown by the data given in Figure 17.  

In order to determine these numbers,  one guess a value for the max lander mass,  which is ignition-at-descent (Wig-des).  The inert fraction times this gives the vehicle inert mass Win.  The ascent and descent payloads are determined vs mission surface duration separately.  The mass ratios already determined are used to estimate propellant masses. 

The ascent propellant mass Wp-asc is determined first as (MR-asc – 1)(Wpay-asc + Win),  then the descent propellant mass Wp-des as (MR-des – 1)(Wpay-des + Win + Wp-asc),  treating the ascent propellant as part of the effective “payload” during descent.  The descent payload plus both propellant masses plus inert mass sum to the result for descent ignition mass. 

The input guess for descent ignition mass is then adjusted iteratively,  until it converges to the result for descent ignition mass.  This is done by simple trial and error in the spreadsheet.  There is such a result computed for each of 4 possible surface durations that divide evenly into the 12 months available.  These results are then the inputs for a characterization of the lander sizing as a function of design surface duration. 

For the selected 2-month duration (corresponding to 6 total lander flights),  those results are given in Figure 18.  These show ascent and descent weight statements,  confirmation of delta-vee capability,  and characterization of vehicle mass fractions,  plus the propellant supply required to cover the appropriate number of flights.  Similar tables exist in the spreadsheet for the other 3 durations,  but those are not shown here.   

Figures 19 and 20 show the trade-off of vehicle sizes and propellant supply sizes versus surface duration options.  The selected design is near the “knee” in the curve of number-of-flights vs surface duration,  at 2 month duration for 6 flights.  For shorter duration,  the required propellant supply is significantly larger.  For longer duration,  the required propellant supply is smaller,  but not so significantly smaller. 

The lander size itself is significantly affected by the design surface duration,  being larger at longer duration.  The 2-month duration selected limits this affect,  without so significantly penalizing the payload fraction (which ranges from about 2 to about 3%).  The selected 2-month duration is also near the “knee” in that curve.  Longer durations do not improve this as much as was gained going from 1 month-12 flights to the selected 2 month-6 flights option. 

For this selected design (6 two-month surface stays),  three landers fueled and loaded with supplies,  less crew,  suits,  and personal equipment,  each massing 376.5 metric tons,  must be sent to Mars along with some 1764 tons of propellant to support all 6 flights.  If 95% of the tank weight is propellant,  the mass of loaded tankage to be sent is some 1856.8 metric tons.  If sent as tanks docked to each of the 3 landers,  that’s a 376.5 ton lander plus 619 tons of loaded propellant tanks. 

The “smart” thing to do from a reliability / self-rescue standpoint is to send the transport return propellant with those same three landers,  so that if one is lost,  the transport can still return safely by drawing the shortfall instead from the remaining lander supplies.  That return propellant was determined above to be 997.26 metric tons of loaded tanks.  Divided by 3,  that’s an additional 332.4 metric tons of Earth return propellant tankage sent to Mars with each lander. 

That makes each lander plus propellant tanks a 1327.9 metric ton item to be moved by solar electric propulsion from low Earth orbit one-way to low Mars orbit.  Each such is thus quite comparable to the departure mass of the manned orbital transport.  That would not be true at the other surface durations. 

There are 6 landings to be made,  and three such landers sent to Mars.  Distributed evenly,  that is two flights per lander minimum,  and 6 maximum.  Bear in mind that only one lander is sent to the surface at a time,  carrying a crew of 3,  while the other three crew do science in orbit,  while acting as the safety rescue “watchdog”,  with at least one functional lander,  even if the other one fails.  The worst case is that all 6 flights are made with one lander.  Thus the lander design must allow for at least 6 flights per vehicle,  justifying in part the higher inert mass fraction used in this design rough-out.

Landers get left in low Mars orbit at mission’s end,  when the transport departs for Earth.  Subsequent missions might utilize these assets,  and reduce the sent mass to only more lander propellant.  That possibility argues for much more than 6 flights per vehicle,  in turn a really good argument for the very robust inert mass fraction of 20% used here. Alternatively,  they could be landed robotically.

Common Engine Design for Transport and Lander?

The lander mass is 378 metric tons at ignition,  and 241 at touchdown,  as just determined above.  The average is 309.5 metric tons.  Also as determined above,  the average deceleration required is 70 m/s2.  That translates to 21,665 KN of retropropulsion thrust required to safely land (nominally 22,000 KN).  This is totaled for multiple engines.  Less may be used for ascent,  as such high gee capability is not required for that.  Something nearer 2 gees at ascent ignition mass 236.3 metric tons (4726 KN thrust) is more appropriate.

As described above,  something near 1170 KN thrust from multiple engines is the minimum required for the orbit-to-orbit transport.   This was set by the min 0.1 vehicle gee capability at max vehicle mass,  and still resulted in only large fractional gee capability at min vehicle mass.  This thrust level selection could be doubled or tripled (or more) with relative impunity.  

A worksheet page was set up in the spreadsheet to explore how this could be done,  in two steps.  The results are shown in Figure 21,  which indicate the possibility of using some number of 3600 KN max thrust engines,  throttleable from 20 to 100%.  In the first step,  I input factored thrust requirements,  plus a number of engines,  and a max number of inoperative engines. 

The thrust requirement for the lander descent is based on slowing the average descent mass (as a constant) from 1 km/s to zero,  in a slant path length of 7.1 km,  using the oversimplified kinematic equation V2 = 2 a s.  This is a very high-gee descent!  Reducing that requires not just supersonic retropropulsion,  but hypersonic retropropulsion (starting retropropulsion earlier in the entry sequence).  It is an inevitable consequence of the high ballistic coefficient producing very low altitudes (on Mars) for end-of-hypersonic deceleration.  This is an area for further design work!

The thrust requirement for the lander ascent is its Earth weight,  factored-up just slightly,  to accommodate flight tests on Earth.  That’s “overkill” for Mars with its lower gravity.

The thrust requirement for the orbital transport is based on its Mars departure mass (largest of the masses under storable propulsion) and a min 0.1 gee vehicle acceleration requirement.  This is arbitrarily factored-up by 3 to achieve commonality,  without exceeding max gees ~ 2 at Earth arrival.

That initial result indicated that something like 3600 KN max thrust per engine would be suitable,  with 9 engines in the lander operating at part throttle in descent,  and 4 engines operating at part throttle in ascent,  able to lose up to 3 engines either way,  and still function within limits.  This was explored further,  looking at vehicle gees and engine throttle percentages,  in the second step. 

Up to 3 of these lander engines could cease operation during ascent or descent.  The remainder could supply adequate thrust at 100% throttle or less,  without waiting for lightoff of any inactive engines. That’s an important safety consideration,  which ethics demands!  Two of these same engines would be adequate to push the orbital transport at part throttle,  with only one operating engine still able to supply much more than the demanded minimum thrust.

In all cases,  engines operate between 20 and 100% throttle setting,  and appropriate gee limits are not exceeded.  Min transport vehicle gee requirement (0.1 gee) is exceeded. 

For descent,  the lander retropulsion operates between about 6 and about 9 gees.  This event is only about 14-15 seconds long!!!  “At the last second” to actually land,  some 8 of the 9 engines must be shut down to reduce thrust to nearer Mars weight of the lander (about 749 KN to 872 KN,  depending upon how much propellant was burned) at touchdown,  with the remaining active engine operating at about 21-24% thrust setting.  This single-engine point is the riskiest aspect of the landing,  but it is mitigated by the facts that (1) this engine is already operating,  and (2) it need only continue to operate at reduced thrust for a second or two.

On ascent with a reduced number of engines,  this is 1.2 to 3.6 gees for the lander at full thrust,  far more than is needed to depart against Mars gravity (only 0.38 gee).  Active throttling reduces that some.

The transport operates between 0.3 and 1.8 gees during the return to Earth.  This exceeds the min acceleration requirement,  but not the maximum.  A 3600 KN engine design for this NTO-MMH common engine would resemble the notional sketch in Figure 22. 

