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.
|Core stage (C)|
|Engine Isp, sec||900|
|Inert Mass, mt||44.99|
|x3 25 klbf NTP Engines||12.32|
(w/ everything else)
|Usable LH2 Mass, mt||41.64|
|RCS Usable Prop Load, mt||17.05|
|Boil-off to ullage, mt||0.20|
|Stage Length, m|
(engines, RCS, I/F)
|Approx. Effective LH2|
PMF / λ
|In-line Tank (I)|
|Inert Mass, mt|
|Usable LH2 Mass, mt||66.40|
|RCS Usable Prop Load, mt||5.51|
|Engine Isp, sec||900|
|Stage Length, m|
(incl. RCS & I/F)
|Approx. Effective LH2|
PMF / λ
|Saddle Truss & Drop Tanks, 1 ½ (D)|
|Inert Mass, mt||38.35|
|Usable LH2 Masses mt||103.30|
|RCS Usable Prop Loads, mt||8.58|
|Engine Isp,||sec 900|
|Stage Length, m|
(incl. RCS & I/F)
|Approx. Effective LH2|
PMF / λ
|Deep Space Hab (stocked)||51.85|
|MPCV (CM+SM, no prop)||14.49|
|Less mass exp.|
prior to TMI, mt
wet mass total, mt
wet mass total, mt
|Saddle Truss & Drop Tanks|
wet mass total, mt
wet mass total, mt
|Mars stack interim total||436.43|
|Total TMI Stack Mass, mt||411.26|
Abbreviations in table:
- 25 klbf NTP Engines: 111.2 kN Pewee-class nuclear thermal engines
- PMF = propellant mass fraction
- λ = Payload Fraction
- PMF / λ = propellant fraction to payload fraction ratio
- RCS = reaction control system
- I/F = I'm not sure but has to do with docking structure used to connect stages
- MPCV = Multi-Purpose Crew Vehicle (Orion spacecraft)
- CM+SM = Orion Crew Module plus Orion Service Module
- TMI = Trans-Mars Insertion, first burn at mission start to leave Terra orbit
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.
|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
|# Engines / Type:||3 / NERVA-derived|
|Engine Thrust:||25 klbf (Pewee-class)|
|Specific Impulse, Isp:||900/nominal - TBD/max sec|
|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:
- core propulsion stage
- in-line tank
- 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.
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.
|Nuclear rocket using|
Indigenous Martian Fuel
|Reactor Power||2,513 MWth|
|ΔV at 1400 K||3,254 m/s|
|ΔV at 2800 K||5,685 m/s|
|Isub>sp at 1400 K||162 s|
|Isp at 2800 K||283 s|
|Ve at 1400 K||1,589 m/s|
|Ve at 2800 K||2,776 m/s|
Nuclear rocket using Indigenous Martian Fuel (NIMF) is from a 1980's Martin Marietta study by Robert Zubrin. The basic idea was to attempt to avoid the tyranny of Every Gram Counts by using in-situ resource utilization for the propellant. So instead of lugging miserly limited amounts of high-performance propellant all the way from Terra, the rocket would make do with unlimited amounts of low-performance propellants available locally on Mars.
There were several vehicles designed: including a supersonic winged craft and a ballistic "hopper". The latter is pictured here. They all featured a single solid core nuclear thermal rocket engine fed by a single propellant tank.
The easiest propellant to manufacture is liquid carbon dioxide. It can be produced from the Martian atmosphere using just high pressure (690 kPa) with no cryogenic cooling needed (a 30 horsepower pump will do, requiring 25 kW, or 80 kilowatt hours per metric ton). The Martian atmosphere is about 95% CO2 so it is not like there is any shortage of the stuff.
Other propellants have superior performance but are much harder to manufacture. Carbon dioxide has enough specific impulse to boost the NIMF from the surface of Mars into low Mars orbit, so the designers figured it was good enough. It also has enough Isp to hop the vehicle from point A to any other point on Mars.
As I mentioned each metric ton of propellant sucked out of the atmosphere takes 80 kilowatt-hours, and the propellant tank holds about 302 metric tons total. About 24,160 kilowatt-hours to fill it. How long it takes to fill the tank depends upon how many kilowatts the power source can feed the pump.
- Power can come from the nuclear reactor in the rocket engine, if you make it bi-modal. It would produce about 100 kilowatts of electricity (kWe).
Advantage: The report says this will fill the tank in 12 days (my slide rule says it will take 10 days at 100 kWe for 24,160 kWH). It would also require zero mass for the power supply, since the rocket engine has already been accounted for.
Disadvantage: is that operating the reactor while the ship is landed with spray deadly radiation all over the landing site. This makes it difficult for the crew to do things like disembark, embark, and linger near the ship.
- Power can come from a solar cell array. It can produce about 25 kWe (averaged around the clock to take account of nighttime)
Advantage: no radiation
Disadvantage: The report does not mention how long it will take to fill the tank but presumably 1.2 as much time as the RTG: 60 days (my slide rule says it will take 40 days at 25 kWe for 24,160 kWH. Why 1.2? 30 kWe / 25 kWe = 1.2). The array is about 3,500 m2 and has a penalty mass of 8.8 metric tons. It also takes three crew members about 2 days to set up and break down, making the total delay about 44 days between flights.
- Power can come from an RTG. It can produce about 30 kWe.
Advantage: practically no radiation. It has a penalty mass of 4 metric tons, about half of the solar cell array. Unlike the solar cell array it requires no setup or breakdown time. The report says this will fill the tank in 50 days (my slide rule says it will take 34 days at 30 kWe for 24,160 kWH), which is less than the solar cell array.
Disadvantage: It has a penalty mass of 4 metric tons as compared to the bi-modal engine. It will take many more days than the bimodal engine.
The paper decided the RTG was the optimal solution.
A reactor temperature of 2800 K (Isp 283) is required to boost the vehicle from the surface of Mars into high orbits (ΔV 5,685 m/s).
But only 1400 K (Isp 162) would be needed for Mars to Mars hops (ΔV 3,254 m/s).
From Top to Bottom:
- Storage Dome
- Flight Deck
- Habitation Deck (crew living quarters, supports a crew of three for more than one year)
- Mechanical Deck (machinery to liquifly atmospheric carbon dioxide)
- Propellant Tank
- Nuclear Engine
The nuclear engine has a shadow shield on top composed of steel, boron, and lithium hydride to protect the crew from radiation when the reactor is operating. A secondary toroidal propellant tank surrounds the reactor to protect crew walking on the surface when the reactor is idling. There is a second radiation shield right under the crew, to protect them from backscattered radiation reflected off the ground during landing.
The reactor elements will need special cladding, since when carbon dioxide is heated to high temperatures inside the reactor it will oxidize the heck out of everything it touches. Reactor elements designed for liquid hydrogen propellent won't work, they will rapidly erode and spray powdered glowing radioactive death while the reactor stops producing power and the ship plummets out of the sky.
Mars Excursion Module
|300 km by|
e = 0.9
|Thrust||133,500 N to|
|Burn Time||48 seconds|
|Total Mass||3,357 kg|
|Deorbit ΔV||200 m/s|
(mom circular) or
|1.5 to 0.15|
|Tanks & System||1,179|
|Ascent ΔV||4,880 m/s|
(mom circular) or
|Ascent Thrust||156,000 kg|
|ASCENT CAPSULE||(2,386 kg)|
|STAGE II||(4,277 kg)|
|Tanks & System||313 kg|
|STAGE I||10,210 kg|
|Tanks & System||730 kg|
The Mars Excursion Module is from a 1966 study by North American Rockwell. This was the first Mars lander designed after the bombshell from Mariner 4 that astronomers had drastically over-estimated how dense the Martian atmosphere was. They had figured it was a useful 85 hectopascals (hPa), in reality it was an almost worthless 6 hPa (just slightly better than a vacuum). By way of comparison Terra's atmospheric pressure at sea level is 1013 hPa.
The poor prior design that was rendered obsolete by the low atmospheric pressure was the Aeronutronic MEM
The low atmospheric blow Mariner 4 dealt to the scientist was just the cherry on top of the sundae. Much more serious was the photographs. The scientists knew there could be no chance of images of scantily-clad Barsoomian princesses, but they were hopefull there would be some lakes and maybe even a canal or two. But nothing but a bunch of freaking craters? The scientists got a sinking feeling in their stomachs, as they could almost see the Mars exploration program go swirling down the toilet right before their very eyes. Once the taxpayers saw these photos the NASA tax dollars would dry up. Mars looks like the freaking moon, for cryin' out loud! And NASA has already been to the moon. Been there, done that, got the T-shirt. No need to go to Moon part deux.
But NASA put a brave face on things, and proceeded to plan for a Mars mission anyway. Sadly, they were right. As I write this it is fifty years later and the movie The Martian is still science fiction, not a documentary. That furry "whumph" noise you hear is RocketCat doing a facepalm.
Given the pathetic whisp of Martian atmosphere, NAR went with a classic gum-drop shape much like the Apollo command module for aerobraking purposes. For one thing all the expertise obtained from Apollo could be leveraged. Plus there was Terra's atmosphere conveniently located for heat shield test purposes.
Though I did read a recent report suggesting that even with the gum-drop design the Martian atmosphere is not up to the task of aerobraking the MEM before it splats into the ground at hypersonic velocities. The report suggested that entirely new technologies are needed.
In a genius move, NAR made the design modular. If you needed a lean and mean mission, you could remove some internal compartments, ascent propellant, and surface supplies to get the total lander mass down to 30 metric tons. Or you could max it out. Or anything in between.
The price of a low mass lander is that it could only support two crew for four days, and the mothership had to be in a low circular Mars orbit for both departure and return to the mother. The high mass lander needed lots more delta V from the mothership, but it could support four crew for thirty days, and the mothership could be in a high elliptical Martian orbit.
You can also make a lander with no ascent stage at all. This can be used to land supplemental equipment, such as an extended-stay shelter, nuclear power module, or a huge Mars mobile lab with fuel supply.
- MEM has a mass of 49,437 kg when it separates from the mothership. The deorbit motors fire for about 200 m/s ΔV. Deorbit motor has a thrust-to-weight-earth of 0.4. The MEM starts falling out of orbit, and the deorbit motors are jettisoned. The MEM now has a mass of 46,078 kg.
- The MEM enters the Martian atmosphere at an angle of attack of 147°. It starts aerobraking, subjecting the crew to about 7 g's. When it slows to a mere Mach 3.5, it pops a hypersonic drogue chute to stabilize then inflates a 18 meter diameter ballute. This will slow the MEM down to Mach 1.5.
- Once the MEM lowers to 3 kilometers of the surface, it jettisons the ballute. The plug in the heat shield over the descent engine is jettisoned. The descent engine is canted about 13° off center, because the MEM center of gravity is off center, because due to design consideration the MEM is not radially symmetric. Mostly because of that pesky crew quarters and laboratory.
- The conical section of the heat shield is jettisoned. The descent engine ignites and burns for 1,070 m/s to 1,450 m/s ΔV, depending upon whether the mothership was in a circular or elliptical orbit when the MEM detached. At this point the engine will have a thrust-to-weight-earth of 1.5 to 0.15.
- The design managed to squeeze in enough extra propellant for about two minutes of hovering (about 457 m/s of ΔV). Which could be a life-saver if the landing site unexpectedly turned out to be full of jagged boulders or something. Instead of hoving, the extra propellant can move the MEM laterally about 6.7 kilometers to an alternate landing site. The design had six landing legs. In concert with the incredibly stable gum-drop shape, they could manage a ground slope of up to 15°. Actually the shape is similar to a no-spill coffee mug, and for the same reason.
- The crew then frantically does as much Martian scientific research as they can cram into 30 days. The pressurized volume is 21.6 m3 (14.4 is laboratory/living quarters, 7.2 is ascent capsule). 20% is taken up by equipment, leaving barely 4.3 m3 per crewperson (right at the ragged limit before claustrophobia strikes).
- When it is time for departure, the descent stage becomes the launch pad (which stays behind on Mars), and the center becomes the ascent stage. It brings the crew and 136 kilograms of Martian geological samples back to the orbiting mothership. It launches as Ascent Stage I.
- When the Stage I tanks run dry, they are jettisoned. The ascent stage continues as Ascent Stage II. The two stages have a combined ΔV of 6,200 meters per second. The ascent has five components.
- Initial burn to 19 kilometer altitude (mothership circular: 4,206 m/s ΔV, mothership elliptical: 4,286 m/s ΔV)
- Coast to 185 kilometer altitude
- Burn to circularize orbit (23 m/s ΔV)
- At appropriate time, burn to ascend for rendezvous (mothership circular: 168 m/s ΔV, mothership elliptical: 1,327 m/s ΔV)
- Rendezvous with mothership at apoapsis
- Total ΔV: mothership circular 4397 m/s, mothership elliptical 5,635 m/s
- The ascent stage docks with the mothership using its Reaction Control System (RCS). It has 100 meters per second of ΔV left for the docking at this point. The rest was burnt during the descent and ascent phases.
The deorbit stage uses a beryllium solid rocket fuel with a specfic impulse of 300 to 325 seconds, a thrust of 133,500 to 204,600 Newtons, and a burn time of 48 seconds.
The reaction control system was supposed to use Chlorine pentafluoride (ClF5) oxidizer with Mixed Hydrazine Fuel-5 (MHF-5). The latter is a devil's brew of monomethylhydrazine, unsymmetrical dimethylhydrazine, diethyline triamine, acetonitrile, and hydrazine nitrate. Which is just as vile as it sounds. It has a specific impulse of 336 seconds.
The one joker in the deck was the specified fuel for the descent and ascent stages. It seems they couldn't quite get the ΔV they needed out of conventional liquid oxygen (LOX) and liquid hydrogen (LH2). With the mass ratio the design had (i.e., the tiny fuel tanks), LOX/LH2 could not even manage the 4,880 m/s ΔV required to reach the mothership in a low circular orbit, much less the 6,200 m/s ΔV required if it was in a high elliptical orbit. The problem was that LH2 takes up a lot of room, but the MEM's fuel tanks are cramped. There wasn't room for enough LH2 even if the entire area was converted into one gigantic tank.
So they used FLOX and liquid methane (CH4) instead. That can do 6,200 m/s ΔV easy because liquid methane is more than six times as dense as liquid hydrogen. So you can cram six times as much liquid methane mass into the same sized tanks. Using FLOX instead of LOX makes up for the lower energy in methane. FLOX/CH4 has a specific impulse of 383 seconds, compared to LOX/LH2's specific impulse of 449 seconds. LOX/CH4 is lucky to get a pathetic 299 seconds.
Fluorine is beyond insanely dangerous. It is incredibly toxic, and will corrode almost anything (some explosively). They don't call it "The Gas of Lucifer" for nothing. The pious hope of the MEM designers was to contain the FLOX in tanks lined with nickel or something similar that would form a passivation layer.
The FLOX mix is 82.5% fluorine and 17.5% oxygen. Mixing liquid fluorine and liquid oxygen is actually relatively safe. For some odd reason those two will not chemically combine without some coaxing. If they do combine, however, you get the dreaded compound Dioxygen Difluoride. This is the compound with the chemical formula FOOF, which coincidentally is the sound your laboratory will make as it blows up. This is the most famous compound in Derek Lowe's hysterical list of Things I Won't Work With (take a minute to read it, the article is a scream).
Another concern is that in a tank the fluorine and oxygen might separate. Then the engine would periodically be sucking pure fluorine, which certainly will not be doing the engine any good.
Carrying entire tankfuls of ultra-corrosive flaming explosive death to Mars seems to be a questionable decision, to say the least. If the MEM lands a trifle hard and the tanks rupture, you won't have the basis for a re-make of The Martian. More like a large melted crater with a few odd pieces of corroded metal and polished skeleton bits at the bottom.
At least the MEM designers saved mass on the ignition system. You don't need any. FLOX/CH4 is hypergolic (because fluorine is hypergolic with almost anything). This is also a help when the ascent stage does staging, you can easily re-start the engines in mid-flight.
I'm doing more research, but apparently the MEM design is so popular, that it was later redesigned just a bit to remove the need for liquid fluorine oxidizer. This would involve removing equipment and increasing the size of the fuel tanks.
These are from Technological Requirements Common to Manned Planetary Missions: Appendix D by the space division of North American Rockwell (1965). They detail a s family of Planetary Excursion Modules (PEM).
Here I will focus on the retrobraking PEMs, that is, the ones that use retrorockets to land because the target planets have no atmosphere to allow aerobraking. These particular PEMs were intended for landing on Ceres, Vesta, Ganymede and Mercury. They are based on the Apollo Lunar Excursion Module with an Ascent Stage on top of a Descent Stage. They are designed for a 30 day stay on the planetary surface, before returning to orbit.
They do, however, have Rockwell's fixation on using the insanely dangerous FLOX as an oxidizer, and the only somewhat dangerous Monomethylhydrazine (MMH) as fuel. The reason is that liquid hydrogen requires preposterously huge tanks, but if you use anything else the specific impulse goes way down. Unless you take the mad-scientist step of using FLOX oxidizer to make up for it.
In the diagrams below, the Ascent Stage is pink, the Descent Stage is green, and the mission crew quarters is gold.
3-Crew Ceres-Vesta PEM
|Propulsion Isp||333 s?|
(30 day, 3 crew)
The mission crew quarters is merged with the ascent stage. Note the Descent Stage Dry Mass does not include the mass of the ascent stage.
10-Crew Ceres-Vesta PEM
|Propulsion Isp||333 s?|
10-Crew Ceres-Vesta PEM
10-Crew Ceres-Vesta PEM
(30 day, 10 crew)
A page appears to be missing from my copy of the document
3-Crew Ganymede PEM
|Wet Mass||17,300 kg|
|Propulsion Isp||333 s?|
3-Crew Ganymede PEM
3-Crew Ganymede PEM
(30 day, 3 crew)
The mission crew quarters is merged with the ascent stage.
10-Crew Ganymede PEM
|Propulsion Isp||333 s?|
10-Crew Ganymede PEM
10-Crew Ganymede PEM
(30 day, 10 crew)
The mission quarters is merged with the descent stage, instead of the ascent stage as is the case with the 3-crew vehicles.
|NTR Specific Impulse||1000 s|
|LANTR Specific Impulse||600 s|
|NTR Exhaust Velocity||9,810 m/s|
|LANTR Exhaust Velocity||5,900 m/s|
|Wet Mass||460,000 kg?|
|Dry Mass||? kg|
|Mass Flow||? kg/s|
|NTR Thrust per engine||1,112,000 n|
|LANTR Thrust per engine||3,336,000 n|
|NTR Thrust total||5,560,000 n|
|LANTR Thrust total||16,680,000 n|
|NTR Acceleration||12 g?|
|LANTR Acceleration||38 g?|
This is from a report called AFRL-PR-ED-TR-2004-0024 Advanced Propulsion Study (2004). It is a single stage to orbit vehicle using a LANTR for propulsion. They figure it can put about 100 metric tons into orbit at a cost of $150 per kilogram. You can read the details in the report.
|Propulsion||NTR Solid Core|
|Engine Thrust||213,000 N|
|Total Thrust||852,000 N|
|Engine Thrust||89,000 N|
|Total Thrust||267,000 N|
|Cargo Mass||16,400 kg|
|Gross Mass||57,100 kg|
This is from a 1964 Ling-Temco-Vought, Inc. study for NASA.
|NTR FIRST LUNAR OUTPOST|
|Engine||Solid core NTR|
|Specific Impulse||900 s|
|Exhaust Velocity||8,800 m/s|
|Engine Mass||3,437 kg|
|Shadow Shield Mass||1,500 kg|
|Total Thrust||330,000 N|
|Total Engine Mass||14,810 kg|
|Inert Mass||33,330 kg|
|Dry Mass||129,330 kg|
|Propellant Mass||67,570 kg|
|Wet Mass||196,900 kg|
|Structural Mass||13,360 kg|
|Avionics & Power||1,000 kg|
|Reaction Control||460 kg|
|x3 Engines||10,310 kg|
|x3 Shadow Shields||4,500 kg|
|INERT MASS||33,330 kg|
|Total Payload||96,000 kg|
|DRY MASS||129,330 kg|
|LH2 Propellant||66,540 kg|
|RCS Propellant||1,030 kg|
|Total Propellant||67,570 kg|
|WET MASS||196,900 kg|
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.
|Trans-Lunar Injection (TLI) Payload|
|96 MT (pilot vehicle and TLI stage adaptor)|
|ΔV||3,200 m/s + gravity losses|
|Initial Orbit||185 km circular LEO|
|Isp||870 sec (graphite)|
900 sec (composite)
960 sec (ternary carbide)
|External shield mass||60 kg/klbf thrust|
|Burn Duration||≤ 30 minutes|
|1% usable propellant|
|Cooldown (effective)||3% usable propellant|
|Residual||1.5% total tank capacity|
|TLI Burnout ΔV||60 m/s|
(30 m/s for trailing edge lunar flyby)
|Geometry||10 m diameter cylindrical tank|
with √2/2 domes
|Insulation||2 inch MLI +|
micrometeoroid shield (3.97 kg/m2)
|All other dry masses||10%|
P.A.R.T.S. is a 2002 study by the Embry-Riddle Aeronautical University for a reusable Earth-Mars cargo spacecraft utilizing a VASIMR propulsion system powered by an on-board nuclear reactor. The report has lots of juicy details, especially about the reactor. Thanks go out to William Seney for bringing this study to my attention.
RMBLR (Rotating Multi-Megawatt Boiling Liquid-Metal Reactor) "Rambler" System. Fuel: Blocks with coolant channels UN+Moly alloy with Rhenium & hafnium, Primary coolant : Potassium, Reactor outlet temperature:1440K, power conversion: Direct Rankine, Specific Mass: 1-2kg/kWe @ 20 MWe assuming a bubble membrane radiator.
|Propulsion||Uprated J2 chemical|
|NERVA Specific Impulse||850 s|
|J2 Specific Impulse||~450 s|
|NERVA Exhaust Velocity||8,300 m/s|
|J2 Exhaust Velocity||4,400 m/s|
|Wet Mass||? kg|
|Dry Mass||? kg|
|NERVA Mass Flow||13 kg/s|
|J2 Mass Flow||25 kg/s|
|NERVA Thrust||110,000 newtons|
|J2 Thrust||110,000 newtons|
|Initial Acceleration||? g|
|Length||30 m + boom|
The Pilgrim Observer was a plastic model kit issued by MPC back in 1970 (MPC model #9001) designed by G. Harry Stine. Many of us oldster have fond memories of the kit. It was startlingly scientifically accurate, especially compared its contemporaries (ST:TOS Starship Enterprise, ST:TOS Klingon Battlecruiser, Galactic Cruiser Leif Ericson).
The model kit included a supplemental booklet just full of all sorts of fascinating details. NERVA engine design, mission plan, all sorts of goodies with the conspicuous absence of the mass ratio and the total delta-V.The kit has been recently re-issued, and those interested in realistic spacecraft design could learn a lot by building one. If you do, please look into the metal photoetched add-on kit, and alternate decals. Round 2 Models (the company who re-issued the kit) have some detailed kit building instructions here.
The design is interesting, and has a lot of innovative elements. For one, it uses a species of gimbaled centrifuge to deal with the artificial gravity problem. It also uses distance to augment its radiation shielding, in order to save on mass and increase payload. This is done by mounting the NERVA solid core nuclear rocket on a telescoping boom.
One major flaw with the Pilgrim's design is the fact that one of the three spinning arms is the power reactor. This means that all the ship's power supply has to be conducted through a titanic slip-ring, since there can be no solid connection between the spin part and the stationary part. Another flaw is if you are going to all the trouble to put the NERVA reactor on a boom to get the radiation far away from the crew, why would you put the radioactive power reactors on an arm right next to the crew?
Anyway, the Pilgrim is an orbit-to-orbit spacecraft that is incapable of landing on a planet. It has a ten man crew (four crew and six scientists), and has enough life support endurance to keep them alive for five years. It could also be used as a space station, in LEO, GEO, or lunar orbit. In launch configuration the NERVA boom is retracted and the spinning arms are locked down. In this configuration it is 100 feet long and 33 feet wide, which fits on top of the second stage of a Saturn V booster. A disposable shroud is placed over the top of the spacecraft to make it more aerodynamic during launch.
After launch, the shroud is jettisoned, the spinning arms deploy, and the NERVA engine's boom telescopes out. The spinning arm array has a diameter of 150 feet. The arms will rotate at a rate of two revolutions per minute (safely below the 3 RPM nausea limit). This will produce about one-tenth Earth gravity at the tips of the arms (Level 1), which fades to zero gravity at the rotation axis. Not much but better than nothing. The spherical center section does not spin, a special transfer cabin is used to move between the spin and non-spin sections.
One arm is the crew quarters, one is a hydroponic garden for the closed ecological life support system, and the third is a stack of advanced Space Nuclear Auxiliary Power (SNAP) reactors using Brayton cycle nuclear power units.
The center section is divided into the Main Control Center at the top and the Service Section at the bottom. The very top of the Control Center has the large telescope, radar, and other sensors. By virtue of being mounted on the non-spin section, the astronomers and astrogators can make their observations without having to cope with all the stars spinning around. Also mounted here is the antenna farm for communications and telemetry.
The Pilgrim carries two auxiliary vehicles: a modified Apollo command and service module, and a one-man astrotug similar to the worker pods seen in the movie 2001 A Space Odyssey. They mate with Universal Docking Adaptors on the non-spin section.
The chemical propulsion system consists of three up-rated J-2 rocket engines with a thrust of 250,000 lbs, fueled by liquid hydrogen and liquid oxygen.
The nuclear thermal propulsion system consists of one solid-core NERVA 2B, using liquid hydrogen as propellant. The NERVA has a specific impulse of 850 seconds, a thrust of 250,000 pounds, and an engine mass of 35,000 pounds (the fact that both the J-2 and the NERVA have identical thrust makes me wonder if that is a misprint). It uses a de Laval type convergent-divergent rocket nozzle. The reactor core has a temperature of 4500°F. The core of the reactor is encased in a beryllium neutron reflector shell. Inside the reflector and surrounding the reactor core are twelve control rods. Each rod is composed of beryllium with a boron neutron absorber plate along one side. By rotating the control rods, the amount of neutrons reflected or absorbed can be controlled, and thus control the fission chain reaction in the reactor core.
There is a dome shaped shadow shield on top of the NERVA to protect the crew from radiation. In addition, the NERVA is on a long boom, adding the inverse square law to reduce the amount of radiation. And finally, the cosmic ray shielding around the crew quarters provides even more protection.
Various attitude control and ullage rockets are located at strategic spots, they are fueled by hypergolic propellants.
The mission will start in June of 1979. Mission is an Earth-Mars-Venus-Earth swing-by. It will have a mission duration of 710 days, as compared to the 971 days required for a simple Mars orbiting round trip. This is done with clever gravitational sling-shots, and use of the NERVA 2B.
