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.
|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.
|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.
The Class 1 Reusable Nuclear Shuttle (above) was designed to be a pre-assembled spacecraft launched into orbit by a Saturn V INT-21 vehicle. But the Class 3 RNS was designed to be assembled in orbit by modules boosted by the Space Shuttle. The major difference is that all the modules must be sized to fit into the Shuttle's cargo bay.
|Spacecraft||Payload Delivered||Recurring Cost Per Flight|
(8 per year)
|Recurring Cost Per Flight|
(2 per year)
|(lb)||(US 1970 dollars)|
|Class 1 RNS (hybrid)||128,000||$73 million|
|Class 3 RNS||108,000||$65 million|
Above table assumes minimum energy lunar missions, and a 20,000 pound payload return (i.e., the "return payload" is the crew habitat module and other items needed for the return to Terra).
The Class 1 has a lower dollar per payload pound, but the Class 2 can be lofted by the reusable Space Shuttle instead of throw-away Saturn heavy lift vehicles. Also the Class 1 requires a 10,000 biological radiation shield, while the Class 3 can get by with no shield but a lot of distance.
Actually, I lied: in some designs they use the "hybrid" engine, which has a cute little auxiliary tank perched on top. This makes the Propulsion modules composed of one big propulsion tank, an auxiliary tank, and a NERVA engine. A NERVA with the small auxiliary tank is just short enough to fit in a Space Shuttle cargo bay, this comes in handy if you are producing both Class 1 and Class 3 RNSs. You can use the same engine for both.
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|
|Class 1: 4,000 kg|
Class 3: 0 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). 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.
PROPULSION TANK MODULE
This is the main propellant tank feeding the NERVA engine. The other propellant modules keep it filled.
This is one propulsion tank module with one NERVA engine. And maybe an auxiliary tank in between if this is a hybrid propulsion module.
SPACE FRAME MODULE
The Space Frame is sort of the backbone of the ship. It holds the spacecraft together and helps transmit the engine thrust to all the components (instead of the components breaking off and falling to the wayside). It is perched on top of the propulsion module and has the propellant modules attached to all six sides. At the top is the command and control equipment module, along with the reaction control jets. The payload is attached to the top of the space frame.
The first word in the spacecraft's name is "Reusable." One of the more worrying concerns about reusing the spacecraft is the dread horror of Neutron Activation. Simply put, parts of the spacecraft that are close to the nuclear engine will gradually become radioactive. More specifically the atoms composing the engine and structural members will swollow low-energy neutron from the nuclear reactor during engine burns and thus be transmuted into radioactive isotopes. These isotopes will emit gamma-ray radiation. This wil make it read difficult to refurbish a RNS for the next trip without exposing the refurbish crew to dangerous doses of radiation.
The charts use the obsolete radiation absorbed dose unit the Rad instead of the more modern Gray. Multiply Rads by 0.01 to get Grays. Since the quality factor of gamma radiation is 1, the dose equivalent of Rem will be equal to the Rad dose (and likewise the Sievert dose will be equal to the Gray dose).
A dose of 3.5 to 3.9 Sieverts means the astronaut is "Singed". A dose over 4.0 Sieverts means the astronaut is "Cooked", which means their NASA career is over (forbidden to work at any NASA job where they might be exposed to radiation). A dose over 4.5 Sieverts is "Fried", which is LD50 (50% chance of death).
The spacecraft is assumed to have a useful life of 10 missions (due to the fission fuel elements filling up with nuclear poisons), then thrown into a radioactive disposal orbit.
Figure 4-68 shows the neutron activation radiation dose rates at various locations, starting at the final engine shut down at the end of the 10th mission. The worst is at location 3, at 0.4 Sieverts per hour. This means a Fried dose in 12 hours. Now, 417 days after final engine shut down the radiation has decayed to the point where location 3's radiation is so weak that an astronaut would have to stay there more than years to get a Fried dose.
The report is of the opinion that any astronaut refurbishment, engine swap, or other operations on the spacecraft should wait until 100 hours (102 hours) after engine shut down. This will allow the neutron activation induced radiation to die down to a less than utter suicidal level.
Table 4-24 is similar to Figure 4-68. Except the locations are different and the doses are in milliRads per hour instead of Rads per hour. 1 milliRad equals 0.001 Rad and equals 0.000 01 Sievert. "Decay Time Hours" and "Hours After Shutdown" are the same, as are the column headers (10,000 = 104)
Table 4-25 is the dosage contributed by each activated part. The values are for the case of the detector at 100 feet. And because the study authors figured things were not complicated enough, the table is in microRads per hour. 1 microRad equals 0.000 001 Rad and equals 0.000 000 01 Sievert. So for instance Table 4-24 lists the 101 shutdown dose as 13.6 milliRad and Table 4-25 lists it as 13,600 microRad.