If the Earth departure stage at 1350 KN uses 5 engines,  each would be approximately 1350 KN max thrust capability operating at 20% thrust.  Up to 4 could be non-functional,  and still easily meet the overall min departure thrust requirement,  without exceeding 100% throttle.  Higher vehicle acceleration than 0.1 gee is easily obtained,  but even with all 5 engines at full thrust,  it is still only fractional gee.  Such a 1350 KN LOX-LH2 engine would resemble the notional sketch in Figure 23.

Sizing the SEP for the Unmanned Assets Sent Ahead

This item is the most speculative,  because (1) it uses the most assumed data,  and (2) this kind of solar electric propulsion has yet to be scaled up to such sizes to push masses this large.  To cover the gravity losses (both planetary and solar),  I simply doubled the required orbital delta-vee data. 

I simply assumed the average characteristics of small Hall effect thrusters operating on iodine could be scaled way up by simple clustering,  at the same thrust/weight and thrust/power ratios.  And,  I just assumed the characteristics of satellite-sized solar panels could be scaled up to the low-hundred kilowatt range at the same power/area and weight/area ratios.

My approach was a self-contained solar-electric propulsion (SEP) “tug”,  that incorporates the clustered thruster unit,  the solar panels to power it,  sized for reduced sunlight at Mars,  a robot guidance package,  and a low-pressure “tank” to contain the easily-sublimated  and inexpensive iodine propellant.  I used published data for two Busek Hall-effect thrusters,  and for a couple of Alta Devices satellite solar panels,  for these estimates. 

This SEP “tug” is coupled to a dead-head payload for the trip from Earth orbit to Mars orbit,  using all of its 120 clustered SEP thrusters to achieve a milli-gee of vehicle acceleration capability at Earth departure.  That payload is one (of the three) Mars landers (fully fueled and supplied),  plus a 1/3 share of the total lander propellant supply,  and plus a 1/3 share of the manned orbital transport’s Earth return propellant supply.  This dead head payload is over 1300 metric tons.

For the return trip (these “tugs” are fully reusable),  there is no dead-head payload,  only the “tug” and its iodine tank,  still containing just enough iodine propellant to get home.  During the trip home,  only one SEP thruster in the cluster need be used to achieve near a milli-gee of vehicle acceleration at Mars departure.  That leaves many “spares in case the one fails”,  insuring utter reliability.  (Outbound,  the cluster is large enough that the loss of a few thrusters is no significant percentage loss of thrust.)

The size of one such thruster (200 mN,  mN meaning milli-Newtons) falls within the range of thrusters produced today.  This produces adequate acceleration of the unladen vehicle.  The scaleup is by clustering,  not by increasing the size of the thrust in such a device.  The clustering-together of 120 of these units produces some 24,000 mN,  needed to move the laden vehicle at adequate acceleration. 

The resulting SEP “tug” design is depicted in the sketch of Figure 24.  I used a big two-stage spreadsheet worksheet to iteratively size this “tug” system,  examining the 4 “burns” individually.  The second stage of this process fully defines the characteristics of the “tug” and its estimated performance.  This is the tabular data in the partial spreadsheet image shown in Figure 25. 

Hopefully,  this rough-sizing is “overkill”,  due to my just-assumed doubling of the orbital delta-vee requirements.  The intent here is to slowly spiral-out of low Earth orbit to escape,  and continue an accelerating spiral about the sun to an appropriate midpoint,  then use a decelerating spiral about the sun toward capture at Mars.  From there,  it follows a decelerating spiral-in to low Mars orbit.  The return uses the same spiraling processes,  just unladen of dead-head payload,  and at far-lower thrust and propellant requirements. 

Sizing the Earth Departure Stage

Of all the items analyzed,  this is the easiest and most straightforward,  because there is one and only one burn (the Earth departure burn).  Then this stage is jettisoned.  The stage layout concept and sized data were already determined as part of the orbital transport propulsion sizing above.  These data were given as part of Figures 12,  13,  and 14 above,  plus part of the common engine discussion just above,  with sized engine dimensions in Figure 23.

Just to summarize,  the departure stage has 5 LOX-LH2 engines each designed for 1350 KN thrust,  weighing an estimated total of 5.139 metric tons.  The stage comprises LOX and LH2 tankage whose combined dry weight is 41.906 metric tons.  The total propellant load is some 796.210 metric tons.  Thus the loaded stage itself is some 843.255 metric tons.

This stage pushes a fully loaded and crewed hab plus Mars arrival propellant tankage that totals some 569.810 metric tons of dead-head payload.  Total orbital transport vehicle mass,  at Earth departure ignition,  is thus some 1413.065 metric tons.  This was shown in Figure 14 above,  including weight statements and performance.

Not considered here is reuse of the Earth departure stage.  Its engine sizing would be fine,  but it needs larger tanks and propellant to accomplish 2 burns.  The first is to put the orbital transport onto a Hohmann transfer ellipse trajectory. 

After releasing the transport,  it burns a second time to enter an ellipse about the sun with an exactly two-year period.  That way the Earth is there when it reaches perihelion,  thus making recovery feasible at all.

It is just not reasonable to expect that cryogens like LOX and especially LH2 will not completely evaporate away over a 2 year interval.  Therefore,  the reusable form of the stage must also incorporate a second propulsion system storable over long periods.  This added propulsion provides the delta-vee to return to Earth orbit from the 2-year solar orbit perihelion conditions.  

Being unmanned,  there is no reason this second propulsion system could not be solar-electric using iodine.  The stage then executes a spiral-in to low Earth orbit after capture.  The alternative is storable propellants like the NTO-MMH. 

Being out of scope here at this time,  these designs have not been explored.  Consider that as a future upgrade.

Totaling Up the Mission and Its Launch Requirements

This mission to Mars requires a fleet of 4 vehicles to be sent from Earth orbit to Mars orbit.  One of these (the manned vehicle) returns to Earth.  The other three are unmanned assets sent ahead earlier by electric propulsion,  for the crew to utilize when they arrive by conventional rocket propulsion. 

The three unmanned vehicles are identical,  comprising a dead-head payload and a reusable solar-electric “tug” that returns to Earth for reuse,  after delivery of the dead-head payload into orbit at Mars.  

That dead-head payload payload is the same for each of these vehicles:  an uncrewed but loaded and fueled reusable Mars landing boat,  plus 1/3 of the total Mars lander propellant supply,  plus 1/3 of the crewed vehicle’s Earth return propellant supply.  That dead-head payload is 1327.9 metric tons for each of these 3 vehicles.

Each of these three unmanned vehicles totals some 2413.5 metric tons as assembled in Earth orbit,  that being the dead-head payload plus the fueled SEP “tug”.

The crewed vehicle (the orbit-to-orbit transport) comprises the crewed and loaded hab section,  plus the loaded Mars arrival propellant tankage,  plus the expendable Earth departure stage that uses cryogenic propellants.  (All the other rocket propulsion uses the same storable propellants,  and the SEP “tugs” use sublimable iodine to keep the iodine “tank” weight down.)  Ready to depart Earth orbit,  the transport and departure stage total some 1413.065 metric tons. 

The grand total that must be assembled in orbit for the fleet of 4 ships is some 8653.6 metric tons.  For that,  you get 6 landings at up to 6 different places on Mars,  all in the one manned trip to Mars.  That’s 1442.3 tons to support each of the 6 landings,  essentially.  These are 2-month max stays at each landing site. You get all this,  plus a “way out” or a self-rescue capability built into the mission at every step,  plus a fully-healthy crew with radiation shielding and artificial gravity during the transits,  and in low Mars orbit. That’s a lot of benefit for the cost.

Getting Landers To Low Earth Orbit

The selected lander design is just about 378 metric tons,  crewed,  loaded and fueled.  Less crew (and their suits and gear),  that’s just about 376.5 metric tons.  Just about 294 tons of that lander weight is propellant.  So,  a loaded,  crewless,  empty-of-propellant lander is just about 82.5 metric tons.  Remove the supplies,  but leave the surface equipment and rover aboard,  and this is about 77 tons.  Completely unladen,  the lander is about 75.6 tons. 