Mission starts with an orbital plane change to a 200 nautical mile circular Earth orbit inclined 23°27' (i.e., co-planar with the ecliptic). Transarean insertion burn is made with the three J-2 chemical engines (D+0). At this point the Pilgrim 1 becomes the Pilgrim-Observer space vehicle. It will coast for 227 days. Then it will perform a retrograde burn with the NERVA to achieve a circumarean orbit (Mars orbit) with a periapsis of 500 nautical miles and a high point of 5,800 nautical miles (D+227).
The Pilgrim-Observer will spend 48 days in Martian orbit (including several close approaches to Phobos). Then the NERVA will thrust into a transvenerian trajectory (D+275). It will coast for 246 days, including a close approach and fly-by of the asteroid Eros occurring 145 days after transvenerian burn (D+320).
The NERVA will burn into a circumvenarian orbit of of 500 nautical miles (D+521). It will spend 55 days studying Venus.
The NERVA will thrust into a transearth injection (D+576). It will coast for 140 days. Upon Earth approach, it will burn into a 200 nautical mile Earth orbit (D+710). The crew will be out shipped by a shuttle craft following extensive debriefing.
I did some back of the envelope calculations, and the numbers look fishy to me. An Earth-Mars Hohmann and Mars capture orbit will take a delta V of about 5,200 m/s. This is done with the J-2 chemical engine, and will require a mass ratio of 3.3. That is not a problem.
The problem comes with the NERVA burns. The Mars-Venus burn and the Venus-Earth burns have a total of about 14,800 k/s. With a NERVA exhaust velocity of 8,300 m/s, this implies a mass ratio of 5.9. I'm sorry but without staging you are going to be lucky to get a mass ratio above 4.0.
The plastic model kit is allegedly 1:100 scale according to the kit instructions. However, expert model builders who did measurements figured out that various parts are clumsily in different scales. The "arms folded mode" diameter is supposed to be 33 feet, to fit on top of a Saturn V, that is 1:127 scale. The rotating arms and the Apollo M are more like 1:144 to 1:200 scale. At 1:100 the arms have a deck spacing of a cramped 5 feet, the passage connecting the arm to the ship proper is only 2.5 feet in diameter, and the command module on the Apollo M is 20% smaller than the real Apollo CM. So the scale of the plastic model kit is a mess.
The Agamemnon is basically the Pilgrim Observer with the NERVA solid core NTR swapped out for an ion drive powered by a deuterium fusion reactor.
IBS Agamemnon Total ΔV 280,000 m/s Specific Power 39 kW/kg
Thrust Power 1.1 terawatts Exhaust velocity 220,000 m/s Thrust 10,000,000 n Wet Mass 100,000 mt Dry Mass 28,000 mt Mass Ratio 3.57 Ship Mass 8,000 mt Cargo Mass 20,000 mt Length 400 m Length spin arm 100 m T/W >1.0 no
IBS Agamemnon (Interplanetary Boost Ship) masses 100,000 tons as she leaves Earth orbit. She carries up to 2000 passengers with their life support requirements. Not many of these will be going first-class, though; many will be colonists, or even convicts, headed out steerage under primitive conditions.
Her destination is Pallas, which at the moment is 4 AU from Earth, and she carries 20,000 tons of cargo, mostly finished goods, tools, and other high-value items they don't make out in the Belt yet. Her cargo and passengers were sent up to Earth orbit by laser-launchers; Agamemnon will never set down on anything larger than an asteroid.
She boosts out at 10 cm/sec2, 1/100 gravity, for about 15 days, at which time she's reached about 140 km/second. Now she'll coast for 40 days, then decelerate for another 15. When she arrives at Pallas she'll mass 28,000 tons. The rest has been burned off as fuel and reaction mass. It's a respectable payload, even so.
(ed note: in reality, the maximum amount of thrust a single ion drive could put out is about 10,000 newtons, not 10,000,000 like the Agamemnon is cranking out.)
The reaction mass must be metallic, and it ought to have a reasonably low boiling point. Cadmium, for example, would do nicely. Present-day ion systems want cesium, but that's a rare metal—liquid, like mercury—and unlikely to be found among the asteroids, or cheap enough to use as fuel from Earth.
In a pinch I suppose she could use iron for reaction mass. There's certainly plenty of that in the Belt. But iron boils at high temperatures, and running iron vapor through them would probably make an unholy mess out of the ionizing screens. The screens would have to be made of something that won't melt at iron vapor temperatures. Better, then, to use cadmium if you can get it.
The fuel would be hydrogen, or, more likely, deuterium, which they'll call "dee." Dee is "heavy hydrogen," in that it has an extra neutron, and seems to work better for fusion. We can assume that it's available in tens-of-ton quantities in the asteroids. After all, there should be water ice out there, and we've got plenty of power to melt it and take out hydrogen, then separate out the dee.
(ed note: 1,100 gigawatts requires burning about 0.014 kilograms of deuterium per second. For 30 days total burn time this will require about 36 metric tons of deuterium.)
If it turns out there's no dee in the asteroids it's not a disaster. Shipping dee will become one of the businesses for interplanetary supertankers.
The Wayfarer is basically a stock Pilgrim Observer, all the way down to the NERVA engine. Except that the arms do not extend and rotate for artificial gravity.
When creating the Pilgrim Observer, G. Harry Stine started with a 1960's study on creating a self deploying space station. Mr. Stine added the propellant tanks and the NERVA NTR to make it into a spacecraft. You will note the box cover says "Space Station", not "Spacecraft". David Portree identified the space station study in question. Actually studies plural, the Pilgrim was based on an amalgam of several.
- Large Orbiting Research Laboratory (1962)
- Manned Orbital Operation, Special Section (1963) pages 52-63. Article by Owen E. Maynard and Rene A. Berglund.
- US Patent #3300162 RADIAL MODULE SPACE STATION (1964). Inventors Owen E. Maynard, Willard M. Taub, David Brown, Edward H. Olling, and Robert M. Mason.
- Modular Multipurpose Space Station Study MMSS (1965). Study by Lockheed commissioned by NASA Manned Spacecraft Center.
- Pilgrim Observer plastic model supplemental booklet (1970)
- Aerospace Projects Review vol 1 number 6 (must be purchased). Article Self-Deploying Space Stations (NAS1-1630) by Dennis R. Jenkins.
|total weight manned|
|diameter (deployed)||150 ft|
|diameter of hub||33 ft|
|length of spoke||50 ft|
|# of decks|
|launch vehicle||2-stage Saturn V|
|orbit||260-n-mi, 29.5° incl|
|Level||2 rpm||3 rpm||4 rpm|
David Portree said the design below is from an NASA Manned Spacecraft Center team under Owen Maynard and dates from 1962. The pressurized cabins and the access tubes are covered with a meteor bumper for protection (0.99 probability of not more than one penetration per month).
GE came up with a modified 35-kw SNAP-8 power system for this design in 9/64. They looked at placing the reactor at the center of rotation, down below the hub, or at the end of one of the arms. Oddly enough (from a balance standpoint), they favored placing the reactor at the end of one of the arms. I think they did this because the nadir surface of the hub was supposed to carry Earth-observation instruments.
You will notice that locating the reactor in one of the arms was copied in the design for the Pilgrim. This is foolish, since unlike the space station the Pilgrim has no Earth-observation instruments on its nadir surface. As a matter of fact, the Pilgrim already has a reactor on its nadir, inside the NERVA.
If was to re-design the Pilgrim Observer, I would not waste an entire rotating arm on the reactor. Instead I'd make the NERVA into a Bimodal NTR, and use the third arm for extra labs or something. The NERVA is not going to be thrusting during the months the ship coasts, so it might as well do something useful. The Bimodal switch would require the addition of some heat radiators, a turbine, a generator, and a condensor, but that should not be hard to incorporate.
However, the fact that the Pilgrim also had the reactor in one of the arms is yet more proof it was copied from the design of this space station.
The 150 foot diameter of the rotating section is the same figure quoted in the Pilgrim plastic model booklet. The Pilgrim however only rotated at 2 rpm, instead of 4 rpm. The patent #3300162 specfied 3 rpm (citing the spin nausea limit). Take your pick.
In the pressurized cabin, each level had an internal floor to ceiling height of 84 inches, an external deck to deck spacing of 100 inches, and the floor had a diameter of 183 inches.
The patent notes that the advantage of the folding arms is that when the station is boosted into orbit the direction of acceleration is the same as when the arms are spinning. This means that the cargo does not shift. I'm sure G. Harry Stine noted that thrust can occur in a deep space exploration ship as well as a station being boosted into orbit.
|Spine unit||8.3 m square|
|Thrust Life||12 hours|
|Engine Mass||7,000 kg@|
|19.5 m x 9m|
|Dry Mass||134,950 kg|
|Tank LH2 mass||~94,000 kg|
|Propellant Mass||752,000 kg|
|Wet Mass||893,000 kg|
This is from Project APEX: Advanced manned exploration of the Martian moon Phobos (Advanced Phobos EXploration) Preliminary Report. Later Report. On the surface it is a plain-vanilla expedition to Mars with a three-NERVA engine spacecraft.
The special part is the mission is not targeted at the Martian surface, rather it heads to the moon Phobos. Specifically to a fuel processing plant delivered by a precursor unmanned mission, so the crew can set up the plant for extraction and storage of water and hydrocarbons from Phobos' interior. In-situ resource utilization, in other words.
The mission does not depend upon fuel from Phobos, the processing plant is intended for future missions. So if the astronauts fail to set up the plant, they don't all die.
An orbital propellant depot established on Phobos could open up Mars and the asteroid belt to chemically-powered rockets.
In addition, the future exploration of the Martian surface could be greatly facilitated by using the Sabatier reaction to turn Martian atmospheric carbon dioxide into methane and oxygen. Trouble is that Sabatier need hydrogen as input, which would have to be imported from Terra at great expense. Importing the hydrogen from Phobos would be vastly cheaper.
This is actually from the Aerospace Systems Design (AE 483) course at the University of Michigan, a requirement for a baccalaureate degree in aerospace engineering. Since the class had only four months to complete the project, the design is incomplete. The report is from 1992, with typical optimism they figure it will launch in the far-future year of 2010. The Project Apex spacecraft is named the Wolverine after the university mascot.
The class had a few assumptions. Phobos is assumed to be a carbonaceous chondrite asteroid containing 20% water by mass. Further it is assumed there will be available a supply of heavy lift vehicles capable of boosting 150 metric tons of payload into orbit. The third assumption is that there will be several unmanned precursor missions. These will carry out surface mapping of Phobos, take surface samples, and deliver the prototype fuel processing plant. The final assumption is that the class can only utilize technologies that are either ready now or will be by the year 2005.
As a side note, this would be a logical start to the events in our Cape Dread future history.
|Prop/Power engines (3)||22,380.0|
|Heat radiator (rear)||1,200.0|
|Common tanks (2 in rear)||167,154.7|
|Fuel tank cluster (7)||585,041.3|
|Power Bus B||200.0|
|Phobos scientific equip.||150.0|
|Portable antenna equip.||650.0|
|Heat radiator (front)||880.0|
|4 Star Trackers||20.0|
|Telescopes & Pointing Sys.||600.0|
|Solar flare detection||100.0|
|Power Bus C||300.0|
|Ext. thermal transport||700.0|
Yes, I know elsewhere in the report it states the ship mass is 893,000 kg. The report has, shall we say, some inconsistencies.
The spacecraft spine is a truss composed of cubic modules that are collapsible and self deploying. Each module is 8.3 x 8.3 meters, with each edge strut only 0.16 meters in diameter. They are composed of graphite-epoxy composite. The main truss is a stack of 11 modules while each communication truss arm is a stack of 3 modules. They are rated for up to 0.56 g in compression, with a 1.4 factor of safety.
The truss modules are collapsible because heavy lift vehicles have limited payload volume. The truss modules are self deploying because orbital assembly is enough of a nightmare without requiring the poor astronauts to assemble edge struts like a zero-gee erector set from hell.
The twin habitat modules are modified International Space Station modules. Each module is encased in a layer of multi-layer thermal insulation, followed by an outer layer of anti-meteor aluminum. Stringers running the length of each module control bending, and bulkheads encircling the cross section control radial expansion. It is designed to withstand an internal pressure of 11 psi. Each module is 4.7 meters tall by 16.9 m long by 4.2 meters wide. Each of the two modules has one airlock, located on the opposite end from its twin.
The life support system is partially-closed. It has a 90% efficiency recycling breathing mix and a 95% efficiency recycling water. 1,000 kg of oxygen and 5,550 kg of water are carried as consumbables for the mission (includs a 15% contingency). In addition 4,720 kg of dried food is carried.
Halon 1301 was chosen for fire extinguishers because "it leaves no corrosive or abrasive residues within the cabin and few ill effects for humans." Sez them. Wikipedia states that At levels between 7 and 10 percent, mild central nervous system effects such as dizziness and tingling in the extremities have been reported. But it does say it is better for use inside a spacecraft because it produces less toxic by-products than does Halon 1211.
The spacecraft uses tumbling pigeon spin gravity. Spin axis is the spacecraft "z-axis", parallel to the long axis of the twin habitat modules. The target value for the spin gravity is 0.5 g. Maximum allowable spin rate is 4 rpm or below. The spin rate will vary between 2.67 to 3.06 rpm, depending upon the location of the center of gravity at specific points in the mission. The CG will shift location as propellant is consumed. So the radius of rotation (distance from CG to crew modules) will vary from 63.75 meters (full propellant tanks) to 47.75 meters (tanks 1/13 full).
The spin gravity will have to be despun for course corrections and for emergency procedures. There is enough RCS fuel to support eight spin/despin pairs (9,360 kg fuel). Four for course corrections en route to Phobos, two for course corrections en route to Terra, and two for emergencies. Emergencies include EVA to repair ship systems and mission abort scenarios. Spinning up or despinning down takes about five minutes and 585.1 kg of RCS fuel.
Radiation exposure is assumed to come from four sources:
- Solar flares (solar proton storms)
- Galactic cosmic radation (GCR)
- Solar wind
- Spacecraft nuclear engines
Maximum total radiation exposure limit for each crew member was set at 0.65 Sievert per year, and 0.33 Sievert per month. Total for the entire mission was estimated to be 0.95 Sievert.
Radiation protection from the nuclear engines is provided by standard shadow shields composed of tungsten and lithium hydride, and by a minimum of 40 meters between the reactors and the habitat modules. The shadow is set such that both the propellant tanks and the communication platfors are within the shadow.
Protection from solar flares is provided by a storm cellar around the sleeping quarters. The cellar has ceiling and walls containing lithium hydride, the floor has a water tank. No protection is provided from solar wind and GCR, the crew knew the job was dangerous when they took it.
The spacecraft has two 9-meter antenna mounted on the communication truss. These transmit on the Ka-band to geosynchronous relay satellites around Terra. These provide 50 megabit per second full duplex connections to Mission Control on Terra. This allows continuous transmission of voice and video communication, experimental data and obsevations, and telementry.
The communication truss is parallel to the tumbling pigeon spin axis. This allows the antennae to be easily de-spun by simply rotating in the opposite direction, so they stay on target.
Guidance, Navigation, and Control (GN&C) of the spacecraft is achieved by the computer-managed interaction of the navigation, telemetry, and propulsion systems. The computer system consists of nine radiation-hardened, spaceready General Purpose Computers (GPC), each providing 16 MIPS of computing power. All computers will be linked in a FDDI-2 network.
The navigation system consists of four star trackers to determine the spacecraft's attitude and position, an Optical Alignment System to recalibrate the star trackers, nine Inertial Measurement Units to sense linear rates of acceleration, and nine Ring Laser Gyroscopes to measure angular rates of acceleration. This system will also monitor the spinning motion of the spacecraft.
The telemetry system consists of a long range, high gain radar and a short range landing radar. These radars will guide the spacecraft into the proper Phobos rendezvous position.
The nuclear engines are Rocketdyne NERVA derivatives using carbide reactors. Nuclear engines were chosen to reduce overall trip time, thus increasing reliability and efficiency. And also reducing crew radiation dose. Each reactor can operate up to 12 hours at full propulsion, and has a lifetime of three years. Three engines are used to escape LEO, only two engines are needed for the rest of the mission. A reactor core can be brought to full power within 60 seconds. After each main burn, propellant is wasted for up to six hours afterwards to cool the blasted reactor core down.
The three engines are stacked vertically (along the Y axis). This increases stability during spin. The middle engine will be jettisoned when it is no longer needed and the mission completed with the remaining two.
There are three unrefrigerated propellant tanks dropped after escape from LEO. Four tanks are dropped in Mars orbit. There are two permanent tanks used for the trip home to Terra. Nine tanks total. Which confuses me since the blueprint appear to show only 8 tanks.
For normal operations the spacecraft requires 175 kWe of electrical power (life support, communication, experiments, cryogenic cooling of propellant tanks). It will need to supply this power for the entire two-year mission. Solar panels were rejected in favor of rigging the engines to be Bimodal (they call it Dual-Mode). The spacecraft only needs one engine rigged as bimodal to provice 175 kW, but the design rigs all three so there are two backups. All three share a common heat radiator, because radiators are heavy suckers that really cut into your payload budget.
The designers wanted to use low-mass liquid droplet or curie point radiators with the bimodal generators, but they proved to be incompatible with tumbling pigeon spin gravity. So they went with standard weighty heat-pipe radiators. The life support uses heat-pipes with an internal two-phase water loop to transport the heat and an external two-phase ammonia loop to reject the heat. The engine in bimodal power mode uses heat-pipes with a helium-xenon mixture. The engine in thrust mode uses open cycle cooling (the exhaust is the heat radiator) so it don't need no stinkin' heat pipe.
The problem with bimodal is that while an engine is providing thrust, you cannot also use it to generate electricity. It would burn up the power generating heat radiator (a given radiator design can only handle heat in a narrow range). So the design needed an auxiliary power supply for use for the duration of the burn. They figured the maximum burn duration was six hours. So they equipped the spacecraft with a regenerative fuel cell. Those usually are paired with solar cells, but in this case it is paired with the bimodal reactor. The fuel cell has enough fuel to provide a bare 20 kWe for 24 hours (enough for basic life support, communication, and computers). Then when the burn is over bimodal power can be used to regenerate the fuel-cell fuel.
The total mission time is 656 days, using an opposition class mission. 318 to travel from Terra LEO to Phobos (including a Venus flyby to conserve fuel). 60 days are spent at Phobos setting up the processing plant and performing experiments. Finally 278 days are spent traveling back to Terra. The intial perigee kick to place the spacecraft on trans-Mars insertion will regrettably send it through the Van Allen radiation belts, but the dose should be only 0.06 Sieverts. The total dose for the first 30 days is estimated to be 0.28 SV, which is below the 0.33 SV monthly limit.
The spacecraft will enter a 9,400 km Mars orbit, about 22 kilometer higher than Phobos. This is only barely within Mars' gravity well, which greatly lowers the required delta-V. Future missions with Mars landings can also get by with a greatly lowered required delta-V, since they can refuel at Phobos instead of having to lug the landing propellant all the way from Terra.
After careful calculation the spacecraft will perform a phasing burn to put it 6 kilometers over Stickney Crater on Phobos. No closer because of the danger of the RCS control jets blowing foreign objects off the surface into the ship's hull. Due to the low gravity of Phobos (1/1000th g or 1 cm/sec2) the ship will not land so much as it will "dock." When it comes within twenty meters of Phobos it will shoot harpoons and reel itself down to the surface. These are needed since the landing site does not have the same velocity as the center of mass of Phobos.
The spacecraft will remain tethered until departure because Phobos' anemic gravity is too weak to prevent the ship from drifting away.
After 60 days of science the ship will cast off the harpoon cables and push away from Phobos using the reaction control system. The main engines will then put the ship into Terra trajectory. Upon arrival it will burn to enter a high elliptical orbit. The crew will be removed by an orbital transfer vehicle. The crew-less ship will then do a slow transfer to LEO under autopilot. The ship can the be refurbished for a new mission.
|ΔV1||4,500 m/s||Terra orbit to transfer ellipse 1|
|ΔV2||4,170 m/s||Transfer ellipse 2 to Mars orbit|
|ΔV3||2,950 m/s||Mars orbit to transfer ellipse 3|
|ΔV4||2,820 m/s||Transfer ellipse 3 to Terra orbit|
|ΔVm||668 m/s||Course correction, Phobos land/launch|
|Gravity loss||125 m/s|
|ΔV Total||15,233 m/s|
The processing facility extracts water from Phobos regolith and turns it into cryogenic liquid oxygen and liquid hydrogen. This will transform Phobos into a transportation node for the inner solar system. This will allow such things as economically reaching the asteroid belt with a chemically fueled rocket.
Phobos is estimated to have 330 cubic kilometers of water ice (assuming it is a Type 1 carbonaceous chondrite body). Depending on one's assumptions, 4 cubic kilometers of ice could support fuel requirements for the next fifty years of space exploration.
Phobos also has resources such as aluminum, magnesium, silicon, iron, and nickel. This can be used as raw materials for factories producing material fibers, glass, silicon chips, ceramics, magnets and space truss elements.
The processing facility is assumed to have been delivered into Phobos orbit by a prior unmanned mision, and will be set up by the crew of the Wolverine. It is also assumed that Phobos' Stickney Crater is solid rock covered by up to 200 meters of regolith. It is composed of five main parts:
- Excavation of Regolith
- Transportation of Regolith to Facility
- Processing of Regolith
- Storage of Resources
The plant will be set up near one of the walls that make up Stickney crater. This will allow for maximum radiation shielding for the plant given by the natural surroundings.
Power requirements were estimated to be about 1 megawatt: 400 kW for oven to bring regolith up to 700°C, 200 kW for electrolysis, and 400 kW for blowers, magnetic separator, crusher, etc. This will be supplied by a pair of SP-100 nuclear reactors with an output of 550 kW each.
The two reactors will be installed using a LEVPU. The LEVPU is a modified version of a Lunar core sampler used in the Apollo 15 and 17 missions. The LEVPU digs a cylindrical hole and places a casing around it to prevent the hole from caving in. The nuclear reactor is then robotically placed in the casing. The Regolith acts as radiation shielding for the reactors. This allows human operations to occur within 300 m of the reactors.
For the inside details about the Orion propulsion system go here.
If you want the real inside details of the original Orion design, run, do not walk, and get a copies the following issues of of Aerospace Projects Review: Volume 1, Number 4, Volume 1, Number 5, and Volume 2, Number 2. They have blueprints, tables, and lots of never before seen details.
If you want your data raw, piled high and dry, here is a copy of report GA-5009 vol III "Nuclear Pulse Space Vehicle Study - Conceptual Vehicle Design" by General Atomics (1964). Lots of charts, lots of graphs, some very useful diagrams, almost worth skimming through it just to admire the diagrams.
The following table is from a 1959 report on Orion, and is probably a bit optimistic. But it makes for interesting reading. Note that 4,000 tons is pretty huge. The 10-meter Orion (the one in all the "Orion" illustrations) is only about 500 tons.
In other words, if you can believe their figures, the advanced Orion could carry a payload of 1,300 tons (NOT kilograms) to Enceladus and back!
|Gross Mass||4,000 tons||10,000 tons|
|Propulsion System Mass||1,700 tons||3,250 tons|
|Specific Impulse||4000 sec||12,000 sec|
|Exhaust Velocity||39,000 m/s||120,000 m/s|
|Diameter||41 m||56 m|
|Height||61 m||85 m|
|Average acceleration||up to 2g||up to 4g|
|Thrust||8×107 N||4×108 N|
|Propellant Mass Flow||2000 kg/s||3000 kg/s|
|Atm. charge size||0.15 kt||0.35 kt|
|Vacuum charge size||5 kt||15 kt|
|Num charges for 38,000 m||200||200|
|Total yield for 38,000 m||100 kt||250 kt|
|Num charges for 480 km orbit||800||800|
|Total yield for 480 km orbit||3 mt||9 mt|
|Δv 10 km/s|
Mass Ratio (Payload)
|1.2 (1,600 tons)||1.1 (6,100 tons)|
|Δv 15.5 km/s|
Mass Ratio (Payload)
|1.4 (1,200 tons)||1.1 (5,700 tons)|
|Δv 21 km/s|
Mass Ratio (Payload)
|1.6 (800 tons)||1.2 (5,300 tons)|
|Δv 30 km/s|
Mass Ratio (Payload)
|2.1 (200 tons)||1.3 (4,500 tons)|
|Δv 100 km/s|
Mass Ratio (Payload)
|cannot||2.2 (1,300 tons)|
|10 km/s||Terra surface to 480 km Terra orbit|
|15.5 km/s||Terra surface to soft Lunar landing|
|21 km/s||Terra surface to soft Lunar landing to 480 km Terra orbit or|
Terra surface to Mars orbit to 480 km Terra orbit
|30 km/s||Terra surface to Venus orbit to Mars orbit to 480 km Terra orbit|
|100 km/s||Terra surface to inner moon of Saturn to 480 km Terra orbit|
Most of the information and images in this section are from Aerospace Project Review vol 1 no 5. I am only giving you a "Cliff Notes" executive summary of the information, and only a few of the images and those in low resolution. If you want the real deal, get a copy of APR v1n5.
Orion drive spacecraft scale up quite easily. However, unlike other propulsion systems, they do not scale down gracefully. Surprisingly it is much more of an engineering challenge to make a small Orion. It is difficult to make a nuclear explosive below a certain yield in kilotons, and small nuclear explosives waste most of their uranium or plutonium. But it is relatively easy to make them as huge as you want, just pile on the megatons.So in the 1960's when General Atomic made their first pass at a design, it was for a titanic 4,000 metric ton monster.
Alas for General Atomic, neither the United States Air Force (USAF) nor NASA wanted it. USAF had no need for a ship sized for being a space going battleship (they thought they did but President Kennedy smacked them down). NASA wanted nothing to do with a spacecraft that would make the Saturn V and its infrastructure obsolete and pitifully inadequate overnight. So General Atomic heaved a big sigh, and started designing a tiny Orion drive craft with only a 10 meter diameter pusher plate.
However, recently declassified documents reveal that the USAF's decision to cancel plans for the 4000 ton Orion was a near thing. If some of the high-ranking USAF officers had slightly different personalties, today there would be a US Space Force with Orion spacecraft sending expeditions to Enceladus.
Since General Atomic was trying to sell the design to a couple of organizations with vastly different missions in mind, GA made the design modular. There was a basic propulsion system that one could attach any number of different payloads, and customizing the amount of propellant was as easy as stacking poker chips.
In this section we will be focusing on the USAF design. After the USAF lost interest, General Atomic started working with NASA to customize the Orion to their needs. This made the NASA design quite different from the USAF design. NASA was losing intererest even before the partial test-ban treaty of 1963 killed the Orion dead.
|USAF 10M ORION|
|Wet Mass||475,235 kg|
The USAF 10M Orion had three main components: the Orion Drive propulsion module, the stacks of magazines containing the nuclear pulse units (the Propellant), and the Payload Stack.
The Propulsion Module containes the cannon firing the nuclear pulse charges. It also has the massive array of shock absorbers allowing the spacecraft to absorb the nuclear explosion without being crushed like a bug. It also contains 138 "starter" nuclear pulse units. These are half strength units used to initiate a period of acceleration.