Figure 4-69 shows the model used to calculate the doses in Figure 4-68. The model shows the location of the various components and their composition. Table 4-26 shows the percentage of the neutron activation radiation contributed by each material, figuring their in the amount of each material and its suseptibility to neutron activation.
At 10 hours after shutdown, Manganese-56 with a half-life of 2.58 hours and Copper-64, with a half-life of 12. 8 hours account for approximately 98 percent of the total dose rate. The stainless steel alloys have a maximum weight fraction of 2 percent Manganese and the aluminum alloys have a weight fraction of 0.15 to 0.9 percent Manganese and 0.4 to 6.8 percent Copper.
At 100-hr decay time, many radioactive isotopes are significant: Copper-64, 12.8 hr; Chromium-51, 27.8 days; Iron-59, 45 days; Molybdenum-99, 66 hr; Cobalt-58, 71 days; and Sodium-24, 15 hr. The latter two isotopes are products of fast-neutron reaction. The Sodium-24 isotope results from an (n, α) reaction with aluminum, while Cobalt-58 and Manganese-54 isotopes result from (n,p) reactions with Nickel-58 and Iron-54, respectively. The dose rates from the stainless steel alloys are dominated by Chromium-51 and Iron-59 isotopes, while Cu-Copper and Sodium-24 dominate the dose rates from aluminum alloys.
At 1,000 hours, the predominant isotope is Cobalt-58, followed by Manganese-54, Chromium-51 and Iron-59. At 10,000 hours, the Manganese-54 isotope becomes predominant, followed by Cobalt-58. The Zinc-65 isotope emerges as relatively significant, while Iron-59 is relegated to a role of minor importance.
As previously mentioned when the nuclear engine is executing a burn, the radiation emitted will be very unhealthy to anything else nearby. A used RNS will emit gamma radiation due to neutron activation, but during a burn it will emit both gamma and neutron radiation.
The intensity of the radiation depends upon reactor power level, distance from the reactor, and intervening masses such as the shadow shield and ship components. Figure 5-6 indicates that even at a distance of 100 nautical miles (102 NM) the dose can be as high as 10-4 REM/sec (0.000 001 Gray/sec) which is significant but not instantly lethal.
In figure 5-7 the sharp reduction in dose rate between 0° and 15° is due to the anti-radiation shadow shield. While the shadow was designed to protect the crew in the habitat module, it will also protect a second spacecraft (as per the next diagram). The 15° line is right where the Neutron curve hits the bottom of the graph. The dose is reduced by distance as per the inverse-square law (if you double the distance the strength drops to 1/4), like all radiation.
REACTION CONTROL SYSTEM
|Specific Impulse||1000 sec|
|Exhaust Vel||9,810 m/s|
|Dry Mass||4,667 kg|
|Propellant Mass||5,584 kg|
|Wet Mass||10,251 kg|
|Initial Accel||8.68 m/s|
|Specific Impulse||399 sec|
|Exhaust Vel||3,912 m/s|
|Inert Mass||83,189 kg|
|Payload Mass||10,251 kg|
|Dry Mass||93,440 kg|
|Propellant Mass||122,016 kg|
|Wet Mass||215,456 kg|
|Initial Accel||10.32 m/s|
|Specific Impulse||295 sec|
|Exhaust Vel||2,895 m/s|
|Inert Mass||61,961 kg|
|Payload Mass||225,708 kg|
|Payload||Stage II +|
|Dry Mass||287,668 kg|
|Propellant Mass||332,030 kg|
|Wet Mass||619,698 kg|
|Initial Accel||12.92 m/s|
Back in the 1950s manned space flight was the new craze. Various companies releases several plastic model kits based on hypthetical designs by such noted rocket experts as Wernher von Braun, Willy Ley, and Krafft Ehricke. These all sold quite well.
Revell Inc. was a kit manufacturer who wanted to get into the act with their own space kit. As it turns out just 26 km down the road was a new company called Systems Laboratories Corporation (SLC) which was doing actual research studies to design future spacecraft. And the founder/CEO John Barnes just happened to know the head of Public Relations of Revell. He suggested that Revell might want to take a gander at their new spaceship design. Revell founder/CEO Lew Glazer couldn't believe his own luck, and promptly accepted.
Barnes gave the job to new employee Ellwyn E. Angle, telling him to design something nice just for Revell.
Angle designed the XSL-01 Moon Rocket in 1957.
The kit sold quite well and Glazer was pleased. Actually it was one of Revell's best sellers for that year. Glazer was even more pleased when he realized that he could sell just the top part as a bargain-priced kit under the name "Moon ship." Angle also wrote an educational pamphlet included with the kit which I've reproduced below.