I looked at SLS (150 metric tons to LEO,  guessing $1,000M per launch),  Spacex’s “Starship” (100 metric tons to LEO,  guessing $150M per launch),  Spacex’s Falcon-Heavy (63 metric tons to LEO flown expendably,  about $85M per launch),  ULA’s Atlas-V (20 metric tons to LEO at about $85M per launch),  and Spacex’s Falcon-9 (20 metric tons to LEO flown expendably,  and $63M per launch). 

The loaded unfueled lander mass of 75.6 metric tons is out of reach of Falcon Heavy,  much less Atlas V or Falcon 9,  even if an 8-meter payload diameter could be flown on any of them.  NASA’s SLS might possibly launch it dry of propellant,  maybe even two of them at once,  although it has yet to fly.  That would be 2 or 3 flights of SLS at $2-3B to put 3 landers into orbit,  unladen of propellant.  It would be 3 flights of “Starship” at $450M total.  The most cost-effective of those two options is “Starship”.  3 “Starships” deliver 3 landers loaded but unfueled. 

At 294 tons of propellant per lander,  and 100 tons per “Starship”,  some 9 “Starship” tanker flights would be required to fuel them fully up.  At 150 tons per SLS,  some 6 SLS flights would be required to fuel them up fully.  At about 60 tons per flight,  some 5 Falcon Heavy flights could be those tankers per lander,   for some 15 Falcon-Heavy flights to fuel the 3 landers up.  At 20 tons per flight,  it would require some 45 flights of Falcon-9 or Atlas-V to fuel the 3 landers in orbit.  The most cost-effective way to deliver these bulk liquid propellant supplies turns out to be 9 “Starship” flights,  with 15 Falcon-Heavy flights a rather close second.  If “Starship”,  the transfer crew need not be sent up separately.

Getting Earth Return and Lander Propellant Supplies to LEO and Docked

Remember,  we must send to Mars each lander loaded and fueled,  plus 1/3 of its Mars landing propellant supply,  plus 1/3 of the transport’s Earth return propellant supply.  These propellant supplies are pre-loaded tanks.  They are 1764.1 tons for the lander operations,  541.3 tons for the transport’s Mars departure,  and 455.9 tons for the transport’s Earth arrival.  That totals some 2761.3 metric tons of propellant,  which must be in tanks,  at about 95% propellant and 5% tank inert.

Unconstrained by other considerations,  I chose to break this up into nominal 60-ton loaded tanks.  The lander supply is 31 of these,  the Mars departure supply is 10 of these,  and the Earth arrival supply is 8 of these.  That’s a total of some 49 tanks to deliver to LEO,  at 60 metric tons each.  The most cost-effective way to do this was 49 flights of Falcon-Heavy,  flown expendably.

We will need a docking crew on-orbit for about a week max to assemble the docked cluster for each of the landers.  This can be a crew of 2 to 4 in a Crew Dragon atop a Falcon-9.  This probably will not happen in parallel for the 3 landers,  but serially.  So plan on 3 manned Falcon-9 launches to support these assemblies.

               Getting the Transport to LEO,  Loaded,  and Assembled

The orbit-to-orbit transport goes up as separate modules (without supplies) to be docked in orbit.  There are six 20-ton modules and one 40-ton center modules,  complete with solar wings that must unfold.  All the listed boosters could launch the 20-ton modules,  only Falcon-Heavy,  “Starship”,  or SLS could launch the 40-ton module.  The most cost-effective means was a tie:  2 flights of “Starship” or 3 flights (expendable) of Falcon-Heavy deliver these 7 modules to LEO. 

There is about 150 tons of supplies,  crew suits,  and crew personal equipment to deliver to the transport and load inside (152 exactly,  per these admittedly-uncertain estimates).   This is separable into lots deliverable by any of the boosters listed.  From a cost-effectiveness viewpoint,  this was another tie:  2 flights of “Starship”,  or 3 expendable flights of Falcon-Heavy. 

This is going to require a temporary docking and loading crew of perhaps 4 to 6 astronauts for a week or so in orbit.  If we send them up in two Crew Dragon capsules atop Falcon-9 boosters,  they can come home in one,  and leave the other Crew Dragon docked to the transport as one of its emergency return escape craft.  Add 2 Falcon-9 flights for the transport assembly crew unless “Starship” is used instead.

Getting the SEP “Tugs” to LEO and Fueled

The SEP “tug” hardware,  empty of the solid iodine fuel,  are not heavy at all.  This crude estimate says they are 14.42 tons each,  and there are 3 of them. That includes the folded solar panels,  the big thruster array,  the guidance package,  and the empty tank which doubles as the vehicle core structure,  about which dead-head payload gets docked.

Any of the listed boosters can get an empty tug to LEO.  The most cost-effective means is 3 Falcon-9 launches,  possibly flown recoverable,  but the expendable price was used here.

The iodine thruster fuel is a sublimable solid,  which can be sent up in portions that fit the various boosters,  determining the number of flights.  For the three tugs together,  we need 3213.54 metric tons of iodine sent to LEO.  (Most of this,  by far,  gets used sending payload to Mars.  Only a few tons with only 1 thruster firing is needed to return to Earth.)

Any of the listed boosters can do this job.  The most cost-effective means is by “Starship”,  with Falcon-Heavy a close second.  That would be 33 “Starship” flights,  or 54 Falcon-Heavy flights flown expendably. 

It will take a crew of 4-6 astronauts to load the iodine fuel and unfold the solar arrays,  plus some checkout.  We probably do not do all 3 vehicles in parallel,  but serially.  If by “Starship”,  that vehicle can carry the crew.  If by Falcon-Heavy,  a separate Falcon-9 launch is needed to send this crew up for a week or two in orbit as the payloads arrive,  which is a huge Falcon-Heavy flight rate!  “Starship” with payload and loading crew aboard is thus the preferred way,  by far.

Getting the Earth Departure Stage to LEO and Fueled

This is assumed an empty stage delivered as one piece of hardware at 47 metric tons,  plus 796.2 metric tons of LOX-LH2 propellants delivered as bulk liquid.  Bulk liquids can be delivered in multiple payloads by any of the listed boosters,  but requires special tankage and a human crew to do the transfers. 

The most cost effective way to deliver the empty stage is by a single Falcon-Heavy,  possibly flown recoverably,  but priced expendably for this analysis.

The most cost-effective means to deliver bulk propellant is 8 “Starship” flights,  followed fairly closely by 14 Falcon-Heavy flights.  These require crews,  which can be aboard the “Starship” flights.  They would have to come up in some 14 Falcon-9 launches with Crew Dragon if Falcon-Heavies were the propellant ferries.  By far,  the preferred approach is 8 crewed “Starship” flights.

Getting the Crew Onto the Transport for the Mission

The Mars mission crew is only 6 people.  This is one Falcon-9 Crew Dragon flight to send them up.  Their Crew Dragon docks with the transport to be its second (and redundant) emergency escape capsule. If not covered earlier,  make this 2 flights so there are two Crew Dragons as escape capsules.

Totaling Up Mission Launch Requirements & Guessing Costs

I totaled-up the launch costs for this mission.  On the assumption that launch costs are 20% of overall program costs,  that puts this mission in a rather modest cost category,  despite the large tonnages.  That is precisely because it does NOT use SLS to launch anything,  at a billion dollars per flight (if not more)!  See Figure 26 for a summary of the launch requirements and costs.  The basis for comparison is the infamous “90 Day Report”,  based on mounting essentially “Apollo-on-steroids-plus” as executed by the long-favored contractors,  to send a crew of 4-to-6 to one site on Mars,  in the one trip. 

Totaling Up What the Mission Accomplishes

This makes the comparison to the “90 Day Report” even more stark.This mission as planned has a “way out” or a self-rescue capability at every step,plus inherently designed-in artificial gravity and radiation protection (to include solar flare events). The likelihood of this crew returning alive and healthy is actually quite high.In comparison,with the “90 Day Report” mission,that likelihood is rather low,because it does not offer those characteristics.

What this mission accomplishes is up to 6 different sites explored in the one manned trip to Mars.With the “90 Day Report” mission design,only one site gets explored.