The Magazine Stack holds the (full-strength) nuclear pulse units. Each magazine holds 60 units. There are six magazines in a layer, holding 360 units. This design can hold up to ten layers depending upon how much delta V it needs, but if it has an odd number of magazines they must be balanced around the thrust axis. A full load of ten layers contains 3,600 pulse units.
The Payload stack has three components: Powered Flight Station, Personnel Accommodations, and the Basic 12-Meter Spine.
The Spine rests on the propulsion module and has the magazine stack frame attached. The spine contains the spare parts, the repair shack, and mission specific payload. Some of the mission payload is attached outside the spine, such as the Mars Lander.
On top of the spine is the Personnel Accomodations. This holds the life support, the crew quarters, laboratories, and workshops.
On top of the Accomodations is the Powered Flight Station. This contains the anti-radiation storm cellar, which contains the flight controls used when the ship is accelerating (since exploding nuclear bombs make radiation). In an emergency, the entire section can turn into a large escape life-boat rocket and fly away from the rest of the spacecraft. The life-boat has about 600 m/s of delta V and enough life support to keep the 8 person crew alive for 90 days.
Note that in the table the mission specific payload is not included. The more of that which is added, the lower becomes the delta V.
|Nuclear Pulse Unit|
|Total mass||79 kg|
|Specific Impulse||3,350 seconds|
|Exhaust velocity||32,900 m/s|
|0.8 to 1.5 sec|
0.86 sec is std
The USAF nuclear pulse units are atom bombs. They were about 0.6 meters tall, had a mass of 79 kilograms, produced a 1 kiloton nuclear explosion, and produced 2.0×106 Newtons of force per pulse unit. They were basically nuclear shaped charges. 80% of the blast was focused on the pusher plate instead of being wastefully sprayed everywhere.
The latter NASA pulse units had more mass, more Newtons of force, but a lower specific impulse.
The propulsion system also carries 138 "starter" nuclear pulse units. These are half-strength (1.0×106), used to start a period of acceleration. A starter pulse is used on a stationary pusher plate, a full strength pulse is used on a pusher plate in motion. The first shot will be a starter pulse, and the remaining pulses will be full-strength for the rest of the acceleration period.
You see, a half-strength push is enough to push the plate from the neutral position up to the fully compressed position. A full-strength push is enough to stop a plate moving downward and start it moving upward. Using a full-strength push on a stationary plate will give it twice as much as it need, driving the pusher hard into the body of the spacecraft and gutting it like a trout.
And using a half-strength push on a plate in motion will just halt the plate but provide no useful acceleration.
Naturally if one of the full-strength units misfires, the pilot will wait for the pusher plate to settle down then start anew with a fresh half-strength unit.
The layer of tungsten propellant should be as thin as possible. However, there are limits to how wide it can be (or a pulse unit will have an inconveniently large diametr) and it should be thick enough to stop most of the neutron and gamma radiation (to reduce the radiation exposure on the ship in general and the neutron activation on the propulsion module in particular). The mass ratio of the tungsten propellant to the beryllium oxide channel filler should be about 4:1.
Each unit had two copper bands, that are bitten into by the rifling of the cannon that shoots these little darlings. The rifling spins the pulse units like rifle bullets, for gyro-stabilization.
|Empty magazine||181 kg|
|Single pulse unit||79 kg|
|Loaded magazine||4,921 kg|
|1 stack layer|
in 1 stack layer
(10 stack layers)
in full stack
Pulse units were packaged in disk shaped magazines, 60 nukes per magazines. The magazines were stacked like poker chips on top of the propulsion module, held in a hexagonal truss. There are six stacks, with a maximum height of 10 magazines.
The bottom six magazines attach directly to the propulsion module's feed system. The open end of a magazine fits onto one of the propulsion module's six "pulse system conveyors". On the magazine, a slot in the side called a "sprocket opening" allowed one of the propulsion system's sprockets to be inserted into the magazine. As it spins, the star-shaped sprocket grabs the next pulse unit and feeds it into the pulse system conveyor. From there the pulse system travels deep inside the propulsion module to the launch position.
Pulse units are drawn simultaneously from the bottom six magazines. "Bottom" because that is the layer which attaches to the propulsion system's pulse system conveyor. "Simultaneously" because you do not want the spacecraft's center of gravity straying from the thrust axis. Those pulse units are heavy, and they do not automatically redistribute the mass like fluid propellant in a tank.
When the bottom six magazines are empty (360 pulse units expended), propulsion is momentarily halted, and the six "ejection actuators" (pistons) push on the ejection pad of their respective empty magazine and catapult the empty into space. Sort of like flicking a bottle-cap off the bar room table with your finger. The "stack drive pinions" then engage the racks on the magazine stack and lower the entire stack down until it engages the propulsion system's pulse system conveyor. Propulsion is restarted.
|Pusher plate mass||47,800 kg|
and launcher mass
|Structural mass||17,200 kg|
|Total module mass||107,900 kg|
|Module diameter||10 m|
|Specific Impulse||3,350 seconds|
|Exhaust velocity||32,900 m/s|
|0.8 to 1.5 sec|
0.86 sec is std
The propulsion module is build around a compressed gas cannon that fires nuclear pulse units downward through a hole in the pusher plate. Once the pulse unit reaches a point 25 meters below the pusher plate, the it detonates. The shaped charge channels the explosion into a 22.5° cone perfectly covering the pusher plate.
Since premature detonation of a pulse unit would probably utterly destroy the entire spacecraft, there are incredibly stringent controls on them. The units are locked into safe mode and as such are as impossible to detonate as the designers can possibly make them. Otherwise no astronaut is going to set foot inside a spacecraft carrying enough nuclear warheads to totally vaporize the entire thing. 3.6 megatons is nothing to sniff at.
If everything is nominal, the arming signal is transmitted to a launched unit when it approaches the 25 meter detonation point. If the engine control computer determines that the synchronization between the pusher, the shock-absorber system, and the pulse unit are within tolerances; the detonation signal is sent when the unit arrives at the detonation point.
If anything is wrong, the computer instead transmits the "safety" signal and the unit enter safe mode again. When the pulse unit is a safe distance away, the computer sends a destruct signal. You don't want unattended nuclear explosives just flying through space.
If the computer sends the standard detonation signal but the pulse unit fails to do so, it is automatically disarmed (we hope). Again the destruct signal is sent once the unit is safely away, hopefully the unit will oblige. But since the unit has already failed to detonation on command once already, something is obviously wrong with it. Whether it will actually disarm then destruct is anybody's guess.
It goes without saying that the various pulse unit radio signals will be heavily encryped to prevent sabotage. Especially if the spacecraft in question is a military vessel. Otherwise an enemy ship could send the detionation code to every single pulse unit on board, and cackle as your ship did its impression of a supernova.
Because the pusher plate has a hole in the center, part of the blast will sneak through and torch the business end of the cannon. The cannon has a plasma deflector cone (with 1.27 centimeters of armor) on the end to protect it. When a pulse unit emerges from the cannon, the deflector cone opens for a split second to let it out. The deflector cone has its own tiny shock absorbers, of course. The rest of the conical base of the propulsion module is also armored, since the deflector cone will be deflecting plasma all over the base.
When the blast hits the pusher plate it gives thrust to the spacecraft, like a nuclear powered boot kicking you in the butt at 32 kilometers per second. To prevent this thrust from flattening the ship like a used beer can, two stages of shock absorbers do their best to smooth out the slam.
The first stage is a stack of inflated flexible tubes on top of the pusher plate. It takes the 50,000 g of acceleration and reduces the peak acceleration to a level that can be handled by a rigid structure. Such as the rigid second stage shock absorbers.
The second stage is a forest of linear shock absorbers. They reduce the peak acceleration further to only a few gs.
The structural frame is welded out of T-1 steel I-beams to take the jolt and transfer the thrust from the shock absorbers to the payload. A set of six torus tanks pressurizes the linear shock absorbers and lubricates their interiors with large amounts of grease. You need that grease, those linear shock absorbers are working real hard. If one seizes up the results will be ... unfortunate.
In between blasts the inflatable first stage shock absorbers oscillates through 4.5 cycles (no doubt making a silent sad cartoon accordion noise), while the second stage absorbers go through one half cycle. The first stage oscillates between being 0.6 normal height to 1.4 height.
As each pulse unit is fired, a fine spray of ablative silicone oil coats the pusher plate to help it survive the blast. With the oil coating, each nuclear charge raises the temperature of the plate by only 0.07° Celsius. Typically acceleration periods use only 1,000 pulse units at a time, which would raise the pusher plate temperature by only 70° C.
The effective thrust is the thrust-per-pulse divided by the detonation interval. So 2.0×106 / 0.8 = 2.5×106 N effective thrust. This means a series of nuclear bombs going off every 4/5ths of a second. Boom Boom Boom Boom Boom!
The compressed gas cannon uses ammonia (NH3), stored in the "gas collector and mixing tank". This is the top-most of the torus (donut) shaped tanks around the core. The tank holds about 8 metric tons of ammonia, and 2 kilograms are used for each shot. Which means the tank is good for about 4,000 shots. A full set of magazine stacks + the start and restart pulse system has 3,600 + 138 = 3,738 total pulse units so this should be ample. If more ammonia is needed, there is room to spare inside the propulsion system. Alternatively extra ammonia tanks could be stored inside the Basic Spine.
The cannon barrel is 12 meters long, aimed straight down. The cannon accelerates the pulse unit at 45 g giving them a velocity of 90 meters per second. The barrel is rifled to spin the pulse unit at 5 rps. When the nuclear charge reaches the 8 meter point inside the barrel, exhaust manifolds frantically try to suck out all the ammonia gas and spit it out the ejector gas exhaust tubes. The idea was to have no ammonia between the pusher plate and the detonating nuclear charge. The shaped charge blast could accelerate the ammonia and damage the pusher plate.
The main source of nuclear pulse units is from the magazines stacked on top of the propulsion module. However, the module carries 138 pulse units internally in its "start and restart pulse system." They are stored at the top of the module in six curving channels holding 23 pulse units apiece. These are special half-strength pulse units (0.5 kt, 1×106N). They are used as the first shot for engine start or in the event of a regular pulse unit misfire. A half-strength unit is used on a stationary pusher plate, a full strength unit is used on a pusher plate in motion.
When I was playing around with my Kerbal Space Program Orion Drive mod, I did discover something unexpected. The blasted thing needs lots of RCS attitude jets, it turns with all the speed of a pregnant hippo. As near as I can figure part of the problem is [A] the propulsion system is very lightweight since it is mostly hollow shells and inflated tubes (see diagram below) and [B] the magazine canisters are very dense since they are jammed full of pulse units composed of uranium and beryllium oxide.
|Mission Payload Stack|
station (8 crew)
station (8 crew)
|Terra ⇒ Mars||750 kg to|
Since General Atomic was trying to market the 10 meter Orion to both the USAF and NASA, they made it modular instead of integrated. That way they could have a common propulsion system for both, with customized payload stacks for each. The example payload stack shown here is for a Mars mission. A chemical rocket using a Hohmann trajectory would take at least nine months to travel to Mars. But the Orion drive rocket could go to Mars and back in four months flat! However the mission that General Atomic finally settled on was a more pedestrian fifteen month mission requiring only 22.2 km/s of delta V. This would only need a mass ratio of 1.93. If my slide rule is not lying to me, this means it needs about 2149 pulse units (36 magazines or 6 layer magazine stack).
All the payload stacks started with a Basic 12-Meter Spine at the bottom, resting on the top of the propulsion module with the magazine supports tied to it. In the Mars mission, this contained the space parts and the repair shack.
On top of the Basic Spine was the Personnel Accommodations. This contains the life support system, crew living quarters, and laboratories.
At the very top is the Powered Flight Station. This contains the anti-radiation storm cellar. The crew shelters inside in case of space radiation storms. The crew also shelters inside while the Orion drive is operating, since a series of nuclear detonations is also very radioactive. This is why the flight deck is located inside. Finally the entire level can detach and turn into an emergency life boat if something catastrophic happens to the main ship.
Usually you have a 10-meter propulsion module topped with a Basic 12-m Spine, topped with an 8-crew Personnel Accomodation, topped with an 8-crew Powered Flight Station.
However it is possible to replace the last two items with a 20-crew Personnel Accomodation topped with a 20-crew Powered Flight Station. The 20-crew Accomodation needs its base modified to attach to the Basic Spine, it was originally designed to attach to the larger diameter spine of a 20-meter propulsion module.
Since the spacecraft is long and skinny, it uses the "tumbling pigeon" method of artificial gravity. This is where the spacecraft rotates end over end, at four revolutions per minute. For a 50 meter long spacecraft this would give about 0.45 g at the tip of the nose, gradually diminishing to zero at the point where the basic spine joins the propulsion module. The amount of gravity will change as pulse units are expended, thus shifting the center of gravity, rotation point, and rotation radius.
This does pose a problem in the internal arrangement. While under acceleration the direction of "down" is towards the pusher plate. But while tumbling, the direction of "down" is where the nose of the ship is pointing. So if you are standing on the "floor" during acceleration, when it switches over to tumbling you will find yourself falling "upwards" and end up standing on the ceiling.
As it turns out, if the ship is accelerating it also means that everybody is huddling inside the storm cellar (or dying of radiation poisoning). Therefore the storm cellar is built with "pusher-plate is down" orientation, and the rest of the ship is build with "nose is down" orientation.
This also means that the entire mission payload stack has to have a structure that can handle tension as well as compression.
|Escape Rockets||600 kg|
|Escape Propellant||4,500 kg|
|Vehicle total||25,680 kg|
|90 day life|
|Content total||4,275 kg|
+ Crew (8)
content + operational
|Inside diameter||2.5 m|
|Bunk room height||1.6 m|
This section contains the flight controls and the reaction control system. The unshielded point at the top is the navigation station. The unshielded room below is full of the emergency supplies.
Since this is the section farthest from the detonating nuclear pulse units, it makes sense to locate the anti-radiation storm cellar here. In the diagram it is the rooms inside the thick radiation shielding. You get extra protection via the inverse square law at no cost in shield mass. The radiation created by operating the Orion drive is also the reason why all the flight controls are located inside the storm cellar. Otherwise the pilots will be forced to be at flight controls during flight which are located inside the deadly radioactive flux from the drive, making it impossible to recruit Orion drive pilots. The storm cellar will also be used in case of solar proton storms.
It is unwise to put holes in the part of the radiation shield protecting the crew from the pulses, radiation will spray through. This is why access to the Flight Station is from the sides not the bottom, via two pressurized passageways attached laterally. The right hand passageway is attached to an airlock in the emergency supply room, and just has a pressure-tight hatch down at the Personnel Accommodation module, at the other end. The left hand passageway is attached to a pressure-tight door on the Propulsion Control center, and has a full airlock down at the Personnel Accommodation module.
The main radiation shield on the "floor" is composed of 55 grams per square centimeter of lead. Below that is 120 g/cm2 of hydrogenous material, probably water. The side walls and ceiling have 25 g/cm2 of water to protect against backscatter. There actually is some extra protection inside each pulse unit in the form of the channel filler and propellant. The estimate was that the shield would keep the crew exposure down to 0.5 Sievert from the Orion drive, and 0.5 Sieverts from solar flares, for a total of 1.0 Sievert per Mars mission. More recent analysis shows that only 0.5 Sieverts from solar flares is a bit optimistic.
The storm cellar will also have lots of fiberglas sound-proofing. The tungsten propellant striking the pusher plate will make a tremendously huge noise, transmitted by conduction to the entire spacecraft. From freqencies of 7,000 cps to 50 cps the noise will be about 100 to 140 decibels. Without sound-proofing it will damage the hearing of the crew members.
The Flight Station can also detach to become an emergency lifeboat if catastrophe strikes the main ship. It has about 600 m/s of delta V and about 90 days worth of life support for the 8 crew members. Part of the floor radiation shield is from the emergency rocket fuel tanks. A bank of solid rocket booster ejects the Flight Station, and liquid rockets are used for maneuvers. The RCS is already a part of this module.
There was some analysis about angling the nuclear pulse units slightly off-center instead of using a RCS, but thankfully cooler heads prevailed.
Operational Payload Mass
|Structural Mass||7,600 kg|
|Main power supply||3,470 kg|
|Spin gravity system|
tankage & nozzles
|Spin Propellant||4,540 kg|
|Life Support||2,977 kg|
|Reserve Life Support||1,170 kg|
|Food Supply||5,398 kg|
|x4 Space Taxi||625 kg|
|x4 Space Taxi Propellant||825 kg|
|Crew (8)||725 kg|
|Contingency (~5%)||3,315 kg|
|General Dynamics 2-Man Space Taxi|
|Specific Impulse||450 s|
|Exhaust Velocity||4,500 m/s|
|Wet Mass||361 kg|
|Dry Mass||155 kg|
|Propellant Mass||206 kg|
This section contains crew living quarters, the main power supply, repeaters for the navigation instruments and communication gear in the Powered flight station, the tumbling pigeon jets, the life support system, and the food.
If there were several Orion vessels in the mission they would have space taxis, since trying to maneuver and dock with an Orion Drive is like trying to thread a needle with a bulldozer.
Since this is an Orion drive and not every gram counts, this section is built solid. The decks are pressure-tight bulkheads, not non-pressure tight walls. Compartments are accessed via airlocks, so a space suited crew member can enter an area that was vented by a meteor strike without killing everybody. The two passageways at the top lead to the Powered Flight Station above. The left hand passageway has an airlock in this module, the right hand passageway just has a pressure-tight hatch.
The module is 7.2 meters in diameter and had two compartments. Each had a center cylindrical section 3.2 meters in diameter (a continuation of the Basic 12 M Spine). Both compartments could be divided into eight wedges via non-structural partitions. The center sections are for labs and workshops. The wedge rooms listed in the table. Each stateroom is double occupancy, to accommodate the 8 crew persons.
|Deck 1||Deck 2|
|Stateroom Alfa||Stateroom Charlie|
|Stateroom Bravo||Stateroom Delta|
The Powered Flight Station plus the Personnel Accommodation has a total pressurized volume of 200 cubic meters, or 25 cubic meter per crew member (not counting the two passageways flanking the Flight Station). This is actually pretty luxurious. NASA figures a bare minium is 17 m3 per person, and a wet Navy enlisted man is lucky to have 8.3 m3. In addition the Basic Spine is available, but it is only pressurized when needed.
A 20-person Powered Flight Station and a 20-person Personnel Accommodation from a 20-meter Orion can be mounted on a 10-meter spine on a 10-meter Orion. In that case the sum of the Flight Station and Accommodation pressurized volume is 490 m3 or 24.5 m3 for each of the 20 crew.
|Basic 12-Meter Spine|
Operational Payload Mass
|Structural Mass||7,600 kg|
|Repair Equipment||2,270 kg|
The spine has an internal volume of about 97 cubic meters. It contains spare parts, a repair bay, and miscellaneous payload.
There is also an airlock on the bottom allowing repair crews to enter the propulsion module. It is constructed out of materials with a low neutron activation potential. In addition, each pulse unit is only about 1kt (not a lot of neutrons), they are detonated 25 meters away from the propulsion unit (inverse square law), and the pulse unit channel filler plus tungsten propellant will provide shielding. It will be radiologically safe for crews to enter the propulsion module a couple of hours after the the most recent nuclear detonation.
|Orion Mars Mission|
|Dry Mass||91,000 kg|
This is from Manned Planetary Exploration Capability Using Nuclear Pulse Propulsion by Paul R. Shipps. Basically it shows how an Orion-powered Mars mission is so superior to a chemically powered mission that it just isn't funny.
The family of Mars missions uses the basic 10-meter pusher plate propulsion module, since that can be lofted by a Saturn V. If you limit mission designs to non-multistage missions, it still has outrageous amounts of delta-V. The study found it could handle mission with delta-V ranging from 12,000 to 34,800 m/s and payloads from 45,000 to 200,000 kilograms.
The Orion can do the same miniscule mission as a chemically powered rocket if the Orion has a total Initial Mass In Low Earth Orbit (IMLEO) of only 290,000 kg. But that's where the chemical rocket maxes out while the Orion is just getting started. You can load it with metric tons of extra propellant and do the mission in 200 days flat instead of three years. Or you can increase the mission to 400 days, but add lots more scientist to the crew along with tons of scientific instruments. You can even add more fuel and return to an elliptical orbit around Terra using a brute-force rocket thrust braking instead of barbecuing the ship with aerobraking. Orion has power to spare and then some.
The standard Mars missions designed use multiple Saturn V launches to loft the components into orbit. But if you want to cut costs and have the political will, you can boost the Orion spacecraft into orbit with one Saturn V launch — if you don't mind it switching to Orion nuclear pulse drive while still in the atmosphere, starting at an altitude of 50 nautical miles (93 km). This is not totally risk-free, but the risk is manageable. But just try explaining that to your hysterical constituents.
Nucler Pulse Propulsion Module
Internal details of the engine can be found here. It has a specific impulse of 2,500 s (exhaust velocity of 24,500 m/s), a dry mass of 91,000 kg, and an effective thrust of 3,500,000 N. "Effective" because the thrust is not continuous, the nukes go off at about 1 second intervals.
The interesting thing is all the various Mars missions can be performed by the basic Orion propulsion module as is. All you have to do is change the number of nuclear pulse units it carres. The raw might of nuclear fission makes this engine very flexible.
The propulsion module does have limited internal space for internal magazines, but the bulk of the nuclear pulse units are a carried in external magazines, which are ejected when empty.
Propulsion module (hot pink in diagram) does not include payload spine (green) nor the (empty) external propellant magazines with magazine support structure (gold). In other words the 91,000 kg dry mass is just the hot pink part. Especially since the number (and mass) of magazines varies with the mission.
The study authors wanted to avoid a lot of tedious calculation so they used a simplification. To do calculations for all the missions, the correct way is to total up the the mass of the needed empty magazines, the RCS propellant, and whatnot to be added to the "dry mass" for mass ratio calculation. This takes forever. The study authors found out that you get much the same answer if you simply downgrade the propulsion module's specific impulse by a fixed percentage. For this module (with some internal magazine storage capacity), a 4% downgrade of specific impulse would account for magazine weight, magazine support structure, and the Reaction Control System (RCS) fuel. But with lots less math.
For the above reason, instead of calculating delta-V as if the propulsion module had a specific impulse of 2,500 seconds, they instead used 2,500 / 1.04 = 2,405 seconds (and an exhaust velocity of 23,593 m/s).
The mass of the empty magazines, magazine support structure, and RCS fuel is more or less considered to be part of the payload (via the 4% downgrade trick). Naturally the mass of the nuclear pulse units proper is considered to be propellant mass (part of the "wet mass").
Mars Mission Velocity Requirements
Because NASA reports contain eternal optimism the report writers analyzed a Mars mission departing Terra in 1982, a mere 17 years from when the report was written.
This was to be a simplistic, inelegant, brute-force mission. All the maneuvers were done by rocket thrust, no fancy aerobraking was used to reduce delta-V requirements.
They only figured the delta-V for two maneuvers in a given mission: Terra-to-Mars (ΔVout) and Mars-to-Terra (ΔVback). You can get away with this simplification if your spacecraft does not use multistaging. The only reason the study authors used two delta-V measures instead of one is because the spacecraft mass changes so drastically. Lots of payload is consumed or left at Mars, particularly the two Aeronutronic landers.
The mission assumes the spacecraft returns to a Terran elliptical orbit (Terran approach velocity of 11,000 m/s), have a reserve of 300 m/s RCS outbound and 460 m/s inbound, plus a 3% performance reserve. The crew is transferred from the spacecraft to Terra by a separate pickup vehicle based in orbit or on Terra (not carried by the Orion).
For each mission, a 40-day Mars orbit capture period is included in the durations. So the scientists landed on Mars can do as much science as they possibly can in one and one-third months.
|450-day||9,100 m/s||12,500 m/s||21,600 m/s|
|350-day||14,000 m/s||14,600 m/s||28,600 m/s|
|240-day||18,000 m/s||16,800 m/s||34,800 m/s|
Remember these mission are brute-force. NASA trajectory analysis can reduce the trip times by about 50 days or so by using swing-by maneuvers and other fancy mission optimizations. Others reduce the delta-V. For instance, NASA has a Venus swing-by maneuver which can do a 450 to 500 day Mars mission for a low-low total delta-V of 12,000 m/s
|7,600 m/s||4,400 m/s||12,000 m/s|
Again, the point is all these missions can be performed with the exact same Orion propulsion unit by simply modifying the amount of nuclear pulse units carried. Other propulsion systems would need staging or totally different designs of propellant tanks sizes. With Orion you just stack another layer of standard bomb magazines in the rack.
Mars Mission Payload and Duration Options
In figure 2, the ordinate is the Orbit Departure Weight (IMLEO) and the abscissa is Total Payload. Abscissa is in units of one-thousand pounds (103 LB or 450 kg). Ordinate is in units of one-million pounds (106 LB or 450,000 kg) on the left, and in units of uprated Saturn V payloads on the right (127,000 kg).
The total payload is assumed to be split 50-50 into the so-called "round-trip" payload and the "destination" payload. The former is payload carried both to and from Mars, the latter is assumed to be all consumed or abandoned on Mars. 50-50 sounds arbitrary, but as it turns out lots of carefully planned mission studies have something very close to that split.
The dotted line at the bottom contains the anemic chemically-powered miniscule mission previously referred to, helpfully labeled with "Minimal Manned Landing Mission". Rubbing salt in the wound, the report authors point out that this chemical rocket can only carry a small number of crew (requiring each person to have multiple functions, and increasing each person's work-load) and the rocket will need a high degree of expensive subsystem development and optimization because Every Gram Counts. Neither of which apply to a rocket driven by exploding nuclear bombs.
Just in case you might have forgotten what you read in the last ten seconds, the report authors reiterate that one single standard Orion propulsion unit can perform any of the mission on the chart, no expensive development and optimization required. The report authors also wrote that at the top of figure 2, just because.
The two points marked "Reference Designs" are based on specific payload breakdowns of about 145,000 kg. The 450-day reference design has an IMLEO of 522,000 kg (about x4 Saturn V payloads), the 250-day reference design has an IMLEO of 839,000 kg (about x7 Saturn V payloads).
Remember the missions assume that the spacecraft does not carry any pickup vehicle. Once it returns to Terran elliptical orbit the crew is rescued by a separate vehicle stationed in orbit or on Terra.
If you examine the chart you will be interested to find that reducing the mission duration does NOT create an outrageous increase in IMLEO. You want a half-year Mars mission? No problem!
Table 2 contains the weight statements for the two reference designs.
The "radiation shelter" is the over-sized storm cellar, found in all Orion control rooms. All long-duration spacecraft need storm cellar to protect the crew from solar proton storms. All Orion need extra-strength storm cellars because being propelled by the equivalent of a small nuclear war is not healthy for children and other living things.
The storm cellar mass is enough to reduce the radiation exposure from the Orion drive to only 0.5 Sievert per mission. This cellar will keep the dose from solar proton storms at 1 Sievert per mission. The two reference missions have the same storm cellar mass. The 250 day mission has a shorter solar exposure than the 450 day mission, but a higher nuclear pulse exposure because more bombs are needed to shorten the trip. So it equals out.