This mission leaves considerable usable assets at Mars for future missions to utilize.That would include the reusable landers,either in low Mars orbit,or on the surface if landed robotically.Plus, there might be some leftover propellant,probably in Mars orbit.The “90 Day Report” mission leaves few (if any) usable assets on Mars for future missions to utilize:maybe a surface habitat structure and a rover or two,and possibly a nuclear power supply item.

See Figure 27 for a listing of what this mission accomplishes,compared to that of the “90 Day Report”.

“Bang-for-the-Buck” Discussion

The first gross indicator is program cost for the one trip to Mars,  divided by the number of sites explored while the mission is there.  For my mission design,  cost per site ranges from $11.7B/site to at most $70.3B/site,  depending upon whether the minimum 1 or maximum 6 sites get explored.  That is factor 6.4 to 38.5 times better cost per site than that of the “90 Day Report”. 

The second gross indicator is the likelihood of getting the crew back alive and healthy.  Because of the features demanded by ethics,  and designed-in from the start,  this mission plan can truthfully claim a high likelihood of accomplishing this.   The “90 Day Report” mission plan cannot truthfully claim that. 

For one thing,  there is no rescue for a crew stranded on Mars.  For another,  there is a high likelihood of a solar flare event during a 31 month mission,  and almost zero chance of surviving that event with no radiation shelter.  And yet another:  there are two 9-month transits in zero-gee,  separated by a 13 month stay on 38% gee Mars,  with undetermined therapeutic effect,  if any.  Should an emergency free return at Earth arrival be required,  that is a high-gee event (likely 10+ gees).  A crew weakened by microgravity diseases is unlikely to survive this.

Now remember,  spaceflight history clearly demonstrates that there is nothing as expensive (economically and politically) as a dead crew.  Especially one dead from a bad management decision.  My mission design raises crew survival probability,  the “90 Day Report” mission design does not;  that survival probability is quite low,  if one is truthful about it.

In order to get both high “bang for the buck” and a high likelihood of getting a crew back healthy,  I had to think way outside the usual boxes.  One of those boxes is “nothing can look much different than what we already did during Apollo,  shuttle,  and ISS”.  Another is “no mission can be affordable if there must be a high tonnage launched”.  A third is “you simply must do direct entry at Mars to save launched tonnage”.  A fourth is “you must use SLS no matter what in order to launch this mission”.

All proved to be false constraints on thinking.  The only one that is true is the one I used:  crew survivability above all,  driven by basic ethics.  In a nutshell:  “provide a way out or a self-rescue capability at every single step”.  That drove me to orbital-based exploration and a manned orbit-to-orbit transport design.

The main possible weakness of my mission design is the low payload fraction of my one-stage reusable landers:  around 2%.  A one-shot two-stage design would have a far higher payload fraction (perhaps 5-6% if you include the safety-required abort capsule,  only higher if you fail this safety requirement),  resulting in a smaller mass sent to Mars for each lander.  But I would have to send more of them (8) to maintain a rescue capability and a spare,  and still visit as many as 6 sites.  This I leave to others to explore.

Final Comments

In terms of both cost and safety,  the comparison of this mission plan to that of the “90 Day Report” demonstrates the unattractiveness of the usual way NASA did things in the past.  There is far more “bang for the buck” and an enormously-higher probability of getting the crew back alive and healthy in my plan.   Not only that,  my program cost is far,  far lower.

The astute reader will observe that I have selected a lot of Spacex hardware as the most cost-effective means to launch and assemble this mission.  That begs a comparison to the Spacex plan just to send multiple “Starships” to Mars by direct entry from the interplanetary trajectory.  According to the presentations released,  that would be about 6 “Starships” initially landed on Mars,  with probably one or at most two of them eventually returning to Earth,  if the local propellant production works,  and it can supply them fully and quickly enough. 

It is as yet unclear whether 5 or 6 “Starship” tanker flights are required for refueling each interplanetary “Starship” in LEO for the journey to Mars.  So somewhere between 36 and 42 total “Starship” launches are required to support their mission.  Using $150M per launch,  and launch costs equal 20% of program cost,  that’s $5.4-6.3B launch cost,  and $27.0-31.5B program cost,  to put their mission onto Mars. 

That program cost scaleup is real,  even for them,  because they are counting on others to supply the local propellant production hardware,  local rover vehicle capabilities,  and local life support capabilities (cannot live in the landed “Starship” forever !!),  not to mention local electric power.  They have their hands full just developing the “”Starship” vehicle.

That’s comparable to my costs,  and (like me) way below the costs in the “90 Day Report”.  The differences are many,  however.  They explore only 1 site,  period.  If the local propellant production fails to meet expectations,  nobody comes home.  They say they will supply radiation sheltering,  but not artificial gravity.  They are counting on Mars’s 0.38 gee being “therapeutic enough”,  when in point of fact,  nobody yet knows that to be true. My mission takes none of those risks and explores up to 6 sites.

There is no aborting or bailing-out during the “Starship” direct entry at Mars.   There is no aborting or bailing out during the landing and touchdown.  They have yet to address soil bearing loads vs landing pad size for Mars,  or rough field landing hazards such as slope,  local roughness   and big boulders.  There is no bailout or abort during the return ascent.  There is no bailout or abort for the direct entry at Earth return.  There is no bailout or abort during the Earth landing and touchdown.    A failure during any one of these events is inevitably a loss of the vehicle and everybody aboard.  My mission takes none of those risks.

Yeah,  you can save the money using “Starship” as the transit vehicle (by about a factor of 2-3 over my plan).  But you are also very much more likely to kill one of your crews if you do (which also very likely would put a stop to the ongoing mission). 

Ethics-driven spaceflight design “from the get-go” seems the more prudent course,  especially when you consider the consequences of killing a crew. 

References

#1. NASA radiation website http://srag.jsc.nasa.gov/Publications/TM104782/techmemo.htm, titled Spaceflight Radiation Health Program at JSC (no cited reference newer than 1992).

#2. From Bigelow Aerospace website http://www.bigelowaerospace.com/b330/  as of 3-7-17

General Dynamics EMPIRE

General Dynamics EMPIRE
Common
Specific Impulse800 s
Crew8
C-22
Height106 m
Diameter21 m
Thrust1,780,000 N
Wet Mass900,000 kg
C-23
Height95 m
Diameter23 m
Thrust1,780,000 N
Wet Mass960,000 kg
C-26 (Duplex)
Height162 m
Diameter10 m
Thrust1,780,000 N
Wet Mass? kg
C-28
Height174 m
Diameter15 m
Thrust1,780,000 N
Wet Mass? kg
C-28V
Height174 m
Diameter
(Saturn V)
10 m
Thrust1,780,000 N
Wet Mass? kg

Information for this entry are from The Empire Dual Planet Flyby Mission by Franklin Dixon, Proceeding of the Symposium on Manned Planetary Missions 1963/1964 Status (starting at page 11), EMPIRE: Background and Initial Dual-Planet Mission Studies by Fred Ordway et al. and the entry in Astronautix.

Unfortunately the real juicy details are hidden in reports that are still classified, due to the details about the nuclear engines. Pity.

Back in 1962, NASA's Marshall Space Flight Center's Future Projects Office (FPO) decided to get serious about manned exploration of other planets. They commissioned a study with the contrived name Early Manned Planetary-Interplanetary Roundtrip Expeditions (EMPIRE). Three mission study contracts were awarded. General Dynamics would study Mars orbital missions. Lockheed would study Mars flyby and orbital missions. And Aeronutronic would study Mars-Venus flybys.

Of the three reports the one from General Dynamics was larger than the other two put together. This is because Krafft Ehricke was a huge fan of expeditions to Mars.

The report focused on the favorable Terra-Mars relationship during the 1973-1975 period. They figured a round trip mission period between 400 and 450 days, with a planetary capture time between 30 to 50 days (ample time for a Mars landing segment).

The report analyzed about thirty difference configurations, they figured that a Terra departure mass of between 1,200 to 1,400 metric tons was realistic for a crew of eight with a mission payload of 45,000 kg. There would be two vehicles traveling as a convoy, one for crew and one for cargo. Plus spares and back-ups. The crew ship would have the Earth Entry Module (EEM) and a space taxi to commute between the two vessels and to tow cargo. And the cargo ship would have a spare EEM. Just In Case.

Crew Ship Tasks
Crew Transport, Navigation, Data Processing and storage, Communication, Control of auxiliary craft, Injection of EEM into proper Earth return trajectory
Cargo Ship Tasks
Transport of auxiliary craft, Transport of spares, Transport of make-up fuel, Transport of spare EEM, Navigational assistance, Back-up crew vehicle

Out of the many possibilities for the life-support section, option L-42 was chosen.

Mainly because they went with a tumbling pigeon artificial gravity instead of a centrifuge. Tumbling pigeon works best with elongated spacecraft. And like most spacecraft with nuclear engines, the design is elongated in order to keep the habitat module far away from the radioactive engines.

The secondary reason is this configuration is narrower, so it fits into a smaller shadow cast by a smaller anti-radiation shadow shield. Shadow shields are heavy and really cut into the payload mass budget. The smaller the shield, the bigger the allowed payload mass.

Tumbling pigeon works best when the habitat module has deck but is very wide ("horizontal LSS"). The advantage is the gravity is constant along the deck. Alas, the width would have the ends of the habitat module protruding out of the anti-radiation shadow and into the deadly radiation.

Therefore, instead the habitat module has multiple decks and is very narrow ("radial LSS"). The advantage is that all of the habitat is safely inside the anti-radiation shadow. The disadvantage is that there is a different level of gravity on each deck.


EEM (Earth Entry Module)

Can accomodate entire crew of 8. Basically an enlarged Apollo command module. This is attached to the fore end of the spacecraft.

If I am reading the flight plan correctly, after Mars exploration is finished, the life support system is jettisoned and the EEM attaches in its place. Then the trans-Terra burn is initiated. Which means the crew has to live in a space barely the size of four telephone booths for the half-year trip home. Personally I think long before they get home the crew will snap and murder each other.


COMMAND MODULE

10 meters in diameter. Upper deck is Command Station with three crew workstations. Lower deck is sleeping quarters for five crew. Encased in radiation shielding so module is also the storm cellar. Bottom of sleeping quarters has the docking port for the EEM. Top of the command station joins to the rest of the habitat module.

The shielding on the storm cellar was initially designed to be plain old water, but it proved to be impossible to carry enough to protect the crew from a largish solar storm ("protect" defined as "reduce crew radiation dose to 0.01 Gray per day"). So the water shielding was supplemented with either solid borated polyethylene or liquid monomethyl hydrazine (MMH). They went with MMH because [1] liquid is easier to jettison in case of emergency and [2] if you carry along some oxygen difluoride (OF2) to use as an oxidizer, it and MMH make a kick-ass hypergolic fuel (Isp = 405 secs). This can be used in the auxiliary craft carried as cargo. The advantage is you are making the mass do double duty: as radiation shielding and as fuel. Because every gram counts.

It is a pity that OF2 is so damn corrosive that it will even oxidize the noble gas xenon. This makes it difficult to make a tank that the stuff won't dissolve a hole in.


HABITAT MODULE L-22

Life Support Section L-42 has a mass of about 39,700 kg. The inside volume varies from 56 kg/m3 to 80 kg/m3. The floor area is about 180 kg/m2. With an 8 crew complement the ecological mass is 2,150 kg/crew

In this diagram, Fore of the ship is to the bottom, Aft is to the top (where the propulsion system resides). Arrangement is because the spacecraft uses the tumbling pigeon approach to artificial gravity. Spin provides 0.25 g.

The central part of the habitat module holds the EEM, Command Module, and Internal Mission Module. The IMM contains [A] Life Support, [B] Food storage, and [C] Repair shop. In addition there are four external mission modules attached around the center. The external module contents depend upon the specific mission.

The spine at top is 23 meters long, to keep the habitat module far away from the radioactive SNAP-8 nuclear power generator and the nuclear engines.

The two design philosophies for the habitat module were Modular and Integrated. Modular takes more mass, but Integrated does not allow one to jettison bits of of the habitat in case of emergency. Integrated also does not allow one to upgrade an existing habitat by swapping out old modules with more modern versions. The designers went with Modular.

Class-22
     Propellant tank clusters, each with a large central tank surrounded by smaller diameter tanks.
     Each major maneuver has its own dedicated pair of engine. Tanks and engines are staged (jettisoned) after performing their maneuver. This is because these early-model engines only had an operating life of 1 hour, not long enough to peform all the manuevers.
Class-23
     Propellant tank clusters, each with a central tank surrounded by tanks with equal diameters.
     With the exception of the initial Terra departure manuever, there is just one pair of reusable engines. Just the tanks are jettisoned after performing their maneuver, the engines are retained and reused. This is because these are new and improved engines with a much longer operating life.
Class-28
     Single tank in tandem for all propellant modules. Tank diameter 15.2 meters
Class-28V
     Same as C-28 except tank diameter is only 10 meters, so as to be compatible with the Saturn V launch vehicle.

General Dynamics EMPIRE Lander

EMPIRE Lander
EngineChemical
FuelUDMH
OxidizerChlorine Trifluoride
Isp280 s
325 s
(300 s)
Area of drag annulus836 m2
Mars Return Orbit1,070 km
Mass Schedule
Fuel Mass1,630 kg
Oxidizer Mass4,940 kg
Descent Mass9,980 kg
Ascent Mass8,440 kg
Return Payload1,360 kg
Mass left on Mars1,540 kg

This is from US Spacecraft Projects #01 and History of Rocketry and Astsronautics ASS History Series, Vol 19, Chapter 1 EMPIRE: Background and Initial Dual-Planet Mission Studies.

It really would not be worth the trip to travel to Mars and not make a crewed landing. General Dynamics didn't spend lots of design time on an excursion vehicle, but they did bang out one that had either a crewed or uncrewed nose section. Both of them returned the top section, loaded with Mars surface samples. The crewed version transported two crew for a seven day surface stay.

The lander uses a chemical rocket fueled by storable hypergolic fuels. Alarmingly they chose the touchy unsymmetrical dimethylhydrazine as fuel, and the insanely dangerous chlorine trifluoride as the oxidizer.

CHLORINE TRIFLUORIDE

It is, of course, extremely toxic, but that's the least of the problem. It is hypergolic with every known fuel, and so rapidly hypergolic that no ignition delay has ever been measured.

It is also hypergolic with such things as cloth, wood, and test engineers, not to mention asbestos, sand, and water—with which it reacts explosively.

It can be kept in some of the ordinary structural metals—steel, copper, aluminum, etc.—because of the formation of a thin film of insoluble metal fluoride that protects the bulk of the metal, just as the invisible coat of oxide on aluminum keeps it from burning up in the atmosphere. If, however, this coat is melted or scrubbed off, and has no chance to reform, the operator is confronted with the problem of coping with a metal-fluorine fire. For dealing with this situation, I have always recommended a good pair of running shoes.

From IGNITION! AN INFORMAL HISTORY OF LIQUID ROCKET PROPELLANTS by John Clark (1972)

But there was more sorrow lurking in the lander design. As with most lander designs meant for those few planets with atmospheres, designers cannot resist the temptation to use aerobraking. Because every gram of landing propellant replaced by aerobraking is an extra gram of payload. So the EMPIRE lander came equipped with a huge drag annulus (read: parachute) with a surface area of 836 m2. Between the parachute and the rocket engines, the lander would be gently delivered to the Martian surface.

Except for one teeny-tiny little flaw.

You see, in 1962 when the lander was designed, scientists were under the misapprehension that the Martian atmosphere was about 25% as dense as at Terra sea level. Oh, calamity and woe! Turns out it is actually 0.7% as dense, pretty blasted close to being a vacuum. The parachute would do diddly-squat to slow down the lander, which would auger into the Martian ground at high velocity and explode into a spectactular metal-fluorine fire.