The propulsion periods when the crew has to retire to the storm cellar are usually short, from a few to about 20 minutes. The nuclear pulse units radiaton flux do not cause significant neutron activation so the crew can access any part of the spacecraft a short time after propulsion shutdown.
The majority of the destination payload is the two Aeronutronic Mars Landers (tail-sitter version). These were designed for a different spacecraft but as it turns out they fit on the 10-meter nuclear pulse rocket with only minor modification (payload spine has to be flattened).
The Exploration Vehicle Configuration
The payload stack consists of the payload spine supporting the flight station at the top, next lower is the personnel accomodations, then the Mars payload including the two landers. The bottom of the payload spine provides crew access to the propulsion module. The lower part of the spine passes through the center of the magazine stack, and encloses a repair-bay/spares-storage room (3 meters diameter by 7.6 meters tall).
The payload spine is flattened in two places to accomodate the landers. If the required pulse unit magazine stack is too tall to fit under the Mars payload, the payload spine might have to be lengthened a bit. This is the only modification the Orion spacecraft is likely to need.
The personnel accomodations is "upside down" because the entire spacecraft is a tumbling pigeon. The center of gravity (CG) of tumbling pigeon rotation moves aft as nuclear pulse units and landers are expended.
Options in Personnel Complement
The dotted line shows how rapidly the IMLEO rises with the number of crew for an 850 second multi-stage NERVA-style nuclear thermal rocket. The solid lines show how modest the IMLEO increase is for extra crew with an Orion boom-boom rocket. Again the report writers harp on the fact that Orion is not subject to Every Gram Counts. With other anemic propulsion systems designers have to have the maximum payload determined at the start of the design process. The max payload is carved in stone. Once you have produced the spacecraft, adding more payload makes it impossible for the spacecraft to do the mission. With the mighty Orion on the other hand, adding more payload just means you just have to add a few more bomb magazines.
Figure 4 illustrates a useful concept called "loading factor".
With the 400-day mission, adding an additional person increases the round-trip payload by about 4,500 kg, once you add in the extra food, water, and air. This additional mass needs additional propellant (pulse units) to propel it. The extra payload plus extra propellant increases the IMLEO by about 11,300 kg. So the loading factor is 2.5 to 1. Which means for every unit of extra payload mass you add, the IMLEO mass increases by 2.5 units.
In other words, for each additional 100 kg of inert weight added (telescopes, cornflakes, meteoroid protection, heavier structure) you need only add 150 kg of propellant to carry it through the journey! (100+150 = 250, which is a loading factor 2.5 to 1) No vehicle change is required, just add more propellant.
For the 200-day mission the loading factor is more like 4.3 to 1.
Options in Terra Return Conditions
The reference design missions assume the spacecraft returns to Terra and uses a modest amount of thrust to enter an economical but wildly ellptical Terran orbit (approach velocity about 11,000 m/s). The missions do not waste payload mass by lugging along a little Terra reentry vehicle, they assume the crew will be rescued by a local vehicle stationed in Terra orbit or on a surface base. The Orion spacecraft will remain in elliptical orbit, available for restocking and reuse.
The report authors looked into two other options.
- Orion spacecraft can be braked into a nice circular LEO orbit, if you are willing to carry additional propellant
- If the priority is to save propellant and reduce IMLEO: you reduce propellant stock, carry a reentry vehicle, and the crew bails out in said vehicle as the Orion goes streaking past Terra on a one-way trip into the dark of the Solar system. The Orion passes by Terra at about 15,000 m/s relative. Another study estimated that the mass for a reentry vehicle for 8 crew and 15,000 m/s is about 6,990 kg.
In figure 5, the ordinate is the Terran Orbit Departure Weight (IMLEO) just like figure 2.
The pair of bars at 50,000 ft/sec (15,000 m/s) is the propellant-saving "abandon ship" option.
The pair of bars at 35,000 ft/sec (11,000 m/s) is the standard reference missions.
The pair of bars at Circular Orbit is the propellant-wasting circular LEO option.
As you can see the "abandon ship" option has a lower IMLEO, though the mass of the reentry vehicle reduces the savings somewhat. And you cannot reuse the Orion. The circular orbit option does have a higher IMLEO, but not by an overwhelming amount.
Single Launch Mission Capacity
The reference missions assume multiple Saturn V launches to loft the components into orbit, where they are assembled. There is a way to use just one Saturn V launch. Unfortunately it involves using the Orion drive. In Terra's atmosphere.
The Orion is used as the top stage, starting at an altitude of 93 kilometers. The savings are substantial, the risks are manageable. But the thought of detonating *Two* *Hundred* *Nuclear* *Bombs* per launch will cause any nukeophobic person to scream in your face at the top of their lungs. Especially if you are a politician and they are one of your constituents.
The reference mission has one Saturn V launch to loft the Orion propulsion module, one launch for operational payload (personnel accomodations unit, remaining vehicle structure, some supplies), one launch for the Mars excursion modules, and a couple of launches carrying nuclear pulse units and miscellaneous small payloads.
And as is typical for any space system, the direct operating costs are dominated by the cost of boosting the stuff into orbit. "Halfway to Anywhere", remember? Reducing the number of Saturn V launches will cut the costs dramatically. Not to mention avoiding the nightmare of orbital assembly.
Figure 6 shows a fully assembled Orion with a gross weight of 635,000 kg (1.4×106 LB) being boosted by a Saturn V with an uprated S-1C stage (since the standard S-1C cannot structurally handle that much payload, plus it needs more thrust and delta-V).
The Orion ignites at an altitude of about 98 kilometers (53 nautical miles) and starts nuking away. This is high enough to protect the eyesight of idiots who cannot be bothered with warnings of not being too close to the launch site and staring directly at freaking nuclear explosions. The Orion arrives at LEO with its mass reduced to 476,000 kg due to burning 159,000 kg of nuclear pulse units.
The Orion then performs some shakedown maneuvers to get all the bugs out. After that the IMLEO mass is about 454,000 kg (1×106 LB). Looking it up in figure 2 we can see that is enough for quite a few mission options. It can do a total payload 250×103 LB (113,000 kg) in a 400 to 450-day mission returning to elliptical Terra orbit. Or even 430×103 LB (195,000 kg) if you are willing to settle for a 450 to 500-day minimum ΔV mission.
You will, however, need one additional launch to boost the crew into orbit. Trying to man-rate a nuclear Orion boost into orbit would be a nightmare. Just man-rating the Orion for deep-space operations is hard enough.
Since the initial Orion gross weight is 635,000 kg and the effective thrust is 3,500,000 N you can see the initial thrust-to-weight ratio during orbital boost is 0.55. This is a pretty low ratio compared to chemical rockets. However the report assures us that detailed trajectory computations (that they do not elaborate on) reveal that for a 2,500 sec Isp rocket this thrust-to-weight ratio actually maximizes the amount of weight delivered to LEO.
SYSTEM ADVANTAGES AND SYSTEM PROBLEMS
There are other advantages to the Orion, besides the flexibility of a single design that the report authors keep mentioning every five minutes. And of course there are disadvantages as well.
Single Vehicle Operational Advantages
Pretty much all the the other Mars mission spacecraft rely upon mult-staging, whether chemical or nuclear-thermal. But not Orion.
One major advantage is a single-stage vehicles can do several test flights and shakedown cruises. You can't do that with multi-stage craft, not if they have to jettison parts of themselves as part of the test. Which means the the brave crew of a mult-stage craft have to set forth on a mission to distant Mars IN AN UNTESTED VEHICLE.
Nothing works perfectly the first time. Shakedown cruises allow debugging the systems, and allow the crew to become familiar with the peculiarities. It also allows any incipient or "break-in" failures to be fixed before departure. Instead of becoming a life-or-death emergency 54.6 million kilometers from the closest help.
Shakedown cruises also allow actual operating performance to be verified, the spacecraft's center of gravity can be trimmed, and unexpectedly high-loss or high-consumption expendables can be supplemented.
Plus any unexpectedly discovered overwhelming problems will result in merely cancelling the mission, instead of a spacecraft lost with all hands in the black depths of space.
Test flights and shakedown cruises are standard procedure in the aircraft, marine, automotive, and other transportation fields. Setting forth on a long voyage without such test is unthinkable, except in the ad-hoc one-shot mult-stage rocket biz. Orion will allow a return to rational testing.
The flexibility of a single design raises its head again, reminding us that it is a vast cost saving to just make one design and reuse it. Instead of making and debugging a freaking new design for every single new mission. This also costs savings in shakedown cruises, since the design bugs will have mostly been already discovered only the specific ship idiosyncrasies will have to be found.
Another advantage is that nuclear pulse units are nicely dense, highly storable, and mostly trouble-free. Other propulsion systems use liquid hydrogen which is pretty much the exact opposite. Liquid hydrogen is annoyingly non-dense, requiring monstrously huge tanks and thus lots of booster vehicles and launch facilities. Liquid hydrogen is not storable at all, suffering from boil-off and thus requiring power-hungry cryogenic cooling equipment. Boil-off also forces closely-spaced successive launches because the longer the hydrogen tanks loiter in orbit waiting for the rest the more hydrogen will be lost. Finally a spacecraft loaded with nice dense nuclear pulse units will have a high ballistic coefficient which will protect it from atmospheric drag deorbiting it. The poor spacecraft loaded with liquid hydrogen will have to depart quickly or suffer a fiery crash.
But the most significant economic advantage is designing mission subsystems while being free of the tyranny of Every Gram Counts. Instead of spending tons of money and time trying to make featherweight (yet reliable) versions of all systems, you can just slap them together out of boilerplate like old Soviet spacecraft. When an additional 100 kg can be carried by simply loading an extra 148 kg of propellant, many subsystem problems become easier to solve.
The main economic disadvantage of Orion is that the pulse units are shockingly expensive. Which is not surprising considering that they are loaded with highly-enriched weapons-grade uranium-235. The official price of HEU is classified, on the black market weapons-grade uranium has a spot price of $10,000 a gram. The back of my envelope says the propellant mass will be roughly 1.4% HEU. Liquid hydrogen on the other hand is about $0.70 US per kilogram.
Enroute Maintenance Capacity
The Orion is far more maintainable enroute than other spacecraft. Especially other nuclear ones.
Orion has very low residual radioactivity, even after a large delta-V maneuver. The nuclear pulse units use beryllium oxide as a channel filler plus tungsten as propellant in order to sop up the neutrons heading for the spacecraft. The idea is it is better to use the neutron energy to accelerate the propellant instead of wasting them and allowing them to turn the butt-end of the ship radioactive. It is safe for the crew to exit the storm cellar surrounding the flight station immediately upon propulsion shutdown. And only a short delay is needed for the neutron activation levels to die down to a safe level, allown crew access to the entire spacecraft. Even the pusher plate.
Nuclear thermal rockets, on the other hand, are neutron activation machines. Once the engine has been used it will be dangerously radioactive for decades to come.
The propellant is packaged in convenient discrete, dense containers instead of being large volumes of liquid hydrogen boiling away in cryogenically cooled propellant tanks. Trying to do maintenance inside a tank of -253°C LH2 is a good way to die. Or trying to do maintenance nearby an LH2 tank. Orions have no cryogentic components (except for maybe the RCS) so all the ship components are easily accessable at temperatures normal for the space environment. This also means the structural members can be composed of ordinary steels, aluminum alloys and titanium instead of exotic hard-to-fix stuff.
To take advantage of this easy access the Orion is designed with a large well-equipped repair bay and spare parts storage area. The ship can be worked on during coasting periods.
If this spacecraft is so great, why ain't NASA using them? Well, there are a few … problems.
The report mentions that there are some uncertainties about the development of the nuclear pulse units, which unfortunately they cannot talk about because it is classified. They are after all basically nuclear weapons.
All such programs have three classes of developmental problems: technical, programmatic (research and development), and poltical. Ordinary rocket projects usually have big problems with the first two classes, but the political problems are minor or nonexistent. With Orion, the bulk of the problems are political.
Orion has the technical problems well in hand, with lots of research and experimentation on ablation, explosive debris — pusher-plate interactions, and impulsive loading on structures.
Orion's programmatic problems are mostly due to the fact that there is no immediate "requirement" for a spacecraft with such a huge thrust and delta-V capacity. So the budgets are limited. The report is of the opinion that if Orion spacecraft are made available, rocket scientists will be falling over each other to take advantage of the oodles of delta-V and thrust they provide.
But Orion's political problems are where the poop hits the fan.
The report says the problem "rather obviously, stems from the fact that nuclear pulse propulsion uses in small scale the same energy source used for nuclear weapons". Translation: the voters are going to scream "OMG!!! YOU ARE TRYING TO MAKE A ROCKETSHIP THAT USES FREAKING ATOM BOMBS!!! ARE YOU CRAZY??!?"
A related political problem is that the Partial Test Ban Treaty forbids civilian nuclear detonations anywhere but underground. Which is a problem for a spaceship. The report optimistically mentions that the treaty provides procedures for its own amendment. Good luck with that.
Aeronutronic Mars Excursion Module
This is from Proceeding of the Symposium on Manned Planetary Missions 1963/1964 Status. The mission carries two of these, the preferred "tail-sitter" version. The "canted" version has problems, and doesn't fit as well on the Orion.
Sadly the design assumed a Mars surface atmospheric pressure of 85 millibars. The discovery by the Mariner 4 probe that the actual value was one tenth of this invalidated the design. This is discussed by David S. F. Portree, where he talks about the successor to the Aeronutronics MEM.
The Aeronutronic MEM was sized for a 40 day stay on the Martian surface with three explorers.
The fuel was a devil's brew of the appallingly corrosive, toxic, and carcinogenic monomethylhydrazine (MMH) mixed with the ever-popular but beyond-insanely-dangerous FLOX. At least it is a re-startable rocket. MMH is hypergolic with any oxidizer, and FLOX is hypergolic with anything.
The reason for this fuel is they needed a specific impulse of at least 375 seconds, but liquid hydrogen fuel just takes up too much blasted room. The designers of the successor to the Aeronutronics MEM had the same problem, so they were forced to use FLOX as well.
This is from
- Nuclear Pulse Space Vehicle Study Vol. I Summary General Atomic division of General Dynamics GA-5009, Vol I
- Nuclear Pulse Space Vehicle Study Vol. III Conceptual Vehicle Designs and Operational Systems General Atomic division of General Dynamics GA-5009, Vol III
- Aerospace Projects Review Vol 1 Number 6
These is from a study of using Orion drive spacecraft to transport cargo to a Lunar base. Since this is Orion, the cargo capacity is huge.
Each nuclear-pulse unit has a mass of about 141 kilograms. The Orion propulsion module carries 900 pulse units internally (126,900 kg), and additional units in magazines stacked on top of the module (92 units per magazine, 90 plus 2 spares. Empty magazine 181 kg, 92 units 12,972 kg, single magazine total 13,153 kg). Pulse units are detonated at 0.86 second intervals to provide the nominal thrust of 3.5×106 Newtons. They have an effective specific impulse of 1,860 seconds (exhaust velocity of 18,250 m/s)
The propulsion module has a mass of 90,946 kg, less the mass of the 900 internal pulse units (126,900 kg). This does not include the mass of magazine rack or any payload support structure. If I am adding correctly a "wet" propulsion module with a full load of pulse units is 217,846 kg. Magazines will be added if 900 pulse units does not provide adequate delta V for the given mission (added in pairs to keep the center of gravity centered).
There were three designs:
|Orbital Ferry||Surface Ferry||Logistics|
Orion Orbital Ferry starts in Low Earth Orbit (LEO) with crew, cargo, and passengers. It travels to Low Lunar Orbit (LLO) under Orion power. In LLO chemical rocket cargo and passenger shuttles transfer cargo and passengers to and from the Orbital Ferry (carried along with the cargo, or sited at the Lunar base). The Orbital Ferry then travels back to LEO under Orion power, where it can be reused.
Orion Orbital Ferry starts in LEO with crew, cargo, and passengers. It travels to LLO under Orion power. It continues down to the Lunar surface. At an altitude of 6 kilometers it switches to chemical rocket power (because landing under Orion power will force the spacecraft to fly through the center of nuclear detonations, voiding the warrenty. And the ship). On the surface the cargo is unloaded by cranes and tractors. The spacecraft lifts off under chemical rocket power until it has delta Ved 640 m/s, then it switches to Orion power. It then travels back to LEO under Orion power, where it can be reused.
Orion Logistics Vehicle starts in LEO with cargo (no crew or passengers). It travels to LLO under Orion power (strictly under remote control/autopilot). There are two flight plans from this point.
In the first, the vehicle parks in LEO. The second stage detaches and lands under chemical power. The spent Orion stages is abandoned in orbit.
In the second the entire vehicle starts landing under Orion power. Near the surface the second stage detaches and lands under chemical power while the spent Orion first stage crashes into the Lunar surface at about one kilometer per second. The Orion stage obviously cannot be reused but the base might be able to salvage the wreckage. The second plan utilizes the awesome might of Orion more fully at the cost of a more risky and complicated flight plan.
The second flight plan drastically lowers the wet mass of the vehicle, allowing a much smaller Saturn or solid rocket booster to loft it into orbit.
The report assumes the Cargo Modules will have a payload density of 272 kg/m3. The modules have a loaded mass of 100,000 kg, a diameter of 10 meters (5 m radius), and a height of 4.7 meters. These are sized so they can be boosted into orbit by a Saturn V. That is, a Saturn V can carry one (1) cargo module into orbit.
The cargo module stack is tied together with wire cables to keep it in compression. The reference designs have a maximum of four cargo modules in a stack, but presumably it could be higher.
The report estimates that for a manned mission: if the lunar base stay time is six months, support is 1,800 kg pwer man-year, and a ferry thrust-to-weight ratio of 0.15, the ferry could transport 400 passengers with the required support. But the report skips over the problem of passenger accomodations during the trip (particularly shielding them from the Orion drive radiation).
Orbit-to-Orbit Lunar Ferry
Remember, this starts in Low Earth Orbit (LEO) with crew, cargo, and passengers. It travels to Low Lunar Orbit (LLO) under Orion power. In LLO chemical rocket cargo and passenger shuttles transfer cargo and passengers to and from the Orbital Ferry (carried along with the cargo, or sited at the Lunar base). It then travels back to LEO under Orion power, where it can be reused.
The ferry carries one Command Module and two Passenger Modules. It has four cargo modules.
The conical structure below the Command Module is the command module adapter section. It supports the Command Module and contains the auxiliary propulsion system. That is used for thrust vector correction and vernier velocity. It uses nitrogen tetroxide, and 50% hydrazine + 50% unsymmetrical dimethylhydrazine (UDMH). Each motor has 5,000 newtons of thrust.
Maximum payload mass fraction is when the thrust-to-weight ratio was 0.15, which is seven cargo modules. The maximum possible was eight cargo modules.
My off-the-cuff estimate of the ship mass:
|Internal Pulse Units|
|x1 Command Module|
(with 3 crew)
|x2 Passenger Modules||5,780|
|x2 Passenger Mod Life Support||1,000|
|x4 Cargo Modules||400,000|
|Command Module Adapter||???|
|x1 Passenger Shuttle|
(no crew, no passengers, no Life Support)
|Pass Shuttle Life Support||1,400|
|x1 Cargo Shuttle||3,520|
|Cargo Shuttle Life Support||400|
The following is me playing number games, no guarantee of accuracy given.
I figure with 92 pulse units per magazine and six magazines plus the 900 internal pulse units the OtO Ferry is carrying 1,452 pulse units. 1,452 @ 141 kg each means 204,732 kg of propellant, for a starting mass ratio of 1.38.
The first leg of the trip from Earth Departure to Lunar Orbit Capture takes a total of 4,296 m/s of delta V.
R = e(Δv/Ve) which means the required mass ratio for the first leg is 1.265. Pf = 1 - (1/R) so the required propellant fraction is 0.2097. Wet mass is 744,114 kg so the total propellant (nuclear pulse units) expended for the first leg is 744,114 * 0.2097 = 156,073 kg. Round up to a whole number of 141 kg pulse units to 156,087 kg (1,107 pulse units).
In Lunar orbit, the cargo and the passengers are ferried to the surface and are now no longer part of the OtO Ferry mass. Neither is the mass for the passenger life support consumables in the Passenger Modules and Passenger Shuttle. Ditto the crew life support in the Passenger and Cargo shuttles. The empty ferries are retained, because Orion has delta V to spare. The new wet mass is 182,141 kg, and dry mass 133,496 kg.
The second leg of the trip from Plane Change to Earth Orbit Capture takes a total of 4,735 m/s of delta V.
R = e(Δv/Ve) which means the required mass ratio for the second leg is 1.265. Pf = 1 - (1/R) so the required propellant fraction is 0.229. Wet mass is 182,141 kg so the total propellant expended for the second leg is 182,141 * 0.229 = 41,624 kg. Round up to a whole number of pulse units to 41,736 kg (295 pulse units).
The OtO Ferry is now in Terra orbit with 49 pulse units to spare.
Remember this starts in LEO with crew, cargo, and passengers. It travels to LLO under Orion power. It continues down to the Lunar surface. At an altitude of 6 kilometers it switches to chemical rocket power. On the surface the cargo is unloaded by cranes and tractors. The spacecraft lifts off under chemical rocket power until it has delta Ved 640 m/s, then it switches to Orion power. It then travels back to LEO under Orion power, where it can be reused.
The ferry carries one Command Module but no Passenger Modules. It has three cargo modules.
The landing module uses LOX/LH2 chemical rockets (specific impulse of 430 seconds). Both the chemical thrust chambers and landing gear are retractable, otherwise the shock from the nuclear pulse charges would snap them off (technical term is "impingement loads"). When deployed the thrust chambers are canted with an angle of 30°.
Scott Lowther is of the opinion that the landing gear does not appear to be capable of extending far enough to clear the pusher plate. The central firing tube protruding from the bottom of the pusher plate is in a most inconvenient location.
Since the entire clanking mess lands, the spacecraft does not have to carry along any cargo or passenger shuttles.
Manned Spacecraft Modules
These are used in the Orbit-to-Orbit Lunar Ferry and Orbit-to-Surface Lunar Ferry. The Logistics Vehicle is unmanned so it needs them not.
The Command Module is shielded to be a storm cellar, and also a shelter from Orion radiation during maneuvering sequences. It is sized for a crew of three. The upper section is the flight deck, the lower is the crew's accomodations. During Orion Drive thrust events and solar proton storms the lower section is also used as the storm cellar for the passengers housed in the Passenger Modules.
The upper compartment is sized at 5 m2 (50 ft2) for a crew of three, while the lower is sized at 5 m3 (180 ft3) per man, assuming that no more than two of the crew is off duty at one time.
Of the 27,510 kg mass of the Command Module, fully 22,380 kg is the radiation shielding. The anti-neutron layer is polyethylene, the anti-gamma-ray layer is depleted uranium.
The side and top shielding is meant to protect from solar storms, the bottom shielding is the shadow shield protecting from the Orion drive. The side and top shielding is 25 cm of polyethylene (25 g/cm2). The bottom shielding is 100 cm of polyethylene (110 g/cm2) and 29 cm of depleted uranium (55 g/cm2). The requirement is to limit an integrated dose to 0.5 Sievert during the nuclear-pulse detonations.
Since the Passenger Modules are unshielded, the maximum number of passengers is limited to how many can cram into the lower part of the Command Module. So the design assumes 20 passengers and two Passenger Modules with a capacity of 10 passengers each. Even though the Orion can carry more than two Passenger Modules, there isn't enough room in the storm cellar for more.
The Command Module can be stretched higher to expand the passenger storm cellar if you simply must add another Passenger Module or two (higher instead of wider in order to minimize diameter of shadow shield). Each additional passenger will add 114 kg to the mass of the Command Module (mostly for radiation shielding, the rest is mostly abort propulsion fuel). Actually to keep the spacecraft balanced it will probably have to have an even number of Passenger Modules.
Passenger Modules have enough life support to keep their 10 passengers alive for five days. Each passenge has 5 cubic meter of volume. The upper deck is the sleeping quarters, the lower is for work, eating, and recreation. The total mass is 4,390 kg. The air pressure is 7 psi.
The Command Module's abort propulsion system can provide 3 g's for three seconds using a solid propellant with a specific impulse of 270 seconds.
The Command Module's crew support mass includes space suits, tools, and crew's personal gear.
Lunar Logistics Vehicle
Remember this starts in LEO with cargo (no crew or passengers). It travels to LLO under Orion power (strictly under remote control/autopilot). There are two flight plans from this point.
In the first, the vehicle parks in LEO. The second stage detaches and lands under chemical power. The spent Orion stages is abandoned in orbit.
In the second the entire vehicle starts landing under Orion power. Near the surface the second stage detaches and lands under chemical power while the spent Orion first stage crashes into the Lunar surface at about one kilometer per second. The Orion stage obviously cannot be reused but the base might be able to salvage the wreckage. The second plan utilizes the awesome might of Orion more fully at the cost of a more risky and complicated flight plan.
The second flight plan drastically lowers the wet mass of the vehicle, allowing a much smaller Saturn or solid rocket booster to loft it into orbit.
It has zero Command Modules, zero Passenger Modules, and four Cargo Modules. There is a rudimentary Forward Module with a few attitude jets and the autopilot, and a chemical rocket landing stage. The landing stage contains the landing gear, which is lower mass than the Orbit-to-Surface Lunar Ferry since it does not have to support the additional mass of the Orion drive. The single chemical engine is centered in the stage instead of being canted at 30° since the Orion drive is jettisoned.
It was assumed that each passenger shuttle would be able to make two trips for each Orion ferry mission. No, I'm not sure what they mean by "trip". Could be from Orion to surface to Orion, or just Orion to surface.
The passenger shuttle consists of three components: passenger compartment, crew/command cockpit, and propulsion module.
The passenger shuttle can transport twenty passenger. The passenger compartment is sized assuming 2.5 m3 per passenger, and a single passenger deck. This gives a diameter of six meters.
The life-support system is open loop since usage time and frequency of use does not justify the expense of a closed-system. System mass requirements are 5.2 kg per man-day, including fixed container weights. Passenger support allowance based on an estimate of 50 kg/man for space suits and personal gear.
The command cockpit is for a crew of two, and can operate completely independent of the passenger compartment (if some idiot passenger vents the compartment to space the passengers will all die but the crew will be just fine.).
The propulsion module uses LOX and LH2 with a specific impulse of 430 seconds. The tanks are sized for just landing, it assumes the tanks can be refilled on the Lunar surface at the Lunar base or propellant depot.
A 2 crew command cockpit is attached to the propulsion module. It has the same mass as the cockpit on the passenger shuttle, but with a contingency mass of 5 percent.