Lift-off back to the mother ship, on the other hand, would actually work. Small solid-fuel rockets on the nose tower would fly the ship high enough so that the rocket exhaust from the main engine would be diluted when it touched the ground. Otherwise it would hurl sand and rock upward and endanger the ship. The ship would have a hull about as strong as tin-foil, just like the Apollo lunar module.

The manned version of the lander also has solid-fuel abort rockets on the nose tower. I'm not sure I understand their function. In case of an abort, they are supposed to drag the lander to a safe place, where exactly?

GCNR Spacecraft

RocketCat sez

You want an atomic rocket? I'll give you an atomic rocket!

Yeah, yeah, this ain't an over-the-top torchship like an Orion Drive ship much less Zubrin's outrageous Nuclear Salt Water Drive. But it is a good working-man's atomic rocket that has the horsepower to Get The Job Done. Orion drives are for battleships, this one is a space trucker hauling cargo.

Bloated chemical drives can barely do a Mars mission in two years, this little atomic number can smoke the mission in 80 days flat! I know that saying the exhaust is radioactive is putting it mildly, but nobody is near enough to it to be harmed (well, except for the poor working-class slobs who are the ship's crew).

Bottom line:

  • It is undisputably an Atomic Rocket

  • It has both high thrust and high specific impulse, approaching torchship levels

  • The design does a clever end-run around the "melting reactor" problem with a solution both elegant and brute force

  • 80 day round-trip to Mars, man! How cool is that?

GCNR Spacecraft
PropulsionNTR-GAS/open
Fueluranium-235
Propellanthydrogen
Specific Impulse2,500 to 6,500 s
Exhaust Velocity24,500 to
63,800 m/s
Mass Flow0.8 to 6.7 kg/s
Thrust20,000 to
430,000 N
Fixed Thrust224,000 N
Thrust Power0.25 to 13.7 GW
Initial Accel0.01 to 0.05g
GCNR Spacecraft
Mars Courier
Mission
Duration
80 days
Wet Mass950,000 kg
Dry Mass290,000 kg
Mass Ratio3.28
Thrust150,000 N
Initial Accel0.016 g
Specific Impulse5,500 s
Exhaust Velocity53,955 m/s
ΔV64,100 m/s
H / 235U Ratio200:1
235U Fuel3,300 kg
Hydrogen
Propellant
660,000 kg

Data from Gas Core Rocket Reactors - A New Look.

This little hot-rod can do a round-trip mission to Mars in 80 days flat! That's only 2.7 months. Using Hohmann trajectories a round-trip Mars mission will take 32.3 months (2.7 years) when you take into account the wait for the Mars-Terra launch window to open.

The report starts off with the common complaint that most rocket propulsion is either high-thrust + low-specific-impulse or vice versa. The problem being that rocket designers want a high-thrust + high-specific-impulse engine. In other words they want a torchship.

The closest thing they can find that is actually feasible is a Gas-Core Nuclear Thermal Rocket. Open-cycle of course, closed-cycle has only half the exhaust velocity. So what if it spews still-fissioning uranium in an exhaust plume of glowing radioactive death?

The report examines the GCNTR's performance to see if it is a torch drive. It comes pretty close, actually.


The higher the specific impulse / exhaust velocity, the more waste heat the engine is going to deal with. They figure that a GCNTR can control waste heat with standard garden-variety regenerative cooling like any chemical rocket, but only up to a maximum of 3,000 seconds of specific impulse. Past that you are forced to install a dedicated heat radiator to prevent the engine from vaporizing. Otherwise the engine vaporizes, your spacecraft has no engine, and perhaps centuries from now your ship will come close enough so that space archaeologists can recover your mummified remains.

As everybody knows, thermal rockets use a heat source to heat the propellant (usually hydrogen) so that its frantic jetting through the exhaust nozzle creates thrust. Solid-core nuclear thermal rockets (NTR) use solid nuclear reactors. They are limited to a specific impulse (Isp) of about 825 seconds, since that corresponds to a propellant temperature of about 2,500 K. Any higher specific impulse raises the temperature high enough that the reactor starts to melt. And nobody likes an impromptu impression of the China Syndrome. If you want an Isp of 5,000 seconds you are talking about a propellant temperature of 22,000 K!

Also as everyone knows the gas-core NTR concept is the result of clever engineers thinking outside of the box and asking the question what if the reactor was already vaporized?

Instead of solid nuclear fuel elements it uses a super hot ball of uranium vapor which is dense enough and surrounded with enough moderator (neutron reflector) that it still undergoes nuclear fission. The fission produces huges amounts of thermal radiation, which heats the hydrogen propellant. The fissioning uranium is like a nuclear "sun" in the center of the engine. The reaction chamber directs a flow of propellant around the sun to be heated.

Since this is using the concentrated energy of fission there is no real limit to the thermal energy generated (think nuclear weapons). Unfortunately there is a limit to the hydrogen propellant's ability to absorb heat. Any heat that the hydrogen fails to sop up will hit the engine walls. If this unabsorbed heat is more than the heat radiator can cope with, bye-bye engine. This puts the upper limit on the engine's Isp capability.

Engine
Cavity Linergraphite +
5% niobium
Moderatorberyllium oxide
Propellant
Presure
5.07×107 to
20.34×107 N/m2
Propellant
Seeding
10% by weight
Moderator
Thickness
0.46 m
Cavity Liner
Thickness
0.0063 m
Engine Cavity
Diameter
2.44 m
Uranium
Plasma Dia
1.8 m
Uranium
Plasma Vol
3.04 m3
Uranium
Plasma
Critical Mass
21 kg
Engine Mass
(including 235U)
40,000 to
210,000 kg

The engine is spherical. The outer layer is the pressure vessel (since both the propellant and uranium gas needs lots of pressure to make this thing work), a layer of beryllium oxide (BeO) moderator (a neutron reflector to help the uranium undergo nuclear fission), and an inner porous slotted cavity liner that injects the cold propellant to be heated. In the center is the furious blue-hot atomic vortex of uranium plasma.

Sadly, this structure does suffer from waste heat:

[1] a bit under 0.5% of the reactor power gets to the slotted cavity liner from thermal radiation emitted by the hot propellant. Which is a problem but not a major one. Most of the thermal radiation is soaked up and removed by the propellant.

[2] A whopping 7% of the reactor power hits all three layers of the engine, because part of the fission output is in the form of gamma-rays and neutrons, instead of useful thermal radiation. Hydrogen propellant does not do zippity-doo-dah to soak up gammas and neutrons, all of it sails right through the propellant to hit the engine structure. Deep inside the engine structure, gamma-rays and neutrons are more penetrating than x-rays.

This waste heat is managed by the engine heat radiator (and a bit managed by regenerative cooling, about as effectively as a 3-year-old helping Daddy wash the car). Most of the engine is the beryllium oxide moderator. It is designed to operate at 1,400 K, which is below the 1,700 K melting point of the BeO but above the 1,100 K radiator temperature (otherwise the radiator will refuse to remove the heat).


The hydrogen propellant is pumped into the engine at about 5.07×107 to 20.34×107 newtons per square meter (which is why the engine needs a pressure vessel).

As it turns out hydrogen propellant is transparent, which means it is lousy at absorbing thermal radiation. That's not good. To remedy this sad state of affairs, it is "seeded" by adding tiny metal bits about the size of particles of smoke, about 5% to 10% seeding material by weight. This is done right before the propellant exits the porous cavity liner into the flood of heat from the nuclear vortex. The seeding absorbs all the thermal radiation and passes the heat to the propellant by conduction. The seeding material will be something like graphite, tungsten, or non-fissionable uranium 238.

Around the exhaust nozzle the seeding concentration will have to be increased to 20% to protect the nozzle from propellant heat. The cold 20% seeded hydrogen will reduce the specific impulse a bit but it has to be done.

The porous cavity liner (in some as yet to be defined manner) magically sets up flow patterns so that the propellant flows around the hot uranium and exits via the exhaust nozzle. Meanwhile miraculously the uranium is trapped in a stagnant cavity in the center so hideously radioactive fissioning uranium does not escape through said exhaust nozzle. Uranium escape not only exposes the crew to deadly radiation, it is also a criminal waste of uranium (that is, it lets get away uranium that is not contributing to the engine's thrust).