Mass of the cargo shuttle is estimated to be 3,520 kilograms, not counting payload. The cargo landing system was specified to handle up to ten Cargo Modules (one million kilgrams total cargo), presumably with two or more shuttles. With one shuttle, the cargo stack would be 47 meters tall (about 150 feet) which would be a formidable challenge to unload on the Lunar surface.
|USAF 4000 Ton Orion|
|Wet Mass||3,629,000 kg|
(4,000 short tons)
|Pulse unit mass|
|Pulse unit mass|
w/support rollers, etc.
|Pulse unit dim.||80 cm dia × 87 cm high|
|52.4 m ± 2 m|
effec: 3,600 sec
|Detonation delay||1.1 sec|
|Pulse Unit Storage|
|Number pulse units|
|15 km/s ΔV||30 km/s ΔV|
|Average initial accel||1.25 g||1.25 g|
|1,233,000 kg||1,252,000 kg|
|1,280,000 kg||2,068,000 kg|
|Payload mass||1,115,000 kg||308,000 kg|
Most of the information and images in this section are from Aerospace Project Review vol 2 no 2. I am only giving you a "Cliff Notes" executive summary of the information, and only a few of the images and those in low resolution. If you want the real deal, get a copy of APR v2n2.
Orion drive spacecraft scale up quite easily. However, unlike other propulsion systems, they do not scale down gracefully. Surprisingly it is much more of an engineering challenge to make a small Orion. It is difficult to make a nuclear explosive below a certain yield in kilotons, and small nuclear explosives waste most of their uranium or plutonium. But it is relatively easy to make them as huge as you want, just pile on the megatons.
So in the 1960's when General Atomic made their first pass at a design, it was for a titanic 4,000 ton monster. By this time they realized that they would never get permission to launch an Orion from the ground under nuclear-bomb power, so the baseline was Mode III: a gargantuan chemical booster boosts the fully loaded Orion into LEO. So it does not carry the pulse units required to achieve orbit. For that the engine section would have to be taller.
This became the basis for the USAF Orion Battleship. They took the 11,000 cubic meters of the payload shell and stuffed it full of weapons.
|Wet Mass||3,629 tonnes|
(4,000 short tons)
effec: 3,600 sec
|Detonation delay||1.1 sec|
|1.25 g||1.25 g|
|Missiles Silos||3 banks of 30 each|
When the Orion nuclear pulse propulsion concept was being developed, the researchers at General Atomic were interested in an interplanetary research vessel. But the US Air Force was not. They thought the 4,000 ton version of the Orion would be rightsized for an interplanetary warship, armed to the teeth.
And when they said armed, they meant ARMED. It had enough nuclear bombs to devastate an entire continent (500 twenty-megaton city-killer warheads), 5-inch Naval cannon turrets, six hypersonic landing boats, and several hundred of the dreaded Casaba Howitzer weapons — which are basically ray guns that shoot nuclear flame (the technical term is "nuclear shaped charge").
This basically a 4,000 ton Orion with the entire payload shell jam-packed with as many weapons as they could possibly stuff inside.
Keep in mind that this is a realistic design. It could actually be built.
The developers made a scale model of this version, which in hindsight was a big mistake. It had so many weapons on it that it horrified President Kennedy, and helped lead to the cancellation of the entire Orion project. The model (which was the size of a Chevrolet Corvette) was apparently destroyed, and no drawings, specifications or photos have come to light.
Scott Lowther has painstakingly done the research to recreate this monster. If you want all the details, run, do not walk, and purchase a copy of Aerospace Projects Review vol2, number 2. He also made a model kit of the battleship for Fantastic Plastic, you can order one here.
Rhys Taylor is a scientist who is also a master of the 3D modeling package Blender. His animation of a launching Orion drive spacecraft is quite famous, and has been seen by most people who type "Orion" into Google. His more recent project is a battle between US and Russian Orion drive ships out around Jupiter, and a rendition of the proposed Orion Discovery from preproduction of 2001 A Space Odyssey.
Like everything else in 2001, the good ship Discovery passed through many transformations before it reached its final shape. Obviously, it could not be a conventional chemically propelled vehicle, and there was little doubt that it would have to be nuclear-powered for the mission we envisaged. But how should the power be applied? There were several alternatives — electric thrusters using charged particles (the ion drive); jets of extremely hot gas (plasma) controlled by magnetic fields, or streams of hydrogen expanding through nozzles after they had been heated in a nuclear reactor. All these ideas have been tested on the ground, or in actual spaceflight; all are known to work.
The final decision was made on the basis of aesthetics rather than technology; we wanted Discovery to look strange yet plausible, futuristic but not fantastic. Eventually we settled on the plasma drive, though I must confess that there was a little cheating. Any nuclear-powered vehicle must have large radiating surfaces to get rid of the excess heat generated by the reactors — but this would make Discovery look somewhat odd. Our audiences already had enough to puzzle about; we didn’t want them to spend half the picture wondering why spaceships should have wings. So the radiators came off.
There was also a digression — to the great alarm, as already mentioned, of the Art Department — into a totally different form of propulsion. During the late 1950’s, American scientists had been studying an extraordinary concept (“Project Orion”) which was theoretically capable of lifting payloads of thousands of tons directly into space at high efficiency. It is still the only known method of doing this, but for rather obvious reasons it has not made much progress.
Project Orion is a nuclear-pulse system — a kind of atomic analog of the wartime V-2 or buzz-bomb. Small (kiloton) fission bombs would be exploded, at the rate of one every few seconds, fairly close to a massive pusher plate which would absorb the impulse from the explosion; even in the vacuum of space, the debris from such a mini-bomb can produce quite a kick.
The plate would be attached to the spacecraft by a shock-absorbing system that would smooth out the pulses, so that the intrepid passengers would have a steady, one gravity ride — unless the engine started to knock.
Although Project Orion sounds slightly unbelievable, extensive theoretical studies, and some tests using conventional explosives, showed that it would certainly work — and it would be many times cheaper than any other method of space propulsion. It might even be cheaper, per passenger seat, than conventional air transport — if one was thinking in terms of million-ton vehicles. But the whole project was grounded by the Nuclear Test Ban Treaty, and in any case it will be quite a long time before NASA, or anybody else, is thinking on such a grandiose scale. Still, it is nice to know that the possibility exists, in case the need ever arises for a lunar equivalent of the Berlin Airlift...
When we started work on 2001, some of the Orion documents had just been declassified, and were passed on to us by scientists indignant about the demise of the project. It seemed an exciting idea to show a nuclear-pulse system in action, and a number of design studies were made of it; but after a week or so Stanley decided that putt-putting away from Earth at the rate of twenty atom bombs per minute was just a little too comic. Moreover — recalling the finale of Dr. Strangelove — it might seem to a good many people that he had started to live up to his own title and had really learned to Love the Bomb. So he dropped Orion, and the only trace of it that survives in both movie and novel is the name.
From Lost Worlds of 2001 by Sir Arthur C. Clarke (1972)
|Reusable Nuclear Shuttle|
|Specific Power||45.9 kW/kg?|
|Thrust Power||1.4 gigawatts?|
|Specific Impulse||816 s|
|Exhaust Velocity||8,000 m/s|
|Wet Mass||170,000 kg?|
|Dry Mass||30,000 kg?|
|Mass Flow||41.7 kg/s|
|Initial Acceleration||0.16 g|
|Payload 8-burn||45,000 kg?|
|Payload 4-burn||58,000 kg?|
This is a 1970's era NASA concept for a nuclear shuttle. Note that in many of the images the shuttle has a Space Tug crew module perched on top. Design is very similar to the Basic Solid Core NTR. David Portree wrote a nice history of the nuclear shuttle: The Last Days of the Nuclear Shuttle.
Phase I design was for an expendable vehicle with a 200,000-pound-thrust NERVA II engine. It was to be used for several rocket stages on their planned Mars mission vehicle.
The Phase II design is what is pictured below the Class 1 Reusable Nuclear Shuttle (RNS). It had a a 75,000-pound-thrust NERVA I engine and a payload capacity of 50 tons. NASA had an optimistic RNS traffic model calling for 157 Terra-Luna flights between 1980 and 1990 by a fleet of 15 RNS vehicles.
The little attachable crew module has a mass of 9,000 kg. The NERVA engine is 18 meters long and 4.6 meters wide, intended to fit inside a Space Shuttle's cargo bay (the propellant tank can be lofted into orbit on a big dumb booster, but a nuke requires the human supervision). The propellant tank is 31 meters long and 10 meters wide.
The RNS is assumed to have an operational life of 10 Terra-Luna round trips (before the nuclear fuel rods were totally clogged). After that the RNS is attached to a chemical booster and tossed into a remote solar orbit.
The NERVA has a 1360 kilogram shadow shield on top. The shadow shield casts a 10 degree half-angle shadow, shielding was intended to reduce the radiation exposure to 10 REM per passenger and 3 REM per crew member per round trip to Luna and back. But in addtion to the shield it also relied upon propellant, structure, and distance to provide radiation shielding for the crew. Obviously as the propellant was expended, the shielding diminished.
North American Rockwell tried to solve the problem with a "stand-pipe", in which a cylindrical “central column” running the length of the main tank stood between the crew and the NERVA I engine. The central column would remain filled with hydrogen until the surrounding main tank was emptied.
McDonnell Douglas Astronautics Company dealth with the radiation problem by developing a “hybrid” RNS shielding design that included a small hydrogen tank between the bottom of the main tank and the top of the NERVA I engine.
D. J. Osias, an analyst with Bellcomm, pointed out that the radiation dosage received by the astronauts riding the RNS was unacceptable. Osias stated that the maximum allowable radiation dose for an astronaut from sources other than cosmic rays of between 10 and 25 REM per year (0.1 and 0.25 Sievert). But the luckless astronaut on board the RNS would get 0.1 Sieverts every time the NERVA did a burn.
Any external astronauts (not in the cone of safety cast by the shadow shield) at a range of 16 kilometers from a RNS operating at full power would suffer a radiation dose from 0.25 to 0.3 Sieverts per hour. Osias suggested that external astronauts not approach a burning RNS closer than 160 kilometers. Which could be a problem if you are an astronaut in a lunar base when the RNS is burning to leave lunar orbit since the blasted thing orbits at an altitude of only 110 kilometers. If you are standing on the ground track of the RNS you'd better get into the radiation storm cellar.
Nowadays the yearly limit of radiation exposure for astronauts is set at 3 Sieverts, with a career limit of 4 Sieverts. Which means an astronaut piloting a RNS through 40 total burns would be permanently grounded by reaching his career limit of radiation.
There are two mission types: the 8-burn mission and the 4-burn mission.
8-burn mission disadvantage: requires 4 extra burns for change-of-plane maneuvers. This increases the required ΔV to 8,495 m/s, and reduces the payload size to 45,000 kg. Advantage: you do not have to wait for a launch window, you can launch anytime you want.
4-burn mission disadvantage: mission launch windows occur only at 54.6 day intervals. Advantage: since you are not required to perform change-of-plane maneuvers the required ΔV is reduced to 8,256 m/s and the payload size is increased to 58,000 kg.
In both of these missions, it is assumed that the full payload is carried to Luna, where the payload is dropped off EXCEPT for the 9,000 kg that is the crew module. Presumably the crew wants something to live in for the trip back to Terra.
RNS NERVA Engine
|Type||NTR Solid Core|
|Specific Impulse||825 sec|
|41.7 kg/s full power|
0.3 kg/s aftercooling pulse
|Operating Life||10 hours|
|Engine Mass||12,577 kg|
|Power req||28 vdc|
2.3 KW normal
3.5 KW peak
This is from McDonnell Douglas Nuclear Shuttle System Definitions Study, Phase III - Final Report - Volume II Concept and Feasibility Analysis - Part B Class 3 RNS - BOOK 2 System Definitions (1971). This is for the Class 3 Reusable Nuclear Shuttle. It may or may not be the same engine as described above. Thanks to Erin Schmidt for bringing this report to my attention.
The engine has a lifespan of 10 hours of total operation and 60 warm-thrust-chill cycles (I assume 10 hours at full thrust). After that it has to be disposed of, preferably into a distant solar orbit. The back of my envelope says this means roughly 10 Lunar missions before the engine is used up. The problem is that the reactor fuel elements are so clogged with nuclear poisons that they won't react any more. By this time the engine has become so radioactive that it isn't worth the effort to try to extract the fuel elements for reprocessing. Which is a pity since only 15% of the nuclear fuel has been burnt.
The NERVA has an internal radiation shadow shield, but that is a weak one just meant to protect the engine gimbals and thrust frame. To protect the crew there is an optional external shadow shield. The ship designers do their best to use liquid hydrogen propellant as radiation protection insteaad of the external shield, since the blasted shield has a mass of four metric tons.
NDICE is the NERVA Digital Instrumentation and Control Electronics. This allows the pilot to control the throttle, gimbal, and other functions. The part of NDICE that is actually mounted on the engine has a mass of 230 kg.
The engine requires up to 3.5 kilowatts to operate the NDICE, the gimbal electric motors, the turbines, control valves, reactor control drums, and whatnot.
The gimbal pivots the engine for thrust vectoring, used to change the course of the spacecraft. The engine can be pointed up to three degrees off-center in any direction. The maximum rate it can change the pivot is 0.25 degrees per second, but it takes time to get up to speed. It can only accelerate to maximum rate at 0.5 degrees per second per second.
|Specific Impulse||1000 s|
|Exhaust Velocity||9,810 m/s|
|Wet Mass||M kg|
|Dry Mass||M/4.6 kg|
|Mass Flow||? kg/s|
|Initial Acceleration||? g|
|Payload||< M/4.6 kg|
This is a 1965 design from NUCLEAR SPACE PROPULSION by Holmes F. Crouch. It seems to be the father of the NASA Nuclear Shuttle design. According to the book, it would have a single solid-core NTR engine with a specific impulse of 1000 seconds (i.e., an exhaust velocity of 9,810 m/s) 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). The book estimates that an Terra to Luna Hohmann trajectory would take about 12,000 m/s ΔV, after you add in all the change-of-plane maneuvers and added an abort reserve. This would require about 60 hours to travel from the Terra to Luna, but that can be reduced to 20 hours by spending an extra 900 m/s.
In the second diagram, the ship is shown docked to something that looks suspiciously like the Space Tug. Note that they dock nose-to-nose so the lunar shuttle vehicle can stay inside the radiation shadow area.
One really exciting nuclear rocket potiential lies in Earth-Moon transport. The Moon is 208,000 n mi from the Earth. The mission concept simply is one of ferrying back and forth between Earth and Moon terminal orbits. We can think of the ferry terminals as 300 n mi Earth orbits and 100 n mi lunar orbits.
The essence of the lunar ferry concept is presented in Figure 11-8 (the one with the Earth-Moon orbits). the lunar vehicle would do all the propulsive legwork in the the terminal orbits and between the terminal orbits. Chemical systems would be employed as shuttle vehicles at the Earth terminius and at the lunar terminus. This would permit specialization in chemical systems where they are most capable: planetary launch and entry.
The nuclear ferry would have one rocket reactor with capability for multiple reuses, in-orbit replenishment, multiple restarts, and full nozzle maneuverability. We would expect the reactor to have a proven Isp on the order of 1000 seconds. It would have proven reliability, man-rating, pilot control, and long life. We would not expect the ultimate in solid-fueled reactor technology but we should be headed in that direction.
Note in Figure 11-8 that the ferry trajectory is in the form of a "figure-8." This is because it is necessary to transfer from one gravitational force center to another. Each section of the figure-8 can be thought of as an elliptical orbit: one focus at Earth and one focus at the Moon. The two ellipses "join" each other at a transfer region which is about 85% of the distance from Earth (the crossover occurs at about 180,000 n mi from Earth or about 28,000 n mi from the Moon). When going from Earth to Moon, the transfer point is called translunar injection. When going from the Moon to Earth, the transfer is called transearth injection. The injection maneuvers actually start well in advance of the trajectory crossover.
Caution is required when interpreting Figure 11-8. It gives the impression that the launching/entry trajectories, the rendezvous/docking orbits, and translunar/transearth ellipses are all in the same orbit plane with each other. This is not the case. We are dealing with noncoplanar orbit trajectories. Furthermore, they are variable noncoplanar trajectories which change from day to day and from month to month. As a consequence, the target plane — that plane connecting the Earth and Moon centers — "corkscrews" around the major axis of the figure-8 flight path. The corkscrewing of the ferry trajectory introduces fluctuations in the ΔV requirements.
Table 11-4 Nuclear Ferry ΔV Requirements Maneuver Feet per second Earth Orbit Docking 1,750 Earth-Space Plane Changes 3,500 Earth to Translunar Injection 10,000 Translunar to Lunar Orbit 3,500 Lunar-Space Plane Changes 1,500 Lunar Orbit Docking 750 Lunar to Transearth Injection 3,500 Transearth to Earth Orbit 10,000 Midcourse Corrections 500 Abort Reserve 5,000 Total ΔV 40,000
A representative summary of the round trip ΔV requirements is given in Table 11-4. This listing includes all contingencies (a lunar mission can be performed with less ΔV than table 11-4 but the risk-potential increases). Note that total ΔV is 40,000 feet per second (fps). A single stage nuclear vehicle with an Isp of 1000 sec would have a ΔV capability of nearly 50,000 fps. Hence, there is some excess ΔV available.
The unused nuclear ΔV can be applied to reducing the trip time. The normal one-way trip time for a chemical propulsion system is about 60 hours (2 ½ days). Because chemical lunar missions border on marginal ΔV capabilities, the chemical trip time cannot be reduced much below 60 hours. In the case of nuclear systems, for an additional expenditure of 3,000 fps, the one-way trip time can be reduced to 20 hours. The effect of other ΔV expenditures on trip time is shown in Figure 11-9 (not shown), It can be seen that if an attempt is made to reduce the trip time below 20 hours, the extra ΔV requirements are disproportionate to the time gained. Therefore, a value of 20 hours will be selected as the nuclear ferry time base.
If the lunar terminal orbit is 100 n mi altitude, the orbit period is about 2 hours. If the lunar terminal activities necessitate as much as two orbit periods fur completion, the nuclear ferry turnaround could be made within 24 hours of Earth departure. If two nuclear ferry vehicles were used, we could have daily service to the moon and back! All-chemical lunar rocket systems could not possibly compete with this schedule.
The advantages of reduced lunar trip time are self-evident There is reduced time of confinement of astronaut, scientific, and technical personnel to the limited quarters of spacecraft. In-transit boredom and monotony are reduced. Less life support equipment is required: less oxygen, less food, less waste disposal. There is less exposure to weightlessness and less exposure to space radiation. The less the life protection equipment required, the more transport capacity for lunar basing supplies.
In the lunar terminal orbit, all exchange activities would take place at the pilot end of the nuclear ferry. This is because the propulsion reactor would be kept idling. The major features involved are presented in Figure 11-10 (middle image above). One feature not always self-evident is the need to off-load chemical propellants from the nuclear ferry to the lunar shuttle. To make the propellaut transfer, special cargo tanks on the nuclear ferry and special piping on the chemical shuttle would be required, It is assumed that chemical propellants for the shuttle vehicle probably could not be manufactured on the Moon and therefore would have to be transported from Earth.
|Max Width||7.4 m|
This design was the result of a nice bit of collaboration between Walt Disney and Dr. Wernher von Braun (architect of the Saturn V).
Disney's TV show "The Wonderful World of Color" had decades of material for the segments Fantasyland, Frontierland, and Adventureland, but zero for Tomorrowland. Disney's concept executive Ward Kimball had been following Collier magazine's awe inspiring series Man Will Conquer Space Soon, detailing von Braun's plans for manned spaceflight. This series would be perfect for a set of Tomorrowland episodes.
Kimball quickly discovered that he was in over his head, but Disney allowed him to hire technical experts. Kimball proceeded to enlist the main tech experts from the Collier's series: Willey Ley, Heinz Haber, and of course Wernher von Braun. Kimball realized that when it got down to the fine details, you'd have to get help from The Man himself. When Kimball made a tentative inquiry to von Braun, the latter jumped in with both feet. von Braun desperately needed favorable publicity for his Moon mission. The Colliers article reached barely three million viewers. A Disney show could reach tens of millions!
The RM-1's mission was a simple loop around Luna, with no landing (the same as the Apollo 8 mission). The only things you needed was a few days of life-support for the crew, and about 2,700 m/s of delta V. And a bit under 100 m/s to brake back into Terra's orbit. So the spacecraft can be built out of bits and pieces of the existing cargo and passenger ferry rockets.
The front part of the RM-1 was the top stage of the passenger ferry minus the wings but including the passenger section, life support, and engine. Six standard propellant tanks were attached to increase the delta V to 2,800 m/s. When the extra tanks were empty, they were retained as protection from meteors (unnecessarily, meteors are not that common), but jettisoned just before the braking burn into Terra orbit to reduce the ship's mass.
On a nose spike was attached a nuclear reactor, for on-board power. A conical shadow shield protects the crew from reactor radiation. The reactor is ludicrously tiny, in reality it would be quite a bit bigger. And the spike would be a bit longer as well.
A dish antenna for radar and communication is on a set of tracks around the ship's waist. Unfortunately the propellant tanks block the view aft.
It also has a belly docking port for a bottle suit, the port is already standard on the passenger ferry.
The deep space ship above (click on the image for full sized view) was inspired by the Travel Planner spreadsheet in the previous post, and modeled in the wonderfully simple and handy DoGA 3D modeler. The shuttle alongside is a rough approximation of the NASA shuttle, and thus a thorough anacronism in this image, but provided as a scale reference.
Of course you want some specifications of the ship. Even if you don't, you get them anyway:
Length Overall 300 meters Departure Mass 10,000 tons Propellant Load H2 5000 tons Drive Mass 2000 tons Keel and Tankage 1000 tons Gross Payload 2000 tons Flyway Cost $5 billion (equivalent)
The payload includes a hab with berthing space for 50-200 passengers and crew, depending on mission duration, and a pair of detachable pods for 500 tons of express cargo, plus service bays and the like.
What this ship can do depends on its drive engine performance. If the drive puts out 2 gigawatts of thrust power — my baseline for a Realistic [TM] nuke electric drive — the ship can reach Mars in three months, give or take. (The sim gave 92 days for a 0.8 AU trip in flat space.) With a later generation drive putting out 20 gigawatts it can reach Mars in a little over a month, or Saturn in eight months.
The general arrangement of this ship is driven by design consideration — a nuclear drive that needs to be a long way from the crew, with large radiators to shed its waste heat; tanks for bulky liquid hydrogen; and a spinning hab section. Most serious proposals for deep space craft in the last 50 years have had more or less this arrangement — the movie 2001 left off the radiator fins, because in those days the audience would have been puzzled that a deep space ship had 'wings.'
A large, long-mission military craft, such as a laser star, might not look much different overall — replace the cargo pods with a laser installation and side-mounted main mirror, and perhaps a couple of smaller mirrors on rotating 'turret' mounts. Discussions here have persuaded me that heavy armor is of little use against the most likely threats facing such a ship.
Within these broad constraints, however, spaceships offer a great deal of design freedom, more than most terrestrial vehicles. Ships, planes, and faster land vehicles are all governed by fluid dynamics, and even movable shipyard cranes must conform to a 1-g gravity field. A spaceship, unless built for aerobraking, will never encounter fluid flow, and the forces exerted by high specific impulse drives — even torch level drives — are relatively gentle.
This ship might have had two propellant tanks, or half a dozen, instead of four. And the entire industrial assemblage of tanks and girders might be concealed, partly or entirely, within a 'hull' of sheeting no thicker than foil, protecting tanks and equipment from shifting heat exposure due to sunlight and shadow. Much of the ISS keel girder has a covering of some sort — in close-ups it looks a lot like canvas — that in more distant views gives the impression of a solid structure.
In fact the visual image of the ISS is dominated by its solar wings and radiators. The hab structure is fairly inconspicuous by comparison, like the hull of a sailing ship under full sail. This would be true to an extreme of solar electric ships; a 1-gigawatt solar electric drive would need a few square kilometers of solar wings. Even nuclear drives, fission or fusion, require extensive radiators — probably more than I showed — with other ship systems needing their own radiators, at varied operating temperatures. Unless the ship has an onboard reactor it must also have solar collectors for use when the drive is shut down.
All of which may do more to catch the eye than heavier but smaller structures such as the hab or even propellant tankage. And then there is color: the gold foil of the main ISS solar wings, for example.
Hollywood knows nothing of this (though I'm surprised they haven't picked up on the gold foil). Hollywood is no more interested in what real spaceships look like than it is in how they maneuver. This is only natural, even though we hard SF geeks complain. Hollywood doesn't care because its audience has almost no clue of what spaceships look like, or act like, getting most of their impressions from Hollywood itself.
The one actual spacecraft to have iconic visual status, the Shuttle, essentially looks like an airplane. The ISS has not yet acquired iconic status, though it may, especially after the Shuttle is retired. And perhaps it looks so unlike terrestrial vehicles that our eye does not yet know quite what to make of it.
As a point of comparison, watch aviation scenes in old movies, especially from before World War II. You'll see airplanes whooshing past (sometimes in pretty unconvincing special effects shots), but you will rarely see what is now a standard shot — a plane filmed from another plane in formation, hanging 'motionless' on the screen, clouds and distant landscape rolling slowly past, until perhaps the plane banks and turns away.
It is a standard shot because it is so very effective. But older movies rarely used it, because audiences would have had no idea what they were seeing. Everyone knew that airplanes were fast, and had at least some idea that their speed is what kept them in the air. A plane apparently hanging in midair would make no sense.
What changed all this, I would guess, is World War II. A flood of newsreel footage included many formation shots, and audiences gradually absorbed a feeling for what midair footage really looks like. When a postwar Jimmy Stewart enlisted for Strategic Air Command (1955), Hollywood — and its audience — were ready to see the B-36 and B-47 showcased in all their glory, including airborne formation shots.
I know what you bloodthirsty people are thinking — one good space war, and everyone will grok the visual language of space travel. Shame on you. Given enough civil space development, and time, people will get the hang of it.
The beauty of spaceships is in the eye of the beholder. The familiar aesthetics of terrestrial vehicles are as irrelevant to them as to Gothic cathedrals (which in some broad philosophical sense are themselves spaceships of a sort). General principles of design will provide some guidance. Even in making the quick thrown-together model above I found that slight changes in proportion could make the difference between a jumble of parts and a unity.
But the real visual impact of spaceships is something we will only learn from experience, by the glint of a distant sun.
|Exhaust Velocity||4,400 m/s|
|Specific Impulse||449 s|
|Thrust Power||7.7 gigawatts|
|Total ΔV||6,100 m/s|
|Engine Mass||7 mton|
|Heat Shield Mass||15 mton|
(15% re-entry mass)
(5% landing mass)
(5% landing mass)
(20% dry mass)
(5% dry m)
|Tankage body||18 mton|
|INERT MASS||75 mton|
|DRY MASS||100 mton|
|WET MASS||400 mton|
This is a splendid spacecraft designed by Rick Robinson, appearing on his must-read blog Rocketpunk Manifesto. This was designed for his Orbital Patrol service, which he covered in three previous posts.
The important insight he noted was that if you can somehow get your spacecraft into orbit with a full load of fuel/propellant, it turns out that most cis-Lunar and Mars missions have delta V requirements well within the ability of weak chemical rockets. So you make a small chemical rocket and lob it into orbit with a huge booster rocket (heavy lift launch stack). This will be the standard Orbit Patrol ship.
It can also be boosted into orbit by a smaller booster rocket, then using the patrol ship's engines for the second stage. So as not to cut into the ship's mission delta V, it will need access to an orbital propellant depot to refuel. At a rough guess, you'll need 9,700 m/s delta V to boost the patrol ship into orbit (7,900 m/s orbital velocity plus gravity and aerodynamic drag losses). So the booster will need 9,700 m/s with a payload of 400 metric tons. Bonus points if the booster is reusable.
Actually, it reminds me a bit of the old Three Man Space Scout.
At a rough guess, Rick figures that if the ship is capsule shaped it will be about 12 meters high by 14 meters in diameter. If it is wedge shaped, it will be about 40 meters high by 25 meters wide by 8 meters deep.
In both cases, total interior volume of 1,200 m3 (of which 900 m3 is propellant), and a surface area of 800 m2
Present day expandable propellant tanks have a mass of about 6% of the mass of the liquid propellant. Rick is assuming that in the future the 6% figure will apply to reusable tanks as well.
If my slide rule is not lying to me, the 300 metric tons of H2-O2 fuel/propellant represents 33.3 metric tons of liquid hydrogen and 266.7 metric tons of liquid oxygen. About 470 m3 of liquid hydrogen volume (sphere with radius of 4.8 m) and 234 m3 of liquid oxygen volume (sphere with radius of 3.8 m). This is a total volume of 704 m3 which falls short of Rick's estimate of 900 m3 so I probably made a mistake somewhere.
Landing on Terra will use retro-rockets, the heat shield for aerocapture, maybe a parachute, and aircraft style landing gear for belly landing. Landing on Luna or Mars will be by tail-landing on rear mounted landing legs. That will also mean reserving some of the propellant for landing purposes.
Note that the heat shield is rated for the ship's unfueled mass (heat shield mass = 15% of ship's re-entry mass), there is not enough to brake the ship if it has propellant left. This assumes a "low-high'low" mission profile: start at LEO, go outward to perform mission while burning most of the propellant, then return to LEO or even land on Terra. So 15 metric tons for heat shield is for a ship with a mass of 100 metric tons at re-entry (ship's total dry mass).
If the ship is going to aerobrake then return to higher orbit, it will need more heat shield mass to handle the extra mass of get-home propellant. This will savagely cut into the payload mass, which is only 25 metric tons at best. For example, if the mission had the ship heading for translunar space from LEO after aerobraking, the extra propellant mass at aerobrake time will increase the heat shield mass from 15 metric tons to 31. This will reduce the payload from 25 metric tons to 8. But by the same token a ship that will not perform any aerobraking can omit the heat shield entirely, using the extra 15 metric tons for more propellant or payload.
Payload includes habitat module (if any) as well as cargo, since hab modules are optional for short missions. The gross payload is 25 metric tons, of which 20 is cargo and the other 5 mtons are payload bay structure and fittings. If you assume two tons of life support consumables per crew per two week mission; then the ship could carry a crew of five plus 12 mtons of removable payload, or a crew of 10 and 4 mtons of payload (the more that payload is consumables, the less mass needed for payload bay structure).
|Low earth orbit (LEO) to geosynch and return||5700 m/s powered|
(plus 2500 m/s aerobraking)
|LEO to lunar surface (one way)||5500 m/s|
|LEO to lunar L4/L5 and return|
|4800 m/s powered|
(plus 3200 m/s aerobraking)
|LEO to low lunar orbit and return||4600 m/s powered|
(plus 3200 m/s aerobraking)
|Geosynch to low lunar orbit and return|
|Lunar orbit to lunar surface and return||3200 m/s|
|LEO inclination change by 40 deg|
|LEO to circle the Moon and return retrograde|
|3200 m/s powered|
(plus 3200 m/s aerobraking)
|Mars surface to Deimos (one way)||6000 m/s|
|LEO to low Mars orbit (LMO) and return||6100 m/s powered|
(plus 5500 m/s aerobraking)
Representative samples of small space craft atop booster rockets:
Payload Crew 25 Hab Module 100 tons Consumables 25 tons Other Payload 75 tons Total Payload 200 tons Propulsion Bus Engine+Radiator 200 tons Tankages+Keel 100 tons Stats Dry Mass 475 tons Loaded Mass 500 tons Propellant Mass 500 tons Wet Mass 1000 tons
The discussion thread about 'Industrial Scale of Space' veered, among other things, into a discussion of patrol missions in space. My first reaction was that (so long as you aren't dealing with an interstellar setting) there is no place in space for wartime patrol missions. But the matter might be more complicated, and for story purposes probably should be.
According to The Free Dictionary, patrol is The act of moving about an area especially by an authorized and trained person or group, for purposes of observation, inspection, or security. This fits my own sense of the word, and is in fact a bit broader, 'security' including SSBN patrols, which are not observing or inspecting anything, just waiting for a launch order if it comes.
In a reductionist way you could say that all military spacecraft are on patrol, since they are all on orbit, and if they are orbiting a planet they have a very regular 'patrol area.' But this is not what most of us have in mind. We picture a patrol making a sweep through an area, looking for anything unusual, ready to engage any enemy they encounter, or report it and shadow it if they cannot engage it.
Back in the rocketpunk era it was plausible that, say, Earth might send a patrol past Ceres to see if the Martians had established a secret base there. But (alas!) telescopes 'patrolling' from Earth orbit can easily observe the large scale logistics traffic involved in establishing a base; watch it depart Mars and track it to Ceres. If you want a closer look you can send a robotic spy probe. If you engage in 'reconnaissance in force' by attacking Ceres, that is a task force, not a patrol.
In an all out interplanetary war there may be plenty of uncertainty on both sides, but very little of it can be resolved by sending out patrols.
But of course all-out war is not the context in which the Space Patrol became familiar. I associate it with Heinlein's Patrol; apparently the 1950s TV series had an independent origin (unlike Tom Corbett, who was Heinlein's unacknowledged literary child).
The rocketpunk-era Patrol, which in turn gave us Starfleet, was placed in the distinctly midcentury future setting of a Federation. This is as zeerust as monorails. But plausible patrolling is not confined to Federation settings. It can justified in practically any situation but all out war.
Orbital patrol in Earth orbital space will surely be the first space patrol, and could be imagined in this century. It might initially be a general emergency response force, because travel times in Earth orbital space are short enough for classical rescue missions. On the interplanetary scale, with travel times of weeks or more likely months, rescue is rarely possible. But eventually power players will want some kind of police presence or flag showing in deep space.
As so often in these discussions, I picture a complex and ambiguous environment in which policing, diplomacy, and sometimes low level conflict blur together. To take again our Earth-Mars-Ceres example, there are kinds of reconnaissance that cannot be carried out by robots (short of high level AIs). If Ceres closes its airlocks to liberty parties from a visiting Earth patrol ship, that conveys some important intelligence information.
The ships that perform these missions will be fairly large (and expensive). They must carry a hab pod providing prolonged life support for a significant crew: at least a commander and staff, SWAT team of espatiers, and some support for both.
Let us say a crew of 25—which is cutting the human presence very fine. Now we can venture a mass estimate. Allow 100 tons for the hab compartment plus 25 tons for crew and stores plus 75 tons other payload, for a total payload of 200 tons. Let the drive bus be 200 tons for the drive, including radiators, and 100 tons for tankage, keel, and sundry equipment.
Our patrol ship with a crew of 25 thus has a dry mass of 475 tons, mass fully equipped 500 tons, plus 500 tons propellant for a full load departure mass of 1000 tons. Cost by my usual rule of thumb is equivalent to $500 million, perhaps $1 billion after milspecking, expensive compared to military planes, cheaper than major naval combatants.
This is no small ship. If the propellant is liquid hydrogen the tanks have a volume of about 7000 cubic meters, equivalent to a 7000 ton submarine. The payload section is about two thirds the mass of the ISS and of roughly comparable size, though the hab is probably spun giving the prolonged missions.
Armament is necessarily modest. The 75 tons of additional payload allowance probably must include a ferry craft for the espatiers and an escort gunship or two, plus their service pod, leaving perhaps 15-20 tons each for kinetics and a laser installation. The laser might be good for 20 megawatts beam power, with plug power from the 200 megawatt drive engine.
This ship is no laser star, but the laser is respectable. Assuming a modest 5 meter main mirror and a near IR wavelength of 1000 nanometers, at a range of 1000 km it can burn through Super Nano Carbon Stuff at rather more than 1 centimeter of per second. Its armament is also rather 'balanced.' My model shows that this laser can just defeat a wave of about 1000 target seekers, each with a mass of 20 kg, closing at 10 km/s—thus a total mass of 20 tons, comparable to its kinetics payload allowance.
Deploying troops, or personnel in general, is impressively expensive: About three fourths of the payload and cost of a billion dollar ship goes to support and equip a crew of 25, with perhaps a dozen espatiers. For comparison the USS Makin Island (LHD-8) displaces 41,000 tons full load, carries a crew of 1200 plus 1700 Marines, and costs about $1.8. So by my model it costs about as much to deploy one espatier as 80 marines.
And this ship is about the minimum patrol package, so standing interplanetary patrol is a costly and somewhat granular business, something not everyone can afford.
Rocketpunk Patrol Ship
Dry Mass 76.2 metric tons Wet Mass 384.6 metric tons Mass Ratio 5 Length Z 73 meters Length Y 20.1 meters Length X 15.2 meters Engine x2 F-26-A LH/LOX Thrust 7.7×106 N Acceleration 0.5 g ΔV 8,200 m/s
This is the same one from the other day, only dressed up with a nice logo and some stats. These are realistic capabilities made courtesy of the charts and other information available from Atomic Rocket and inspiration from Rick Robinson's Rocketpunk Manifesto.
My PL differs from the one in Rick Robinson's article in a few key areas. The main difference is that it is not made for long hauls. It only has a delta v of about 8200 m/s. This will not get one far in the solar system but it allows a forward deployed Patrol Craft a sufficient "range" to perform many of the missions we discussed in the last post on Building a Space Navy. Our little A-Class has enough Delta V to shape a light-second orbit around a convoy in deep space, conduct SAR missions anywhere in cis-lunar space, or to reach any moon of Saturn from any other moon. Obviously, this rocket is mostly propellant (mass ratio 5). If you drew lines through the side view of the rocket that bracket the docking rings, you would encompass the entire pressurized volume. I've got to say, it's nice to work on a warship for a change — I don't have to make it economical to run!
One of the interesting things about this design is actually the freedom the little carried craft gives me. It was a throw-away touch, originally — a design borrowed from another project. But as I got to looking at the little thing, I realized that it's about the size of the Saturn V stage/Apollo/LM stack. That means it should be able to go from Earth Departure to Lunar orbit. That means that it has the Delta V to ferry crew to and from a Patrol Craft on station away from the convoy. That means, like submarines, our Patrol Craft can have two crews and stay out for a lot longer than otherwise. This is one of those realistic touches that I hope add to the charm of the rocket's design.
ed note: a 1500 nanometer near infrared laser with a 10 meter fixed mirror can have a 4 centimeter spot size out to 220 kilometers or so. A 4 meter mirror can have a 4 centimeter spot size out to 87 kilometers or so.
Round the Moon Ship
|Total ΔV||≅6,120 m/s|
|Total Thrust||2,250,000 N|
|Max Width||8 m|
|Dry Mass||≅50,290 kg|
|Wet Mass||≅412,378 kg|
Wernher von Braun's Round the Moon Ship first appeared in the famous Collier's Man Will Conquer Space Soon! series (and later collected in the book Across the Space Frontier). You can find it in PDF form here in the Horizons Newsletter July/August 2012 Issue on page 60. The spacecraft became sufficiently iconic that it was plagiarized for the "Space Age" poster.
The main thrust of the first half of the Collier's series was a large expedition to Luna. First there was a large ferry rocket used like a space shuttle to transport pre-fab section of a space station into orbit. The space station would then help assemble the fleet of huge ships for the lunar expedition.
Now it would be real nice if a tiny ship could be sent in advance to scout out some promising landing sites for the big lunar expedition. It would be most unfortunate if the expedition landed in a field of huge dagger-like rocks and everybody died. The scout did not have to land, just make a close orbital pass and take lots of photos. Which means the scouting spaceship does not need any landing legs.
For such a scouting mission von Braun wanted something quick-and-dirty. He remembered that the third stage of the ferry rocket (the part that actually reached orbit) had a cluster of five rocket engines. So the idea was to cannibalize the cluster from one of the ferrys floating in orbit and build on top a flimsy cage made out of as few low mass girders as he could get away with. The cage would be a spaceframe, the base of the cage resting on the cluster is the thrust frame. Then hang off the spaceframe some super low mass fuel tanks and hab modules which were little more than large balloons. One quick-and-dirty spaceship, coming right up.
Everything had to be low mass because the Hydrazine/Nitric Acid chemical engine had a truly pathetic specific impulse of 328 seconds at best, and von Braun was assuming the engines would actually manage barely 296 seconds. It's a good thing that the scout doesn't need landing legs, those things are heavy.
Why did von Braun use Hydrazine/Nitric Acid instead of something more powerful? William Seney did some research:
First off, Hydrazine/Nitric Acid is not cryogenic, which means it will stay in the fuel tanks indefinitely without needing electrical cooling. The alternatives all required liquid oxygen (LOX) which is regrettably cryogenic.
Secondly, the Round the Moon Ship design dates from 1952. The only other fuel that was in active use at that time was LOX/Alcohol, with a barely better specific impulse of 338 s, compared to Hydrazine/Nitric Acid's 328s.
LOX/RP-1 has a specific impulse of 353 s, but work was not done on it until 1953, and it didn't fly until the late 1950's. LOX/Liquid Hydrogen has a great specific impulse of 451 s, but it didn't fly until the early 1960's.
The top of the spacecraft had the inflatable habitat module with an airlock hanging off the bottom. Below were the inflatable hydrazine fuel tank and the inflatable nitric acid oxidizer tank. Each tank had an associated compressed nitrogen tank. The nitrogen kept the tank pressurized, encouraging the fuel to flow to the engines.
All three inflatables had several square arrays of passive thermal control slats. If a sphere got too cold, black slats would deploy to suck up the Sun's heat. If a sphere got too hot the black slats would retract, revealing the mirrored surface which rejects the Sun's heat (alternatively they may be like Venetian blinds with one side black and the other mirrored). Looking at the illustrations I count about 12 slat arrays per sphere.
Near the bottom was a torus (donut) shaped hydrogen peroxide tank. This the fuel that runs the Walter turbines, which pumps the rocket fuel at high speed into the rocket engines.
Each engine produces 450 kiloNewtons of thrust, the five engine cluster produces a total of 2,250 kiloNewtons. The four outer engines are swivel mounted to allow the spacecraft to be steered. The center engine is fixed.
The spaceframe sports a single radar/communication dish antenna aimed at Terra. On the opposite side (for balance) is the solar mirror/mercury boiler power plant, used because photovoltaic solar cells arrays have not been invented yet. According to Roger's Blueprint, the solar mirror has an aperture of 1.2×6.5 = 7.8 square meters. At Terra's distance to the sun, solar energy is about 1366 watts per square meter, so the aperture is admitting about 10.6 kilowatts. von Braun was assuming the mercury boiler was about 28% efficient, giving an output of 2.97 kW.
But according to the best figures I've manage to find, von Braun was being wildly optimistic. A mercury boiler is lucky to be 11% efficient, giving the power plant a wretched 1.17 kilowatts of output. If you retro-fit a NASA standard photovoltaic array of the same area you'd get more like 3.07 kW.
Finally there were four oddly-shaped storage compartments squeezed into the oddly-shaped free space between the hydrazine and nitric acid tanks.
Now it is time for me to do some pointless playing around with numbers.
What von Braun wanted for this mission is a "free-return trajectory". The spacecraft starts in low Terra orbit, does a specfic maneuver with the rocket engines, the spacecraft then falls along a large figure-8 trajectory looping around Luna and eventually arriving back at Terra Orbit with no further rocket burn required.
NASA used the free-return trajectory for the Apollo missions as insurance. If the Apollo SM main engine broke the spacecraft would automatically return to Terra, instead of sailing off into the big dark with the destination being a lonely death for the astronauts and a public-relations nightmare for NASA. Which paid off big-time with Apollo 13, when the SM main engine did break.
According to figure 9 on page 16 of Trajectories in the earth-moon space with symmetrical free return properties, the lowest delta V you can manage for a circumlunar (not cis-Lunar) free return is about
10,860 meters per second 6,120 meters per second.
(ed note: William Seney set me straight on that point. 10,860 m/s includes boosting from Terra's surface into LEO, which is not needed with this mission profile. 6,120 m/s is 3060 m/s to leave orbit and another 3060 m/s to break back into orbit on return, no aerobraking required.)
Close enough for a back-of-the-envelope estimate (yes, kids, envelopes were paper containers for letters, which were physical emails people used to send in olden days. Engineers would use them as impromptu calculation scratch pads).
A hydrazine-nitric acid chemical engine has an abysmal specific impulse of 328 seconds, and von Braun figured the ferry rocket third-stage cluster would be lucky to get 296 seconds. This implies an exhaust velocity of 2,900 m/s.
Delta V is 10,860 m/s (ignoring braking into Terra's orbit at the end, assume a rescue ship). Mass ratio (R) is equal to e(Δv/Ve) which comes out to a truly ugly 29.2. Which is pretty bad, since one generally does not see a mass ratio above 4.0 without multistaging. A mass ratio of 29.2 means the spacecraft will have to be made out of foil and soap bubbles. (again William Seney showed the 10,860 m/s figure is incorrect. )
Roger's Blueprint say both the hydrazine fuel tank and the nitric acid oxidizer tank have a diameter of 6.5 meters, implying a volume of 143.8 cubic meters (less the bubble-skin walls). Given the densities the hydrazine tank has a mass of 144,950 kilograms and the nitric acid tank at 217,138 kilograms. Total is 362,088 kilograms, which is the spacecraft's fuel mass (Mpt).
The spacecraft's dry mass (with empty fuel tanks) is equal to Mpt / (R -1) which comes out to...
a miserly 12,840 kg or only 12.8 metric tons. Including crew and life-support. Spacecraft's wet mass is 374,928 kg or 375 metric tons
...50,290 kg or 50 metric tons. Spacecraft's wet mass is 412,378 kg or 412 metric tons.
|Rotating Bed Rocket|
|Fuel Mass||140 kg|
This is from Advanced Propulsion Systems Concepts For Orbital Transfer Study, vol I and vol II (Boeing documents D180-26680-1 and D180-26680-2). Additional information from French Wikipedia entry for nuclear thermal rocket (missing from English Wikipedia).
The study was an attempt to find advanced propulsion alternatives to the standard hydrogen-oxygen chemical rocket. It studied all sorts of systems, including solar powered ion, laser thermal, fusion, nuclear lightbulb, magnetothermodynamic, and others.
It found several systems worthy of study, but there was only one feasible propulsion was both better than LH2/LOX and suitable for use for manned missions: the Rotating Fluidized-Bed Nuclear Rocket (RBR). The others either had too low a thrust for manned missions or were considered not feasible (too long a timeline before useable hardware became available).
The core of the engine is a rotating drum (the "rotating structure") which is made out of a porous material with the high-tech name of "frit." It is encased in a squirrel cage type support structure.
Inside the drum is 140 kilograms of fissionable uranium 235 fuel pebbles, coated with zirconium carbide like an M&M candy is coated with a hard candy shell. This prevents the uranium from vaporizing and escaping into the exhaust plume, leaving a trail of glowing blue radioactive death. "Melts in your mouth, not in your hands".
The frit drum is spun with enough rpms (about 1000 r/min) to generate sufficient artificial gravity to stick the fuel pebbles to the frit, instead of floating aimlessly in free fall. The hydrogen propellant is injected through the squirrel cage and poros frit with enough velocity to "fluidized" the fuel pebbles (lift and separate particles). The propellant is heated by passing through the fissioning fuel pebbles, then goes shooting through the exhaust nozzle producing thrust. It is easy to adjust the pebble bed to match any desired propellant mass flow rate by simply altering the spin rate of the frit drum.
Why are we bothering with such a Rube-Goldberg contraption? Because:
- Since the fuel pebbles are from 100 to 500 μm in diameter (dust sized), the total fuel mass has an astronomically high surface-area-to-volume ratio, especially compared to NERVA and other solid core nuclear thermal rockets.
This makes the pebble bed super efficient at transferring the fission heat from the fuel into the gaseous propellant.
Bottom line: the pebble bed engine will have a much smaller reactor core size than pretty much any other nuclear thermal rocket, much lower mass as well.
- For the same reason: while the propellant will become very hot, the squirrel cage and other supporting structure will stay cold. Since the fuel pebbles are fluidized, they are not actually touching the frit, the only thing they touch is propellant. This is not the case with other NTRs.
This means the pebble bed design does not have to worry about thermal stress and other factors that plague other NTR designs. The only thing that matters is the stabilty of the fuel pebbles (ensure that they do not melt off their coating and let the radioactive uranium escape).
Bottom line: the pebble bed rocket has the highest specific impulse of all solid-core NTRs.
- The fuel and fuel support of a pebble bed is about 1/6th the volume and mass of a conventional solid core NTR. This is because the high surface-area-to-volume ratio allows the heat exchange zone (the layer of fuel pebbles) to be very narrow. This drastically lowers the diameter of of the engine.
Bottom line: it is quite easy to remove the reactor core of a pebble bed rocket for maintenance and to swap out the nuclear fuel. For conventional NTRs it is so difficult that it is more economic to just throw away the entire freaking engine when the fuel elements clog up.
Putting it all together, the 420 megawatt pebble bed engine has an initial thrust-to-weight ratio of 6.5 (because the engine is so low mass). A conventional solid-core NTR is lucky to have a T/W of 2.4.
This advantage grows with higher reactor power levels. A 6.5 gigawatt pebble bed engine with a thrust of 1.8 megaNewtons would have a T/W of 17.0, a corresponding solid-core NTR would be hard pressed to have a T/W of 4.0.
As with all nuclear powered rockets, the major draw-back is the dread spectre of deadly atomic radiation.
The study decreed that for each manned mission, the maximum allowable radiation dose experience inside the crew habitat module was 0.03 Sieverts per mission (3.0 rem).
A standard liquid hydrogen (LH2) propellant tank is shaped like a cylinder with elliptical (√2) end caps (that is, shaped like a hot dog). At the aft end is the nuclear engine, the other has either the habitat module or a second LH2 tank then the habitat module.
As it turns out, if you change the shape of the tank at the nuclear engine end, you can drastically reduce the radiation that penetrates through to the habitat module.
Looking at the graph above, the highest radiation dose is when the nuclear engine end cap is a √2 elliptical, the lowest is when the entire engine side half of the tank has a 10° taper. Why?
- A 10° has a lousy volumetric efficiency, which makes the tank longer if it holds the same amount of propellant, which makes the hab module farther from the nuclear engine, which gives extra radiation shielding due to the inverse-square law.
- The graph below somewhat confusingly indicates that most of the radiation dose happens in the last few seconds of the final engine burn, when the radiation-protecting depth of liquid hydrogen propellant in the tank is at its minimum. The 10° taper tank retains a thick layer of LH2 for a longer period, which reduces the total integrated radiation dose.
In addition to the radiation shielding provided by the LH2 propellant, there are two shadow shields: a 1,220 kg disk on top of the nuclear engine and a 240 kg shield on the bottom of the habitat module. This mass directly reduces the spacecraft's payload mass. Naturally the engineers tried to figure out some kind of trick to reduce the shadow shield size.
They noted that the highest radiation dose happened during the last burn, when the propellant level got low. If they could somehow make it so the crew wasn't present when the last burn happened (and have the spacecraft be autopilot controlled), the shadow shields could have their mass reduced since they would only have to protect against lower doses. But how to remove the crew?
Ah, what if the habitat module ejected from the spacecraft, that would remove the crew.
The problem now is that the last burn is when the spacecraft is approaching Terra, and has to brake into a circular Terran orbit. If the habitat module is separated from the engine, it won't be braked. The habitat module has no engine, adding one would eat up the mass saved by reducing the shadow shield size. How can the hab module brake without an engine?
By using that standard NASA sneaky trick: Aerobraking! Give the habitat module an inflatable ballute and use Terra's atmosphere to brake its excess velocity. Then it can rendezvous with LEO station. Just like the Leonov in the movie 2010.
This will allow the shadow shield to be reduced by 450 kilograms. In addition, the amount of required propellant is reduced by 4,000 kg because when it is time to brake into Terra orbit, the spacecraft will be lighter by an amount equal to the mass of the now-absent habitat module.
|Cargo Tug Slingshot|
|Total ΔV||6,000 m/s|
|Specific Power||1.5 kW/kg|
|Thrust Power||764.4 gigawatts|
|Exhaust velocity||280,000 m/s|
|Wet Mass||512,600 mt|
|Ship Mass||1,600 mt|
|Payload Mass||500,000 mt|
|Dry Mass||501,600 mt|
|Deuterium Fuel||16 mt|
|Initial acceleration||0.01 m/s2|
The Cargo Tug Slingshot is from Jerry Pournelle's short story Tinker. In the story, it rescues the BoostShip Agamemnon.
The spacecraft's spine is a strong hollow tube built to transmit thrust from the aft engines to the fore array. The array is composed of detachable fuel pods of deuterium fuel and cadmium reaction mass. Fuel and remass are fed to the engines through the center of the ship's spine. The cargo goes fore of fuel pod. There are a couple of pods of fuel/remass attached to the hull.
Crew cabins are torus-shaped, arranged around the outside of the spine. Foremost torus is control deck. Next aftwards is living quarters for crew. Next comes deck with office and passenger quarters. Furthest aft is deck with shops, labs, and main entryway to the ship. Entryway doubles as a small store catering asteroid miners, to supplement the ship's income. Decks are connected by airlocks for safety.
|0.6 g U|
These are from the report Affordable Development and Demonstration of a Small NTR Engine and Stage: How Small is Big Enough? by Stanley Borowsky et al (2015). The scientists wanted to promote the development of a right-sized solid core nuclear thermal rocket. Because NASA's budget is always so tight, the scientists wanted an engine that was as small as possible, but no smaller.
They looked at a 33,000 Newton (7.5 klbr) engine design but it was a bit too small to do anything useful, even in a cluster of three. A 73,000 Newton (16.5 klbf) engine on the other hand had quite a few useful possiblities. They called it the Small Nuclear Rocket Engine (SNRE).
The engine uses a graphite composite core, because that allowed them to build on the expertise from the old NERVA program (instead of starting back at square one, "affordable development" remember?).
The scientists explored three different missions that could be performed using a cluster of three SNREs:
- Near Earth Asteroid Mission
- Lunar Cargo Mission
- Lunar Crew Landing Mission
All three missions used a common Nuclear Thermal Propulsion Stage (NTPS). The stage had a cluster of three SNREs, a length of 26.8 meter (6.1 m of which is SNRE) and a diameter of 7.6 meters.