The interior of the engine (cavity diameter) is 2.44 meters in diameter (7.61 cubic meters), and the incandescent ball of violently fissioning uranium is planned to have a diameter of 1.80 meters and a volume of 3.04 cubic meters. This gives a fuel-to-cavity radius ratio of 0.74. The idea is for the uranium sphere to be 40% of the volume of the entire chamber. However since hydrogen propellant is going to diffuse into the atomic vortex, the uranium sphere might be up to 50% hydrogen. This means the effective volume of pure uranium will be closer to 20% to 30% of the entire chamber.


The uranium can be injected by pushing a very thin rod of solid uranium into the chamber. The uranium penetrates the BeO moderator inside a tunnel lined with a cadmium oxide neutron poison, because otherwise there would be a nuclear explosion once the uranium was surrounded by BeO. This is a bad thing. The engine was designed to have the nuclear reaction happen in the core of the chamber, not in the walls.

As the uranium rod enters the chamber, the heat of the fission ball vaporizes the rod so the fresh uranium atoms can join the party.

A problem is how to get the process started. At startup, there ain't no ball of fissioning uranium to heat up the rod. The report says that the engine will have to be started by first blowing in some hydrogen and somehow injecting some powered uranium metal into the stagnant cavity until it reaches critical mass. Sounds tricky to me.


Figures 2a through 2c above are for a reactor of the following characteristics:

  • Spherical geometry
  • Uranium-235 fuel
  • Beryllium-oxide (BeO) moderator
  • Fuel-to-cavity radius ratio 0.67
  • Cavity liner thickness 0.63 cm
  • Cavity liner graphite + 5% niobium

Figure 2a shows that the 235U critical mass ranges from 10 to 35 kilograms for the cavity diameters and moderator thicknesses considered (all the curved lines are more or less above the 10 kg line and below the 35 kg line). Now for a given cavity diameter, you can reduce the critical mass required by adding more BeO neutron reflector. This means the pressure inside the engine can be lowered, which means the mass of the pressure shell can be lowered. Alas the increased penalty mass of the BeO moderator more than offsets the mass saving on the pressure shell.

(Elsewhere the report notes an optimal thickness of BeO to be 0.46 m, and a cavity diameter of 2.44 m. Eyeballing the graph implies a critial mass of 21 kg of 235U.)

Figure 2b shows that if the BeO moderator thickness is fixed, increasing the cavity diameter will decrease the critical density (the curved line will be closer to the bottom of the graph). Not shown in the table is the unfortunate fact that increasing the cavity diameter also has the side effect of increasing the total BeO weight.

(Elsewhere the report notes an optimal thickness of BeO to be 0.46 m, and a cavity diameter of 2.44 m. Eyeballing the graph implies a critial density of 18 kg/m3. If the uranium plasma ball has a volume of 3.04 m3, at that density it will contain about 55 kg of uranium, which is more than the 21 kg (from eyeball value above) it needs for criticality. However, since propellant seepage will make the sphere about 50% hydrogen, this means it will have about half of 55 kg. Which is a reasonably close eyeball value to a second eyeball value. I'm just playing number games with the graphs, do not put too much credence to these speculations on my part.)

Figure 2c shows that there is an optimum BeO moderator thickness which gives a minimum critical density for a given BeO moderator weight.

Why is there an optimum BeO moderator thickness?

If the BeO is too thin there is excessive neutron leakage (the purpose of the BeO moderator is to reflect escaping neutrons back into the fissioning uranium, basically kicking the out-of-bounds neutrons back into play). Excessive neutron leakage means the blasted cavity diameter will have to be extremely large to avoid very high critical densities.

If the BeO is too thick, the total BeO weight becomes very large. Even though you can get away with smaller cavity diameters without the heartbreak of very high critical densities.

Figure 2c is telling you that the optimum BeO thickness is 0.46 meters (for a reactor of the specified characteristics). 2c goes on to tell you that above a moderator weight of 40,000 kg larger cavities only give a slight reduction in the critical density (the curved lines are almost horizontal).

So all the engine weight estimates below are assuming a BeO thickness of 0.46 meters.


Experiments show that an effective fuel volume is about 20% to 30% of the cavity volume, for a uranium flow rate less than 1% of the hydrogen flow rate.

The paper assumes the engine can accelerate at about 0.01 to 0.05g (0.098 to 0.491 m/s)


The idea is to get the maximum thermal radiation from the fissioning atomic fireball into the cold hydrogen propellant, and the minimum thermal energy escaping the hot hydrogen propellant (which reduces the specific impulse and scorches the heck out of the cavity wall).

Figure 5b shows experimental data for tungsten-seeded hot hydrogen. It says that adding just a few percent by weight of tungsten will increase the thermal absorption cross section to between 2,000 to 100,000 square centimeters per gram. The figure also shows the thermal absorption increase at elevated pressure, which is a good thing since the engine is a high-pressure rig.

These cross sections are high enough to protect the cavity wall from damage for Isp from 4,000 to 7,000 seconds.


Figure 6 is the straight dope on the gas-core NTR engine parameters. The critical density of uranium given cavity size and moderator is as per figure 2. Thermal absorption of seeded hydrogen is as per figure 5. Heat tranfer analysis is used to determine maximum specific impulse that will keep heat load on cavity wall below 1,000 K. Engine pressure is whatever is required to have a critical mass of uranium.

The engine weight is assumed to be the sum of the three major components: BeO Moderator, Pressure Shell, and Heat Radiator. Plus 4,000 kg or less for the uranium fuel.

Pressure Shell assumes a strength-to-density value of 1.7×105 N-meters/kg.

Heat Radiator assumes a unit weight of 140 kilogrmas per megawatt of radiated power. Heat depostion rate is assumed to be 7% of reactor power. Heat radiator operates at 1,100 K (instead of 945 K), which reduces the required radiator surface area by a factor of 2. This kind of radiator more than doubled the specific impulse without adding enough weight to offset the gain. Future radiator designs with even lower unit weights would give even more specific impulse gains.

  • Spherical geometry
  • Uranium-235 fuel
  • Beryllium-oxide (BeO) moderator
  • Beryllium-oxide (BeO) thickness 0.46 meters
  • Fuel-to-cavity radius ratio 0.67
  • Fuel volume 30% of cavity volume
  • Uranium loss rate is 1% or less of hydrogen flow rate
  • Cavity liner thickness 0.63 cm
  • Cavity liner graphite + 5% niobium

In figure 6, the abscissa for both charts is engine thrust. The charts are for thrust levels from 20,000 to over 400,000 Newtons.

The ordinate of the upper chart (Fig 6A) is specific impulse, engine weight for lower chart (Fig 6B). Specific impulse ranges from 2,500 to 6,500 seconds. Engine weight ranges from 40,000 to 210,000 kg.

The curved lines are Engine Pressure, for ranges between 0.5×108 to 2.0×108 N/m2. Note in Fig 6A the three curves are labeled "Low", "Nominal", and "High". These labels are used in the Mission Chart below.

A higher engine pressure allows higher specific impulse because higher pressure makes the hydrogen propellant more opaque. But higher pressure also makes the engine heavier.

Higher thrust increases the specific impulse because there is more propellant flow to cool the cavity wall (note this is the opposite of what occurs when shifting gears). But this also makes the engine heavier.

The two reason above are why it is impossible to chose the "best" engine. What you have to do is specify a specific mission in order to have enough determining factors to figure which engine would be best.


The spacecraft is composed of a gas-core engine (with heat radiator and uranium fuel), a command module, payload, various jettisonable liquid hydrogen propellant tanks , and interconnecting structure.

The engine provides four burns:

  1. Terra orbit escape/target planet trajectory insertion
  2. Target planet orbital capture
  3. Target planet orbit escape/Terra trajectory insertion
  4. Terra orbital capture

After each burn the associated empty propellant tanks are jettisoned, except for the last burn. This is because the command module is attached to the last tank, and the crew would object strongly to being cast off into deep space. The command module also relies upon the hydrogen in the last tank for extra engine-radiation shielding.