Next came an in-line liquid hydrogen propellant tank, the length of which varied according to the mission. The NEA mission held the in-line tank within a saddle truss, so the tank could be jettisoned.
Finally came the mission specific payload.
All missions start from Low Earth Orbit, where the spacecraft components are boosted via multiple launches and assembled.
In Conventional and Bimodal Nuclear Thermal Rocket (NTR) Artificial Gravity Mars Transfer Vehicle Concepts they refer to a Mars mission using this architecture. This was for the Mars Design Reference Architecture (DRA) 5.0 study. There were two unmanned cargo vessels, and a manned vessel called Copernicus. A variant called Copernicus-B used tumbling pigeon spin gravity.
Near Earth Asteroid Mission
|NEA Mission ΔV Budget|
This would make a quick manned visit to asteroid 2000 SG344. NASA had been eyeballing this asteroid once they discovered it was going to whizz very near by Terra in April 2028. The mission would have a duration of 327 days, including a 7 day lay-over at the asteroid.
The payload includes:
- Orion Multi-Purpose Crew Vehicle (MPCV)
- TransHab module outfitted for 4 crew
- Deployable photovoltaic arrays for primary power
- 2-person Multi-Mission Space Exploration Vehicle (MMSEV)
- Short saddle truss
The aft side of the TransHab is connected to a short tranfer tunnel with two docking ports. The MMSEV is docked to the aftward port.
The spacecraft's initial mass in Low Earth Orbit (IMLEO) is 184.6 tonnes.
The mission will require 5 primary burns which will expend a total of 76.2 tonnes of liquid hydrogen propellant, and 46.2 grams of uranium (15.4 grams per engine, about 0.026% of the 59.6 kilograms contained in each engine).
Lunar Cargo Mission
Mission ΔV Budget
The Lunar Transport System (LTS) is a reusable work-horse. It carries no crew. The second liquid hydrogen tank varies in length depending upon the propellant requirements. The total length varies from 20.7 to 23.7 meters (of which 15.7 to 18.7 is tank, the rest is adaptors, propellant feed lines, electrical connections and RCS).
The maximum payload capacity is 60 tonnes, the mass of a chemical rocket Lunar habitat lander with surface mobility (i.e., it has wheels or walking legs). It takes 72 hours to travel from LEO to Low Lunar Orbit (LLO), and the same time to return. The mission will require 5 primary burns which will expend a total of 74.5 tonnes of liquid hydrogen propellant, and 45.15 grams of uranium (15.05 grams per engine, about 0.025% of the 59.6 kilograms contained in each engine). Once in LEO, the LTS can have its propellant tanks reloaded for a new mission.
Lunar Crew Landing Mission
|Lunar Crew Landing|
Mission ΔV Budget
|LDAV ΔV Budget|
For crewed missions to Luna, a small saddle truss is attached to hold the payload. Inside the truss is an Orion MPCV. On top of the truss is the Lunar Descent/Ascent Vehicle (LDAV). The LDAV is a "heritage" design (i.e., an already created design made in the early 1990's) which carries a crew of 4 plus 5 tonnes of surface payload. The payload is carried in two side pods which swing down after landing. The LDAV has a wet mass of 35.3 tonnes: 2.5 t crew cab, 6.1 t descent/ascent stage, 20.9 t LOX/LH2 propellant, 5 t surface payload, 0.8 t 4 crew and their suits. The crew can operate on the lunar surface for 3 to 14 days with the carried payload, or longer (180 days) if pre-deployed habitat landers are present. The LDAV has enough propellant to return 100 kg of lunar samples along with the crew up to LLO.
Once the vehicle is assembled in LEO, the crew launches in the Orion and docks. The vehicle then departs for Luna. As with the cargo mission, both the trip to Luna and the trip back to Terra take 72 hours.
The mission will require 5 primary burns which will expend a total of 83.2 tonnes of liquid hydrogen propellant, and 50.4 grams of uranium (16.8 grams per engine, about 0.028% of the 59.6 kilograms contained in each engine). Once in LEO, the LTS can have its propellant tanks reloaded for a new mission.
On September 27, 2016 Elon Musk unveiled SpaceX awe inspiring Interplanetary Transport System. This was displayed as part of the SpaceX plan to colonize Mars, but the system could transport explorers all over the entire solar system.
The plan seems grandious, but Mr. Musk has a track record of delivering on his promises.
The system has three components:
- ITS Super-heavy lift launch vehicle
- Interplanetary Spaceship
- ITS Tanker
The ITS launch vehicle is used to boost either the Interplanetary Spaceship or the ITS Tanker into Low Earth Orbit (LEO)
- All three components are reusable and capable of returning to Terra. Including the launch vehicle. This is a huge advantage.
- The launch vehicle has a jaw-droppingly monsterous payload capacity of 300 metric tons if reused. And 550 metric tons if expended.
- The tanks will be autogenously pressurized, using gasified propellant for both tank pressurization and for RCS. Conventional rockets use helium gas for pressurization, which creates problems.
- All of the components use subcooled methane/liquid oxygen propellant. The important point is this propellant can be produced on Mars by using the Sabatier reaction. This creates local propellant depots which dramatically increases the effective delta V of the spacecraft. In-situ Resource Utilization for the win!
This makes up for the fact that CH4/LOX has a much lower Isp than LH2/LOX (382s compared to 450s)
- The Interplanetary Spaceship is designed to allow in orbit refueling. This allows it to burn most of its propellant to climb into LEO, then have its tanks refilled by a series of ITS Tanker launches.
|ITS Launch Vehicle|
|Propellant mass||6,700,000 kg|
|Dry mass||275,000 kg|
|Exhaust Velocity||3,750 m/s|
|Thrust (atmo)||128 MN|
|# Engines (atmo)||x42 !!!|
|Max Diameter||12 m|
When loaded with 300 metric tons of payload, this monster is x1.1 as tall as a Saturn V, and has x3.5 the mass. It uses titanic carbon fiber cryotanks, which SpaceX has already produced examples of (thanks to William Black for this link).
It returns to the landing site, using 7% of its propellant for boostback burn and landing. It guides itself back with the famous SpaceX grid fins.
|Propellant mass||1,950,000 kg|
|Dry mass||150,000 kg|
|Cargo mass||300,000 kg|
to 450,000 kg
|Exhaust Velocity||3,750 m/s|
to 5,430 m/s
|Thrust (vac)||31 MN|
|# Engines (vac)||x6|
|# Engines (atmo)||x3|
|Solar Array||200 kW|
|Max Diameter||17 m|
Remember that if you have orbital refueling, a puny chemical rocket can take you all over the solar system. And remember that boosting from Terra's surface into LEO is halfway to anywhere. This is why one of the most important features of the ITS Spacecraft is its orbital refueling capability.
The ITS Launch Vehicle lofts the spacecraft most of the way to LEO, and the spacecraft expends most of its propellant climbing the rest of the way (about 50 metric tons of propellant left). But then it waits in LEO parking orbit.
There follows a series of five more launches of ITS Tankers. Each one reaches orbit with about 380 metric tons of cryogenic methane and liquid oxygen, used to fill the spacecraft's tanks. Total of 1,900 metric tons, so the spacecraft's tanks are totally filled with 1,950 metric tons.
Since the ITS Launch Vehicle and the ITS Tanker are both reusable, all five launches could be of the same two vehicles.
Using the Oberth effect, the bare minimum delta V needed to leave LEO and enter Hohmann Trans Martian Injection is about 3,600 m/s. It will take 8.6 months (258 days), all the while exposing the passengers to deadly galactic cosmic rays and microgravity damage.
However, an ITS Spacecraft with only 300 metric tons of cargo has almost twice that: 6,280 m/s. It can do a high-energy Hohmann and get there in about 80 to 150 days, a vast improvement. It will only have to reserve a bit of fuel for the last bit of the Mars landing, the bulk of the landing delta V is by aerobraking.
On the Martian surface, it can be refuelled by the on-site Sabatier reaction generators.
|Wet Mass||388 metric tons|
|Mars Lander mass||57 metric tons|
ascent stage mass
|27 metric tons|
|Ion propulsion mass|
|123 metric tons|
|153 metric tons|
|Rotation rate||1 rpm|
|Artificial gravity||0.2 g|
|Spin radius||179 meters|
|Wet mass||309 metric tons|
|226 metric tons|
This is from Ernst Stuhlinger 1966 hybrid NERVA-Ion Mars mission proposal.
The idea is to avoid the drawback of the ion drive, the fact that the pathetic thrust of around 100 Newtons means it had an equally pathetic acceleration of about 0.0001 meters per second. Ordinarily this would not be a problem, except it means the spacecraft takes over twenty days to crawl through that glowing blue field of radioactive death they call the Van Allen Belts. A NERVA style nuclear thermal rocket can zip through the belt in a couple of hours, but its abysmal exhaust velocity makes it a propellant hog.
Stuhlinger's plan was a two-stage spacecraft. The NERVA-II stage gets the spacecraft through the radiation belt before the astronauts are fried, then that stage is ditched. The ion drive with its vastly superior exhaust velocity then takes over and gets the expedition to Mars using only a tea-cup's worth of propellant.
In Phase 1, for each of the four spacecraft in the expedition, 3 Saturn V will boost the ion-drive stage components into orbit, where the components will be assembled (12 Saturn V launches total).
In Phase 2, for each of the four spacecraft, 2 Saturn V will boost the NERVA components into orbit (one for the NERVA, one for the propellant tank), where the components will be assembled (8 Saturn V launches total). The NERVA stages will be attached to the ion stages.
There are four spacecraft in the expedition, in case one or more have to be abandoned for whatever reason. In a pinch a single spacecraft can carry all 16 expedition members home, abet in cramped conditions.
The mission starts with the crew inside the landers. If anything goes wrong during the initial burn, the landers will be the crew's abort-to-Terra vehicles. The NERVA-II stage burns for 30 minutes, passing through the Van Allen belts in 2 hours. About 17 minutes into the burn, exhaust is vented to spin up the spacecraft to 1 revolution per minute, for artificial gravity. The burn terminates when the spacecraft is at an altitude of 3450 kilometers.
The crew leaves the lander, and climbs down the 179 meter arms to the habitat modules. The NERVA stage is jettisoned, and the ion engines are started. They will burn for a while, then the ship will coast.
145 days into the mission, the ion engines are restarted to decelerate into high Mars orbit. The crew enters the Mars lander and land on Mars.
The unmanned spacecraft will continue the ion burn 24 days to move the ship to a low 1000 kilometer orbit. It would take even longer if the spacecraft had to deal with the mass of the lander.
After a month on Mars frantically doing sciene, the crew enters the lander's ascent stage and blast of to rendezvous with the orbiting ion spacecraft. The ascent stage is discarded to save on mass. This allows the spacecraft to spiral out to Terra transfer orbit in only 18 days.
The trip home will take 255 days, with deceleration starting halfway through.
|1st stage ΔV||2,440 m/s?|
|1st stage Specific Power||4.3 kW/kg|
|1st stage Propulsion||Chemical, plug nozzle|
|1st stage Fuel||LO2/LH2|
|1st stage Specific Impulse||382 to 439 s|
|1st stage Exhaust Velocity||3,750 to 4,310 m/s?|
|1st+2nd stage Wet Mass||10,900,000 kg|
|1st+2nd stage Dry Mass||5,940,000 kg?|
|1st stage Mass Ratio||1.83?|
|1st stage Mass Flow||3,160 kg/s?|
|1st stage Thrust||13,600,000 n|
|1st stage Initial Acceleration||1.25 g?|
|Staging velocity||2,440 m/s|
|2nd stage ΔV||19,500 m/s?|
|2nd stage Specific Power||28 kW/kg|
|2nd stage Propulsion||OC Gas Core NTR|
|2nd Engine size||3500K|
|2nd Number of engines||4|
|2nd stage Specific Impulse||2,000 s|
|2nd stage Exhaust Velocity||19,600 m/s?|
|2nd stage Wet Mass||5,940,000 kg|
|2nd stage Dry Mass||2,190,000 kg?|
|2nd stage Mass Ratio||2.7?|
|2nd stage Mass Flow||324 kg/s?|
|2nd stage Thrust||6,350,000 n|
|2nd stage Initial Acceleration||1 g?|
|Total Wet Mass||10,900,000 kg|
|total ΔV||21,800 m/s|
|Total height||134 m|
|1st stage Diameter||45-52 m|
|2nd stage Diameter||36 m|
This is a heavy-lift vehicle designed to boost absurd amounts of payload from the surface of Terra, using deadly open-cycle gas-core nuclear thermal rockets in the second stage. If you want all the hard details, run and purchase a downloadable copy of Aerospace Projects Review vol. 3 no. 1. You get a lot of info for your downloading dollar.
This monster is the Uprated GCNR Nexus grown to three times the size. The document says that it can deliver 453 metric tons (one million pounds) not to LEO, but to Lunar surface. Doing some calculations on the back of an envelope with my slide rule, I estimate that it can loft 4,600 metric tons into LEO. But also with a proportional increase in radioactive exhaust. The data in the table is for the Terra lift-off to Lunar landing mission.
|Dry Mass||1,016,000 kg|
|Wet Mass||6,113,600 kg|
In his novels Michael McCollum postulates lots orbital antimatter factories that in one year will consume outrageous amounts of energy and produce 25 miserable kilograms of antihydrogen, conveniently packaged in a magnetic torus to prevent it from touching any normal matter and blowing everything to tarnation. These are useful for moving valuable ore-rich asteroids into Terra orbit. And as fuel for antimatter torchships.
Mr. McCollum stated the following:
- Antimatter Gas Core Engine
- Fuel: 4.5 grams of antihydrogen
- Propellant: Liquid hydrogen
- Ship carries eighteen propellant tanks each carrying 4,000 cubic meters, total 72,000 m3
- Reaction chamber temperature: 100,000 degrees R, which according to the table in TAOSF vol 1 corresponds to a specific impulse of 5,680 seconds and an exhaust velocity of 55,720 m/s
- Ship can make the trip from Terra to Jupiter in six months (whereas a Hohmann transfer is more like six years)
- Ship has a delta V of 100,000 meters per second
Thus endeth the canon knowledge.
The other figures are me playing with numbers.
R = mass ratio
ΔV = transit delta-V (m/s)
Ve = exhaust velocity (m/s)
ex = antilog base e or inverse of natural logarithm of x, the "ex" key on your calculator
Delta-V is 100,000 m/s, exhaust velocity is 55,720 m/s, so the mass ratio is 6.0176
There is 72,000 m3 of liquid hydrogen propellant. Liquid hydrogen has a density of 70.8 kg/m3 so the total propellant mass Mpt is 5,097,600 kg.
Me = Mpt / (R - 1)
Me = dry mass (kg)
Mpt = propellant mass(kg)
R = mass ratio
Propellant mass is 5,097,600 kg and mass ratio is 6.0176 so dry mass is 1,016,000 kg
M = Me + Mpt
M = wet mass
Me = dry mass (kg)
Mpt = propellant mass(kg)
Dry mass is 1,016,000 kg and propellant mass is 5,097,600 kg so wet mass is 6,113,600 kg
Playing around even more, I took the ship diagram as a blueprint into the Blender 3D modeling program. The diagram had a bar labeled as 100 meters long, so I scaled the model to that.
The hydrogen tanks were stated as canon to have a volume of 4,000 cubic meters each. Mathematically this meant they had a diameter of about 19.7 meters, which matched the blueprint reasonably closely. Adding the habitat module gave me a ballpark figure of it being 27 meters in diameter with a volume of 10,000 cubic meters. The main body had a diameter of 27.8 meters, a height of 33.6 meters, and a volume of 21,000 cubic meters (assuming it is a cylinder). The total length was about 150 meters.
Since these figures are from playing around with a quickly done diagram (which does not agree with the cover illustration very well), I would not put too much faith in them.
This is from NUCLEAR SPACE PROPULSION by Holmes F. Crouch (1965). It is a solid-core nuclear thermal rocket used by the outer space version of the Coast Guard to rescue spacecraft in distress. In the diagram below, note how the rear fuel tanks are cut at an angle. This is to prevent any part of the tank from protruding outside of the shadow cast by the nuclear shadow shield. Also note that while the central tank must be load-bearing, the strap on tanks do not. This means the side tanks can be of lighter construction.
|TRW Mars Mission|
|Outbound time||200 days|
|Mars Stay time||10 days|
|Return time||220 days|
|Total time||450 days|
This is from a 1963 study done by TRW for a NASA Ames contract (all the tedious details can be found in NASA-TM-X-53049. The only reason I found that document was by reading David Portree's well worth reading Humans to Mars: Fifty Years of Mission Planning, 1950-2000).
The contract was to develop a manned mission to Mars using non-nuclear propulsion. Chemical propulsion means the spacecraft would need its mass drastically reduced, and the required delta V lowered by quote "innovative mission scenarios" unquote.
TRW figured out how to lower the spacecraft mass by a whopping factor of five! The major mass reduction came from using aerobraking instead of thrusters at both Mars and Terra (assuming a Martian surface pressure of 10% Terran). Delta V requirements for the return trip were obtained by having the ship do a gravity assist at Venus instead of heading directly to Terra.
A conventional mission using rocket thrust for braking would have a mass of around 3250 metric tons, TRW's design was only 650 metric tons.
The mission was a fast opposition-class, with a duration of 400 to 450 days but only ten days spent on Mars. See "The Short-Stay Mission".
Six or seven Saturn V launches are required to boost all the spacecraft components into orbit, where they are assembled (see diagram). In one variant, a single launch is for the monolithic Earth departure engine (containing no fuel) and the other four are tanker spacecraft to fuel the monolithic engine's tanks. In another variant four launches are four modular Earth departure engines with full tanks, which are assembled into the engine unit. The monolithic engine variant has the advantage of assembling the spacecraft using simple docking, and the disadvantage of the nightmare of free-fall propellant transfer. The modular engine variant has the advantage of avoiding free-fall propellant transfer, and the disadvantage of the nightmare of free fall component assembly.
One variant uses conventional liquid oxygen/liquid hydrogen fuel. If you look closely at the blueprints below you will notice in that variant the oxidizer tanks are not labeled with "LO2 (liquid oxygen) but rather with LF2 (liquid fluorine!?!!). A designer uses fluorine oxidizer only if they are really desperate for delta V, that stuff is unbelievably dangerous.
The command station doubles as the storm cellar. The radiation shielding is basically a huge tank of hydrazine (N2H4) fuel enveloping the command room. The hydrazine is borrowed from the Earth re-entry module deorbit engine fuel tanks. There are about twenty other variants, using different shielding material and covering different areas. One actually has no storm cellar, just bloated water balloon suits, one for each crew member.
Spacecraft uses a bola artificial gravity system (see diagram). The spin radius is 22.86 meters (75 feet), the spin rate is 2.56 RPM giving 0.167 g of artificial gravity (1/6 g or one Lunar gravity). The cable is 136 meters long even though the ship's spin radius is 22.86 meters because the center of rotation is quite far away from the geometric center. This is because the spacecraft has a mass of 213 metric tons but the counterweight is only 32 metric tons.
During the Terra-Mars transit, the counterweight for bola spin is the spent Earth departure engine. During the Mars-Terra return transit, the spacecraft splits into two parts. The lower section (the "exhausted Mars departure stage") becomes the counterweight.
Meanwhile from the base of the Mars Departure Stage are deployed two solar power panels. In one variant they are solar thermal collectors, another variant uses solar photovoltaic arrays. You can see the solar photovoltaic arrays here in dark blue, note how they are hinged at the edge so they can flip outwards. The solar thermal collectors can be seen here.
In both designs there are two solar arrays each with a collecting surface of 70 square meters. As a rough guess, while at Mars the solar thermal will generate about 9 kilowatts and the photovoltaic will generate 24 kilowatts (583 w/m2 at Mars, 140 m2 of collector, thermal is 11% efficient, photovoltaic 29%).
The report says the spacecraft requires 5 kilowatts: 2.6 for the life support system managing 6 crew members (water and air regenerated), and 2.0 kilowatts for television transmissions between Mars and Terra.
Instead of using rocket thrust, spacecraft maneuvers into an elliptical Martian orbit via aerobraking. Gotta get that ship design mass down somehow. The solar arrays and antenna are retracted first, obviously, or they will be torn off. The spent Earth departure engine is jettisoned and the bola cable is reeled in. Once orbit is achieved, a little bit of rocket thrust is used to raise the perigee of the orbit above the top of the atmosphere.
After surveying the surface, a landing site is selected and the Mars Excursion Module transports two crew members for a ten day exploration of said site. At the end of the period, the upper part of the Excursion Module carries the astronauts back to the spacecraft. In one variant the Excursion module also uses liquid fluorine oxidizer.
The Mars departure stage burns to put the spacecraft into Trans-Terra injection.
Ordinarily the spacecraft will approach Terra at about 20 to 21 km/s. The problem is that TRW wanted to return the crew via an aerobraking Earth re-entry module, instead of using rocket thrust. Unfortunately no known re-entry vehicle could handle 20 km/s.
So the TRW mission designers had the spacecraft do a gravity-assist maneuver at Venus. This reduced the Terra approach velocity to 14 km/s, which the re-entry module could handle.
At the end of the mission when the spacecraft approaches Terra, the crew enters the Earth re-entry module and abandons the spacecraft. The empty spacecraft goes sailing off into deep space and into an eccentric solar orbit. The re-entry module does a deboost burn into Terra reentry trajectory, then jettisons the external deboost engines and propellant tanks. The module aerobrakes using its ablative heat shield. The crew is seated with their backs and the acceleration couches facing the heat shield. This ensures the deceleration pushes the crew into their couches instead of hanging from the couches eyeballs-out with the straps slicing their bodies into chunks.
In one variant re-entry module was a half-cone lifting body, 6.5 m long, 1.97 m high, and with a span of 3.84 m. In another variant, the re-entry module is a cone much like the Apollo command module. During the mission, the re-entry module doubles as the sleeping quarters.
- Earth Departure Stage: Boosts spacecraft from Terra orbit into trans-Mars trajectory. Spent stage acts as artificial gravity counterweight.
- Mars Mission Module: Crew habitat module. Nose has aerobraking heat shield to enter Mars orbit.
- Mars Excursion Module: Lands expedition on Mars and returns it to spacecraft.
- Mars Departure Stage: Boosts spacecraft from Mars orbit into trans-Terra trajectory. Spent stage acts as artificial gravity counterweight.
- Earth Reentry Module: Transports crew from abandoned spacecraft to Terra's surface, using aerobraking.
Here is a partial list of variants:
- Chemical Fuel: Oxygen-Hydrogen / Fluorine-Hydrogen
- Mars Excursion Module: Nose extend into Mission Module / Nose is below base of Mission Module
- Solar Power: Thermal boilder / Photovoltaic
- Earth Re-entry module: Conical Apollo CM style / Half-cone lifting body
- Storm Cellar: about 20 different designs
- Earth Departure Booster: Monolithic fueled in orbit / Modular assembled out of sections fueled on the ground
|Below mission module|
|Penetrates mission module|
|Crew Mod||10,068 kg|
|Shadow Shield||4,500 kg|
|Dry Tanks||14,367 kg|
|DRY MASS||97,350 kg|
|TLI Burn LH2||78,200 kg|
|LOI Burn LH2||19,990 kg|
|TEI Burn LH2||12,390 kg|
|EOC Burn LH2||26,580 kg|
|RCS fuel||2,070 kg|
|TOTAL FUEL||137,170 kg|
|WET MASS||234,520 kg|
|Actual ΔV||~ 8,620 m/s|
|Engine mass||159 kg|
|Crew Mod||9,950 kg|
|Dry tanks||1,977 kg|
|DRY MASS||17,840 kg|
|SSF to LTV burn||28 kg LH2|
139 kg LOX
|Descent burn||3,600 kg LH2|
18,000 kg LOX
|Ascent burn||2,120 kg LH2|
10,600 kg LOX
|LTV to SSF burn||9 kg LH2|
44 kg LOX
|TOTAL FUEL||34,540 kg|
|WET MASS||52,380 kg|
This is from Lunar Transportation System Final Report (1993) by the spacecraft design team of the University of Minnesota. The goal was to design infrastructure capable of cheaply transporting large payloads between LEO and the lunar surface.
The result had two main components. The Lunar Transfer Vehicle (LTV) is a nuclear powered spacecraft that ferries payloads to and from Lunar orbit. It has a habitat module for the crew. The LTV carries the Lunar Excursion Vehicle (LEV) which ferries crew of six and cargo from Lunar orbit to the Lunar surface and back.
There is an unmanned cargo version of the LEV. It has no crew module, no fuel for ascent, and carries (I calculate) about 48,000 kilograms of cargo. It will be ferried to Luna by an unmanned lunar transport vehicle controlled remotely from the Johnson Space Flight Center. The LTV will return to Terra after the cargo LEV lands.
The LEV is also used to ferry the crew from Space Station Freedom (hah! That dates it!) to the LTV at the start of the mission, and ferry the crew back at the end. This is because NASA is not going to let a spacecraft with a live nuclear reactor get anywhere near the space station. The designers initially wanted to park the radioactive LTV in between missions in a 1,200 kilometer high orbit. This was at a safe distance from the Space Station in its 400-odd km orbit, and was also high enough so if the LTV suffered a catastrophic failure no radioactive debris would reach the ground in any concentration. Unfortunately that orbit contained lots of debris from Soviet weapons testing, which would tend to cause the aforementioned catastrophic failure. The designers were forced to settle for a parking orbit that was about 10 kilometers higher than the space station's orbit, and hope for the best.
The LEV was also added as a component by the designers so it could be used as a "lifeboat" in case of emergencies. The designers had learned well the lesson taught by the Apollo 13 mission.
The initial design of the LTV was chemically powered. They switched to solid-core nuclear rocket propulsion after struggling with the inordinately large fuel masses required by chemical rockets. The chemical design also used aerobraking for the Earth Orbit Capture stage of the mission, as most chemical rocket missions do in a desperate attempt to reduce the fuel mass. The aerobraking was dropped with the switch to nuclear rockets because [a] NTR don't need no stinkin' aerobraking because they have delta-V to spare and [b] aerobraking a nuclear powered spacecraft is just begging for a radioactive disaster and a public relations nightmare.
The LTV habitat module is designed for a crew of six with enough life support for six days, plus a 48 hour contingency.
Each crew is supplied with 0.62 kg of food and 15 kg of water per day. Water must be supplied from LEO since the power is from solar cell arrays, not fuel cells who helpfully provide water as a by-product. The crew's water supply is 720 kg (including contingency), plus 280 kg of water for the science station. The total water supply is 1,000 kg.
The life support system carries 200 kg of oxygen and 650 kg of nitrogen. This is enough for 6 days plus 48 hours, and for six repressurizations of the habitat module.
The average power requirements for the habitat module is 3.1 kWe.
The LEV habitat module has far less life support. In normal operation it only has to supply the crew for a few hours, during transit to and from the Lunar surface. Most of the time the life support comes from the LTV hab module or from a pre-landed Lunar base. In an emergency the LEV may have to act as a lifeboat for up to three days. It carries 7.44 kg of food, enough for one (1) meal for each of the six crew. For the rest of the time they will just have to fast for a couple of days. There is enough breathing mix for three days plus 24 hours as contingency, and for six repressurizations (630 kg total).
Given the shadow shield screening the habitat module, it is estimated that the crew will receive from the nuclear engine a dose of 0.0548 Sieverts per mission (0.0274 Sv per transit leg). They estimate that the exposure from galactic cosmic rays is about 0.009 Sv per mission. So the total radiation exposure is 0.0638 Sv per mission (six days). This is well below NASA's guidelines of 0.25 Sv per 30 days.
But if a solar proton storm erupts, the crew is in big trouble. The LEV habitat module can be used as a partial storm cellar, because it is surrounded by tanks of liquid hydrogen and liquid oxygen. At least before it burns all the fuel by landing and ascending from Luna. In addition the shadow shield can be aimed at Sol for some more partial shielding.
The shadow shield is composed of Borated Aluminum Titanium Hydride (BATH), which was developed for the old NERVA nuclear engines. The shield is 2.54 meters in diameter, 0.186 meters thick, and weighs four metric tons.
The mass ratio of the lunar transport vehicle is difficult to figure out given the sparse information in the report. Simplistically it is about 2.4. But that does not take into account how the mass goes down after the lunar excursion vehicle expends all its fuel mass midway through the mission. It burns all its fuel landing and lifting off from Luna. Given the delta-V and specific impulse specified in the report, I calculate the effective mass ratio is more like 2.59.
|Lunar Orbit Insertion|
|Earth Orbit Capture|
Freedom to LTV
|LTV to Space|
A mission starts with the LEV docked to the space station, and the LTV at a respectable distance in its parking orbit.
For an unmanned cargo mission, there are two launches of Heavy-Lift Launch Vehicles (HLLV). One boost the cargo lander with payload, the other boosts the required propellant.For a manned mission there is only one HLLV launch, carrying the propellant. The crew travels to the space station via space shuttle or other personnel launch system. They use the LEV docked to the station to travel to the LTV. There it will dock to the LTV and be carried to Luna to proved access to the Lunar surface.
The propellant is loaded into the LTV by "wet-tank transfer", that is, the HLLV boosts into orbit propellant tanks that are already full of liquid hydrogen. These are strapped onto the LTV. The alternative, trying to pump liquid hydrogen into empty tanks on the LTV, is complicated, messy, and dangerous. The next week will be spent in vehicle check-out before it is cleared for the mission. Then and only then will the crew arrive in the LEV.
The NTR reactor is fired up and the Trans-Lunar Injection burn (TLI) starts. The burn lasts for 35 minutes and gives the ship 3,100 m/s of delta-V. It now has a three day coast before reaching Low Lunar Orbit (LLO). At some time during the coast the ship will burn for about 5 m/s of delta-V and jettison the TLI tanks. These are aimed to impact somewhere on the Lunar surface. The burn is slightly dangerous since it takes the ship off the free-return trajectory vital for an emergency mission abort (if the avionics or RCS break down or something).
After tank jettison the ship maneuvers to prepare for Lunar Orbit Insertion (LOI). The ship burns for 9.05 minutes and 1,100 m/s of delta-V and enters Low Lunar Orbit.
The ship adjusts its orbit into the proper inclination for the desired landing spot. The crew enters the LEV, which separates from the ship and does its descent burn of 17.64 minutes and 2,000 m/s of delta-V. At this point the mission elapsed time is T+72 hours.
Once on the Lunar surface, the first task of the crew is a systems check of the LEV. Because if something is wrong with your ticket back up to the orbiting ship you want the maximum amount of time to fix the blasted thing.
Assuming everything checks out the crew puts on their space suits, exit the LEV, and enters the Lunar habitat delivered by a prior unmanned mission. They then perform the scheduled 14 day mission, using life support supplies included in the Lunar habitat.
At the end of the 14 day surface mission, the Return Mission starts. The crew enters the LEV and does an ascent burn of 10.13 minutes and 1,900 m/s of delta-V. In LLO they rendezvous with the LTV. The orbital inclination is adjusted into the proper angle for Trans-Earth Injection (TEI) trajectory.
When the TEI burn starts the Return Mission elapsed time is T+5 hours. The burn is for 5.15 minutes and 1,100 m/s of delta-V. The return trip to LEO will take about two days. During this time mid-course corrections will be performed as needed. As LEO approaches, the ship will be oriented into the proper position for the Earth Orbital Insertion (EOI) burn.
The EOI burn is for 10.82 minutes and 3,000 m/s of delta-V. The ship's orbit is adjusted to bring it within parking distance of the space station (but no closer). The 20 day mission is over.
As previously mentioned the designers started out with a chemical engine. After they got tired of pounding their heads on a brick wall, they gave up and went with a solid-core nuclear engine. You can see the chemical designs in the report.
On the plus side, nuclear engines drastically reduced the required propellant mass, and eliminated the need for aerobraking (since NTR have more than enough delta-V). On the minus side the design had to be changed to protect the crew from nuclear radiation. They did try keeping the aerobrake shield as a back up deceleration method, just in case the nuclear engine malfunctioned. But they finally concluded it was not worth the mass.
The first design pass was a One-Tank configuration. A single huge tank was used to contain propellant, be the truss spine of the ship, and provide radiation shielding for the crew.
The drawback is since the tank is integral to the ship (since it is the spine), you have to use "refueling fluid transfer" to fill it. That is, at the start of each mission a fleet of tankers have to rendezvous with the ship and try to fill the tank with hoses. As previously mentioned this is complicated, messy, and dangerous. Even with a chemically powered ship. Add the fact that you are trying to get this done while in close proximity to a nuclear reactor, nope, too dangerous. Granted the reactor is not terribly radioactive when shut down, but if a tanker crashes into it you'll have dangerously radioactive fuel rods flying everywhere!
Additionally, a monolithic integral tank means you cannot do any staging, jettisoning spent tanks to increase efficency.
So the designers went with a Four-Tank layout. Two tanks stored the propellant for the Trans-Lunar Insertion (TLI) burn, and two smaller tanks were for the Trans-Earth Insertion (TEI) burn (as well as the LOI and EOI burns). This allowed the TLI tanks to be jettisoned after use, to reduce the ship mass by staging. This also allows the tanks to be "filled" by using the previously mentioned "wet-tank transfer". The integral tank was replaced by a truss, a long truss since distance is radiation shielding that cost very little mass.
This arrangement created a new problem.
The truss is only three meters square. But the tanks are so fat that they cannot be closer than one and a half meters to the truss or they bump into the other tanks. Having the tanks on 1.5 meter outriggers from the central truss is a big problem, structurally. So the designers looked into two possible strategies.
First they tried moving the fatter TLI tanks away from the engine, "upwards" so to speak. This allowed both sets of tanks to join directly to the truss and not bump into each other.
Sadly this created a new problem. The fuel lines for the TLI tanks will have to be eight meters in length or longer, which drastically reduces the efficiency of the fuel transfer to the engine.
The final solution was to use a dual truss. Most of the truss was three meters square, but the section the tanks are attached to is four meters square. This allows the tanks to not bump into each other, while keeping the fuel lines short. Everybody happy.
|Specific Power||9.6 kW/kg|
|Propulsion||Solid core NTR|
|Specific Impulse||198 s|
|Exhaust Velocity||1,942 m/s|
|Wet Mass||123,000 kg|
|Dry Mass||30,400 kg|
|Total ΔV||2,740 m/s|
|Total Propellant||92,600 kg|
|Boost Propellant||75,700 kg|
|Landing Propellant||16,900 kg|
|Boost ΔV||1,859 m/s|
|Landing ΔV||881 m/s|
|Mass Flow||155 kg/s|
|Initial Acceleration||0.25 g|
|Tank Length||8.5 m|
|Total Length||11.9 m|
|Guidance Package||0.45 tons|
and Feed Lines
|Landing System||0.68 tons|
|25% Growth Factor||2.09 tons|
The Lunar ice water truck is a robot propellant tanker design by Anthony Zuppero. Its mission is to boost 20 metric tons of valuable water from lunar polar ice mines into a 100 km Low Lunar Orbit (LLO) cheaply and repeatably. It is estimated to be capable of delivering 3,840 metric tons of water into LLO per year.
This design uses a nuclear thermal rocket with currently available materials, and using water as propellant (a nuclear-heated steam rocket or NSR) instead of liquid hydrogen). This limits it to a specific impulse below 200 seconds which is pretty weak. However, numerous authors have shown that a NSR could deliver 10 and 100 times more payload per launched hardware than a H2-O2 chemical rocket or a NTR using liquid hydrogen. This is despite the fact that the chemical and NTR have much higher specific impulses. NSR work best when  the reactor can only be low energy,  there are abundant and cheap supplies of water propellant, and  mission delta-Vs are below 6,500 m/s.
The original article describes the water extraction subsystem at the lunar pole. It is a small reactor capable of melting 112.6 metric tons of ice into water (92.6 metric tons propellant + 20 metric tons payload) in about 45 hours. This will allow the water truck to make 192 launches per year, delivering a total of 3,840 metric tons of water per year.
Since the water truck is lifting off under the 0.17 g lunar gravity, its acceleration must be higher than that or it will just vibrate on the launch pad while steam-cleaning it. The design has a starting acceleration of 0.25 g (about 1.5 times lunar gravity).
The landing gear can fold so the water truck will fit in the Space Shuttle landing bay, but under ordinary use it is fixed. The guidance package mass includes radiation shielding. In addition, the guidance package is on the water truck's nose, to get as far as possible away from the reactor. The thrust structure and feed lines support the tank and anchor the reactor. The 25% growth factor is to accommodate future design changes without having to re-design the rest of the spacecraft. The reaction control nozzles perform thrust vector control. They take up more mass than a gimbaled engine, but by the same token they are not a maintenance nightmare and additional point of failure.
The reactor supplies about 120 kilowatts to the tank in order to prevent the water from freezing. The reactor mass is 50% more than minimum. The lift-off burn is about 20 minutes durationa and consumes 0.7 kg of Uranium 235.
|Specific Power||31 W/kg|
|Propulsion||Solid core NTR|
|Specific Impulse||190 s|
|Exhaust Velocity||1,860 m/s|
|Wet Mass||299,030,000 kg|
|Water tank mass||25,000 kg|
|Sans Payload Mass||148,000 kg|
|Payload mass||50,000,000 kg|
|Dry Mass||50,148,000 kg|
|ΔV|| 802 m/s|
 1280 m/s
 752 m/s
|Mass Flow||[1,2] 903 kg/s|
 2,684 kg/s
|Thrust||[1,2] 1,680 kiloNewtons|
[1,2] 4,990 kiloNewtons
|Nozzle Power||[1,2] 4.9 gigawatts|
 1.6 gigawatts
|Engine Power||[1,2] 12.1 gigawatts|
 4.1 gigawatts
|Initial Acceleration|| 0.0006 g|
 0.0009 g
 0.005 g
The Water Ship is a robot propellant tanker design by Anthony Zuppero. Its mission is to deliver 50,000 metric tons of valuable water from the Martian moon Deimos to orbital propellant depots in Low Earth Orbit (LEO) cheaply and repeatably. It is not much more than a huge water bladder perched on a NERVA rocket engine. It might have integral water mining equipment as does the Kuck Mosquito, or it might depend upon a seperate Deimos ice mine.
Mass of water bladder is 25 metric tons (rated for no more than 0.005 g). Mass of nuclear thermal rocket plus strutural mass is 123 metric tons (struture includes computers, navigation equipment, and everything else). Mass without payload is 25 + 123 = 148 metric tons. Payload is 50,000 metric tons of water. Dry mass is 148 + 50,000 = 50,148 metric tons. Propellant mass is 248,882 metric tons. Wet mass is 50,148 + 248,882 = 299,030 metric tons.
At Deimos, only about 4.55 megawatts will be needed to melt 299,000 metric tons of ice into water (50,000 tons for payload + 249,000 tons for propellant). The engine nuclear reactor can supply that with no problem. The water must be distilled, because mud or dissolved salts will do serious damage to the engine nuclear reactor. By "serious damage" I mean things like clogging the heat-exchanger channels to cause a reactor meltdown, or impure steam eroding the reactor element cladding resulting in live radioactive Uranium 235 spraying in the exhaust plume.
Nuclear thermal rocket was designed to be a very conservative 100 megawatts per ton of engine. Engine will have a peak power of 12,142 Megawatts (for stage  and ). This works out to a modest engine temperature of 800° Celsius, and a pathetic but reliable specific impulse of 190 seconds. A NERVA could probably handle 300 megawatts per ton of engine, but the designer wanted to err on the side of caution. This will require much more water propellant, but there is no lack of water at Deimos.
This design uses a nuclear thermal rocket using water as propellant (a nuclear-heated steam rocket or NSR) instead of liquid hydrogen). This limits it to a specific impulse below 200 seconds which is pretty weak. However, numerous authors have shown that a NSR could deliver 10 and 100 times more payload per launched hardware than a H2-O2 chemical rocket or a NTR using liquid hydrogen. This is despite the fact that the chemical and NTR have much higher specific impulses. NSR work best when  the reactor can only be low energy,  there are abundant and cheap supplies of water propellant, and  mission delta-Vs are below 6,500 m/s.
It is true that electrolyzing the water into hydrogen and oxygen then burning it in a chemical rocket will get you a much better specific impulse of 450 seconds. But then you need the energy to electrolyze the water, and equipment to handle cryogenic liquids. These are just more things to go wrong.
In the table, , , and  refer to different segments of the journey from Deimos to LEO.
-  Start at Deimos. 497 m/s burn into Highly Eccentric Mars Orbit (HEMO). At apoapsis, 305 m/s burn into Low Mars Orbit (LMO)
-  At LMO periapsis, 1,280 m/s burn using the Oberth Effect to inject the water ship into Mars-Earth Hohman transfer orbit
-  270 days later at LEO periapsis, 752 m/s burn using the Oberth Effect to capture the water ship into Highly HEEO
- [x] Water ship does several aerobrakes until it reaches an orbital propellant depot in LEO
Total thrust time is about 10 hours.
Water ship's propellant has 15,137 metric tons extra as a safety margin. When it arrives, hopefully some of this will be available. It will take 322 metric tons of propellant for the empty water ship to travel from HEEO to Deimos, or 1,992 metric tons to travel from LEO to Deimos. Plus 0.139 gigawatts of engine power and 10 hours of thrust time.
Traveling from Deimos to LEO will consume about 12.7 kg of Uranium 235. Given the fact that Hohmann launch windows from Mars to Earth only occur every two years, the fuel in the engine nuclear reactor will probably last the better part of a century before it has to be replaced. The engine will be obsolete long before then.
For more details, refer to the original article.
This is from a 1963 study called Application Of Nuclear Rocket Propulsion To Manned Mars Spacecraft by Thomas Widmer. Unfortunately I cannot seem to find a copy, so most of the data comes from abstracts. It is an expansion of an earlier 1960 Lewis Research Center study.
|Lewis Nuclear Mars Mission|
|Propulsion||Solid core NTR|
|Delta V||19,800 m/s|
|Mars lander mass||40,000 kg|
|Terra lander mass||13,600 kg|
|Terra lander wingspan||6.7 m|
|Wet mass||614,000 kg|
|Mass per crew||102,000 kg|
The Lewis vehicle would have a habitat module with two levels, and 35 square meters of floor per level (3.3 meter radius). The storm cellar is a cyliiner at the centerline, and doubles as a sleeping quarters. The mass of the storm cellar depended upon the maximum allowable radiation exposure for the 420 day mission:
|Lewis storm cellar mass|
|Max 1 Sievert , no solar flares||21,400 kg|
|Max 1 Sievert , one solar flare||74,500 kg|
|Max 0.5 Sievert , no solar flares||127,000 kg|
I believe the Lewis design went with the 21,400 kg storm cellar.
The Lewis mission would use an opposition-class trajectory. Terra-Mars trajectory takes 150 days, Mars surface mission takes 40 days, Mars-Terra trajectory takes 240 days. Total mission time is 420 days. Spacecraft requires seven Saturn V launches to boost all components into orbit, each launch boosting 100,000 kg.
|Widmer Nuclear Mars Mission|
|Propulsion||Solid core NTR|
|Total Burn||3,900 sec|
|Reactor Power||2,600 MW|
|Specific Impulse||830 s|
|Exhaust Velocity||8,140 m/s|
|Engine Mass||3,200 kg|
(1 Sv mission dose)
|Payload Mass||33,960 kg|
|Dry Mass||103,000 kg|
|Wet Mass||399,000 kg|
|Delta V||16,730 m/s|
|Hydrogen per tank||20,000 kg|
|8 kW APU|
|2.5 kW APU|
|Life Support||5,220 kg|
|Hab Module||2,720 kg|
|Mars Excursion||16,450 kg|
|Storm Cellar||4,760 kg|
|Mars Excursion Module|
|Deorbit Rocket||270 kg|
(ΔV 5,330 m/s)
(10% boil off)
Parts of this section are from Application Of Nuclear Rocket Propulsion To Manned Mars Spacecraft by T. Widmer.
The Widmer vehicle was sized to have four crewmen for a 15 month mission to Mars. Just like the Bono Mars Glider, it was optimistically scheduled to depart in 1971, to take advantage of the next Hohmann launch window.
The solid core nuclear thermal rocket used a fast spectrum refractory metal core, with an inherent re-start capability and resistance to fuel cladding erosion allowing long burnning times. Long engine life and multiple restarts are extremely important factors in reducing gross vehicle weight, since they permit a low initial thrust to weight ratio (small engine), and eliminate the need for staging engines after each firing interval.
Furthermore, the smaller size of a fast metallic core provides an engine weight advantage of at least two to one over a thermal-graphite core engine of the same thrust rating. Smaller core frontal area also permits a similar reduction in shield weight. That is, the smaller the top of the nuclear reactor core, the smaller the anti-radiation shadow shield has to be, and thus the lower the shield mass.
The spacecraft components are boosted into orbit by four Saturn V boosters, one launch for the propulsion/payload module and three launches containing 4 loaded propellant tanks each. There will be a total of twelve propellant tanks. Each tank contains 20,000 kilograms of liquid hydrogen. A SNAP-9 or SNAP-50 nuclear power unit provides electricity to the cryogentic re-condensation system. The SNAP radiator is the cone shaped area just forward of the rocket engine.
Auxiliary Power Unit (APU)
The spacecraft will require about 8 kilowatts, increasing to 30 kW if the designers go with a cryogenic recondensing system in an effort to save on propellant tank insulation mass.
Fuel cells are mass hogs, they require about 16 kg of fuel and tankage per kilowatt-day (about 59,000 kg total for the mission). Solar cell arrays are massy as well, and the pesky inverse-square law dilutes the solar power available around Mars to about 43% of the energy at Terra orbit.
So the designers went with nuclear power plants. Apparently they hadn't heard about bimodal nuclear rockets because they used a second SNAP reactor perched on top of the nuclear rocket engine. In that position the center propellant tanks would shield the crew from deadly radiation (as long as the tanks were full), and the shadow shield on top of the rocket engine would prevent neutron radiation from causing the auxiliary power reactor from going all Chernobyl on them (the technical term is "neutronic decoupling").
Since the radiation from the APU reactor will kill the crew if the center propellant tank becomes too empty, the APU is turned off at that point. The APU is mostly to supply electricity to keep the hydrogen tank cool. No hydrogen, no need for electricity. The crew's modest power needs can be met by a small fuel cell or solar cell array, since at that point they will be approaching Terra.
There are two choices for power conversion equipment: Turboelectric and Thermoelectric.
Turboelectric takes hot working fluid from the APU reactor and uses it to spin a series of turbines. The turbines run conventional electrical generators, converting rotary motion into electricity (technical term is "turbo-alternator").
Advantages: lower mass than thermoelectric, can generate at power levels of 30 kW or higher. Disadvantages: turbines have a limited life, the system has so many moving parts that reliability suffers, breakdowns cause entire system to halt.
This is why multiple turbines are used, to provide some redundancy. For example an 8 kW plant might have a single SNAP-2 reactor running two operating turbines, with a third turbine sitting idle as a back up. If one turbine fails, the back up can be brought into service by activating a valve on the working fluid pipe. In the same way a SNAP-8 could energize eight turbo-alternators with one or more standby units waiting.
Thermoelectric takes the thermal gradient created between the hot and cold working fluid and converts the gradient into electricity by the Peltier-Seebeck effect (remember it does NOT convert heat into electricity, it converts the gradient into energy). The working fluid is a sodium-potassium alloy (NaK) in two loops connected by a heat exchanger full of thermoelectric elements. The primary (hot) loop starts and ends at the SNAP-8 reactor. The secondary (cold) loop starts and ends at the external heat radiator, wrapped around the end of the cryogenic hydrogen tank. The thermoelectric elements bridge the gap between the hot and cold loops, generating electricity.
Advantages: thermoelectric elements have no moving parts which increases reliability, modular construction with large numbers of thermoelectric elements means malfunctions cause a gradual degradation of power instead of a total loss. Disadvantages: can only produce up to 12 kW of electricity, has a greater mass than a turboelectric system.
|180 kg||320 kg||320 kg|
|Rad shield||640 kg||910 kg||910 kg|
|110 kg||450 kg||450 kg|
|250 kg||820 kg||998 kg|
|(25 m2)||(86 m2)||(100 m2)|
|Total||1,180 kg||2,500 kg||2,678 kg|
|The propulsion and payload module is shown in its launch configuration. The hydrogen tank and crew compartment secions are 6.7 meters in diameter. Attached to the forward end of the tank, a chemically propelled Mars excursion module will permit the landing of a two man exploration party, after the spacecraft has attained Mars orbit.|
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|One of the three tanker vehicles is shown in the launch configuration. A structural shell supports four nearly spherical tanks, each of which contains over 20,000 kilograms of liquid hydrogen. By employing auxiliary structure to reinforce the tanks during booster ascent, the weight of the tankage can be minimized. After installation on the nuclear rocket spacecraft, the light weight tanks will be exposed to only moderate acceleration (less than 1g), rather than the 7 or 8g experienced in attaining initial orbit.|
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|The separate hydrogen tanks are being attached to the propulsion module in low Earth orbit. Each tank is insulated with multi-layer radiation foils to minimized hydrogen boil-off. In addition, a cryogentic re-condensation system may be employed for those tanks which are not emptied until the later phases of the mission. This system would be powered by a SNAP-9 or SNAP-50 type nuclear electric generating system located between the main propulsion reactor and the aft end of the central tank. The radiator for the SNAP powerplant can be seen just aft of the tank. In practice, it may be necessary to move this radiator into a position well to the rear of rocket engine during coast periods, so that head load on the hydrogen tank will be minimized. An attractive possibility exists for eliminating the auxiliary power reactor by integrating a liquid metal heat exchange loop with the rocket reactor core. This approach not only reduces system weight, but also tends to minimize the problem of after-heat removal from the engine.|
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|In this view, the general arrangement of the crew quarters can bee seen. A two deck command module will contain the life support system, living accommodations, communications gear, experimental equipment, and a control center. Solar flare protection is provided by a vacuum jacketed capsule projecting downward into the main hydrogen tank. This "storm cellar" is lined with carbon shielding to augment the 2.4 meter thick annulus of liquid hydrogen which surrounds the capsule. Shielding is designed to restrict the integrated crew dose to less than 1 Sievert for the complete mission.|
Note that the proposed configuration does not provide an artificial "g" capacity. If zero "g" cannot be tolerated for the long duration of an interplanetary mission, a rotating cabin section could be factored into the design. However, this approach would result in a substantial increase in spacecraft gross weight due to structural integration problems with an artificial "g" design.
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|The orbital launch maneuver is shown here. A total of six tanks will be emptied to depart from Earth orbit and achieve the Mars transfer ellipse. In the event of an abort during the escape maneuver, the chemically propelled Mars landing craft could be used for return to Earth orbit.|
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|Staging of tanks during Earth escape propulsion is shown. Total propellant consumed up to injection for the Mars transfer is about 127 metric tons. In coast configuration two of the six tanks emptied during Earth escape will remain attached. This provides a degree of redundancy against the possibility of a meteoroid puncture in any of the loaded tanks, since propellant could be transferred into the remaining empty tanks. If no puncture occurs, the empty tanks are released immediately prior to the firing interval for Mars capture.|
Transit time to Mars is about 180 days.
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|The Mars capture maneuver produces an eccentric orbit of about 560 kilometer perigee and about 5,000 kilometers apogee; thereby minimizing propellant requirements, while still providing a close view of the planet for final evaluation of landing sites. Four of the last six external propellant tanks are emptied during capture, but only two are jettisoned. Two are retained for meteoroid puncture redundancy until just before the Mars escape firing interval.|
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|After transferring to the Mars excursion module, two of the four crew members fire braking rockets to bring the entry vehicle orbit perigee into the planetary atmosphere. The major portion of the deceleration is then accomplished by aerodynamic drag. After maneuvering to an altitude of about one kilometer, the landing craft is maneuvered into a vertical attitude for final approach. One minute of hovering capacity allows for some possible changes in landing site, and three shock absorbing struts are extended for the final touchdown. The winged entry vehicle represents one of several possible shapes, and lenticular or conical configurations might also be employed, depending upon the degree of aerodynamic maneuvering desired during entry.|
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|The Mars excursion module is shown in its landing position. In addition to the two man crew capsule, approximately 2,300 kilograms of scientific equipment and portable life support gear can be transported to the Martian surface. Equipment will include a portable meteorological station, a powerful radio for communication with Terra, and a tracked car for exploration.|
Gross weight of the excursion module prior to departure from orbit will be about 15,900 kilograms if hydrogen/oxygen propulsion is used. Stay time on the planet is restricted to about 5 days, due to limited payload and the rapid deterioration in launch window for the Earth return phase of the mission.
Note that the upper part of the Mars excursion module is a modified Gemini.
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|All equipment, except for the minimum life support capsule and 150 kilograms of soil samples, will be abandoned on the surface. The chemically propelled second stage of the landing vehicle uses the first stage structure as a launching platform for the return to Mars orbit.|
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|After rendezvous with the nuclear rocket spacecraft, the excursion module second stage is abandoned in the eccentric parking orbit.|
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|This illustration shows the Mars escape configuration of the spacecraft. During this maneuver, the last two external tanks are emptied, as is the aft compartment of the main tank. The forward end of the main tank, which surrounds the solar flare shelter, still contains hydrogen throughout the Mars-Earth transfer.|
Transit time to Terra is about 200 days
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|Upon approaching Earth, the two empty tanks are released, and the nuclear rocket engine is used to brake the vehicle into a high altitude parking orbit. The crew will then transfer to a ferry vehicle for Earth re-entry. Alternatively, it would be possible to reduce the velocity increment required of the interplanetary spacecraft by employing direct re-entry from the Mars transfer ellipse. However, this would require that an Earth re-entry vehicle be transported through the entire mission, thereby increasing the weight carried on the spacecraft. Since direct re-entry alleviates the need for a large propulsion maneuver at the terminal end of the mission, little or no propellant would be available for solar flare shielding during the return flight coast period. The flare shield weight would then have to be increased to insure crew protection in the "empty" vehicle.|
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