Initial Mass In Orbit
ItemMass
Command Module50,000 kg
Payload to Planet150,000 kg Science/Exploration
0 kg Courier
Expendables50 kg/day
Propellant Tankage20% of hydrogen mass
Interstage Structure2% of transmitted load
Thrust Frame5% of thrust
Gas-core Engineas per Figure 6, including uranium storage and supply
Parking orbits600 km circular at Terra
high ellipse at target planet
Propulsive Effortideal ΔV from ref. 19
gravity-loss corrections Cg from ref. 20
Propellant Fraction1 - exp(-((ΔVi * Cg) / (Isp * g0)))
Ref. 19. Fishbach, L. H., Giventer, L. L., and Willis, E. A., Jr., "Approximate Trajectory Data for Missions to the Major Planets," TN D-6141, 1971, NASA, Cleveland, Ohio.
Ref. 20. Willis , E. A., Jr., "Finite Thrust Escape from and Capture into Circular Elliptic Orbits," TN D-3606, 1966, NASA, Cleveland, Ohio.

Propellant Fraction equation comes from combining these four equations into one big equation:

Pf = 1 - (1/R)
R = ev/Ve)
1/ex = e-x
Ve = Isp * g0

where:

Cg = gravity-loss corrections Cg from ref. 20
Δv = ship's total deltaV capability (m/s)
ex = antilog base e or inverse of natural logarithm of x.
g0 = acceleration due to gravity = 9.81 (m/s2)
Isp = specific impulse (seconds)
Pf = propellant fraction, that is, percent of total rocket mass M that is propellant: 1.0 = 100% , 0.25 = 25%, etc.
R = mass ratio (dimensionless number)
Ve = exhaust velocity (m/s)

In the charts below

  • SCNR: Solid-Core Nuclear Rocket (an old-fashioned NERVA)
  • REGEN GCNR: Gas-Core Nuclear Rocket cooled with Regenerative Cooling (choked down to avoid need for heat radiators)
  • RAD GCNR: Gas-Core Nuclear Rocket cooled with Heat Radiator (uses heat radiators so it can run full-bore)
  • FUSION: Fusion Rocket (for comparison purposes)
  • SCIENCE/EXPLORATION: A mission where you bring along tons of scientific payload, and stay on Mars for 40 days to do some science.
  • COURIER: A mission with no payload just a Very Important Person. And no staytime on Mars, just a quick unloading/loading and immediate return to Terra.

In many of the charts Initial Mass in Earth Orbit (IMEO) is used to measure efficiency. The lower the IMEO value, the more efficient. Usually because it means lower propellant requirments, and may allow more payload.

Figure 6 shows that the radiator-cooled gas-core nuclear rocket becomes more efficient (higher Isp and lower specific weight) as the thrust level is raised. So the GCNR is best for missions with large payloads and/or big thrust-to-weight requirements. The missions depicted in the charts below were chosen with this in mind.


This chart shows the effect of changing the duration of the mission on the Initial Mass in Earth Orbit (IMEO). You want IMEO to be as low as possible. The shorter the mission duration, the more propellant you have to pack to increase ΔV, so the higher IMEO becomes. Obviously you can lower IMEO by increasing the mission time, but who wants to spend years on a Mars mission?

The scientific missions assume a 40 day stay on Mars to do science stuff.

The patheticaly weak SCNR (NERVA style solid-core nuclear rocket, shown with yellow curved line) has minimum mass at around 500 days and 1.5×106 kg IMEO (very roughly). This wimp ain't gonna manage a trip time below 400 days, not with a practical IMEO it isn't.

The first gas-core nuclear rocket (green curved line) show an immediate performance improvement. This is the gas-core with no heat radiator, deliberately throttled down so it can make do with mere regenerative cooling. If it is given the SCNR's 1.5×106 kg IMEO, it can do the mission in half the time, only 250 days. Its lowest IMEO is about 0.7×106 kg (700 metric tons) with a mission time around 480 days.

But the other gas-core rocket is even more powerful.

The gas-core nuclear rocket with a heat radiator (blue curved line) lowest IMEO is 0.4×106 kg (450 metric tons). This is only twice the payload (150 tonne payload + 300 tonnes = 450 tonnes). If it is loaded at a IMEO of 0.7×106 kg (the regenerative GCNR's minimum) it will do the mission in 250 days flat instead of 480 days.

With performance this high, the 40 day stay on Mars becomes an appreciable fraction of the total mission time. However low transit times mean high ΔVs and high propellant fractions.

So we now present "courier mode." This has a zero day stay on Mars, instead it immediately turns around to return to Terra. No payload either, except for something way under 1 metric ton (like a Very Important Person or a box of serum to treat the Martian Anthrax-Leprosy Pi epidemic.). The entire mission is nothing but Terra/Mars transits.

A gas-core rocket with radiator on a courier mode mission (hot pink curved line) has truly jaw-dropping performance. It can do an entire mission in only 80 days!

Just for comparison sake, the report includes a fusion rocket with typical high specific impulse but miniscule thrust (orange curved line). The fusion ship has a power plant specific mass ("alpha" or "α") at a very advanced 1 kg/kW. It has extremely low IMEO's if the mission time is greater than 250 days. But below that mission time the fusion ship's performance is lackluster. This is because the fusion drive is low thrust and is power-limited. In order to accelerate up to cruising speed in sometime less than a decade it has to increase its thrust at the expense of the specific impulse. Which sends its IMEO skyrocketing.

Unlike the fusion drive, the radiator-cooled gas core nuclear rocket is not power-limited, it is specific impulse limited (as shown in Figure 6A, see how it rapidly reaches a plateau?). This means if it trades thrust for specific impulse, it isn't reducing the specific impulse very much at all. It can crank up the thrust so it gets up to cruising speed in only two or three days. Then it can drop down to high specific impulse fuel economy gear for the rest of the 80 day mission, at a vast savings in IMEO.

Actually one can calculate the functional equivalent of α for the gas-core drive by using Figure 6. Thrust power is:

Fp = (F * Isp * 9.81) / 2

where Fp is thrust power in watts and F is thrust in Newtons. Divide Fp by the engine weight We' to get the engine α. When you do that with Figure 6, all the engines have an α in the range of 0.01 to 0.1 kg/kW, which makes the fusion drive look like a hippo.


Since these rockets were designed to be reusable, it is important to look into the difficulty of refurbishing one for a new mission.

Insipid solid-core nuclear rockets are woefully weak, but at least their nuclear fuel elements don't go anywhere. They stay safe inside the reactor ready for the next trip. Gas-core on the other hand have the drawback that the nuclear fuel elements eventually spew out the exhaust nozzle. The gas-core rocket's uranium requirement for one mission may be considerably less than the solid-core. Unfortunately the solid-core can re-use its uranium several times before more has to be added, while the gas-core has to restore its entire supply with each mission.

In figure 9 the H/U numbers are Hydrogen-Uranium flow ratios. So for instance, a rocket with a H/U of 200 will expend 200 units of hydrogen propellant for each single unit of uranium. The green SCNR curved line has no H/U number, it is a solid core rocket so zero units of uranium are expended regardless of the hydrogen flow (unless there is a catastrophic engine malfunction).

The family of yellow lines of the scientific/exploration missions show several flow ratios. There is only one flow ratio for the courier mission (200), the one in orange.

Since these are ratios you can take the uranium fuel requirement, multiply by the flow ratio, and thus calculate the hydrogen propellant requirement. For example, the 80 day Mars courier mission requires 3,350 kilograms of weapons-grade uranium-235 (98% enrichment) at a H/U of 200. Therefore the hydrogen propellant requirement is 3,350×200 = 670,000 kg.

Due to the fact that solid-core rockets can re-use their uranium a few times, a gas-core needs a H/R of 200 or more to have a lower uranium fuel bill. In 1971 (when the report was written) uranium fuel was roughly $10,000US/kg. Which means the 150-day Mars courier mis