Glazer commissioned Angle to make a second design, for a space station. Sadly this kit did not do nearly as well, which is a pity because it is nice kit. Or so I've heard, it is so rare that I've never seen a vintage kit offered at a price I could afford. The reasons for failure were varied: it was so big it was quite a bit more expensive ($4.98 as compared to $1.98, about $44.48 in 2018 dollars), and after Sputnik went up people had soured on space. So Revell commissioned no more kits from Angle.
The XSL-01 (eXperimental Space Laboratory) was a classic "arrow" design. That is, it looked like sharp pointy thing perched on a rod. The pointy thing was the winged Moon Ship that actually performed the mission: LEO ⇒ lunar transit ⇒ lunar orbit ⇒ lunar landing ⇒ exploration ⇒ lunar liftoff ⇒ Terra transit ⇒ aerobraking ⇒ Terra landing.
The rod was a two-stage rocket whose sole purpose was just to get the Moon Ship (stage three) from the ground into low Terra Orbit. "Halfway to Anywhere" strikes again. The original design had stage I and II chemical rockets using liquid oxygen and and alcohol.
For the model kit, Angle had to shorten the stage I and II tanks to keep the kit within Revell's planned price range. The booklet says the overall length is 34 meters, I'm not sure if that with the shortened stages or not. The instruction sheet says the scale is 1/8 or 0.125 = 1 foot (1 mm = 0.08 m). Using calipers on my Moon Ship's astronauts makes this scale seem reasonable. The distance from the Moon Ship's nose to the rear of the wings is 12.0 meter on this scale. I do not have the XSL-01 model, but measuring from a couple of different images I get an overall length of 27.9 meters. Make of that what you will.
With the truncated tanks Angle was forced to use the more powerful (but insanely dangerous) oxidizer Liquid Fluorine, which has probably killed more rocket researchers than any other chemical. Or any chemist for that matter. It is sometimes used with liquid methane when you need the specific impulse of liquid-oxygen/liquid-hydrogen but cannot afford the voluminous fuel tanks required. Angle then doubled-down on danger by using hydrazine instead of methane. Hydrazine is not quite as deadly as its close cousin Unsymmetrical dimethylhydrazine (which Troy Campbell calls "explosive cancer") but it is certainly bad enough.
The Moon Ship (stage III) does not play around with feeble chemical engines, it has a full blown nuclear thermal rocket. When I look at the mass budget, I find it difficult to believe it also has a full blown radiation shadow shield thick enough to protect the crew from a lethal dose. Even if it did, the Moon Ship's wings and propellant tanks stick outside the shadow, so they will backscatter harmful radiation all over the place.
Upon return to Terra, spacecraft uses aerobraking by a series of braking ellipses over a period of two days. The drags covering the hydrogen propellant tanks do most of the work. When the velocity slows enough, the drags and the propellant tanks are jettisoned. It then does a dead-stick landing using the wings and aerodynamic control surfaces exactly like the old NASA Space Shuttle.
The column SPEED — MILES PER HOUR is a running total. The final value is the delta-V total for the first burn, the one that takes the spacecraft from the launch pad to Trans-Lunar-Insertion. This requires two chemical stages and a short burn from the nuclear engine on the Moon Ship.
The numbers in pink look like an error to me. It is impossible to have a larger propellant pounds than gross weight pounds, unless the inert mass is negative or something impossible like that. I swapped the positions of the numbers for my calculations.
21,600 miles per hour is six miles per second. This is referred to in the flight program below, at +1380 seconds. That's how I know that the final delta-V value is only for the first burn, not the entire mission.
Here is the above table in metric, with Atomic Rocket standard headers:
The ΔV Total of 9,656 m/s means Stage III (the Moon Ship) contribution was 1,945 m/s.
In addition, the Moon Ship also has to land on Luna (~2,470 m/s), lift-off from Luna (~2,222 m/s), and do a Trans-Terra Insertion (~1,076 m/s). I'll assume that it need negligable delta V to aerobrake. So more delta-V will be needed than 1,945 m/s.
This means it will need a total of about 1,945+2,470+2,222+1,076 = 7,713 m/s.
Assuming the nuclear engine has a maxed-out specific impulse of 1,000 seconds, it can manage this with a mass ratio of 2.2. This means 4,667 kg of dry mass and 5,584 kg of liquid hydrogen propellant (I tried with a more reasonable 800 second nuclear engine, but the mass ratio got ugly).
I doubt 5,584 kg of hydrogen will fit in the small external aerobrake drags since liquid hydrogen is annoyingly non-dense. The entire rear of the Moon Ship is probably full of LH2 as well.
|-10 sec||0||0||0||Stage I ignition|
|-1 sec||0||0||0||Stage I full thrust|
|0||0||0||0||Blast Off with Stage I|
|+85 sec||48 km||Start of gravity turn|
|+200 sec||Stage I throttle-down|
|+205 sec||177 km||+2,222 m/s||2,222 m/s||Stage I cut-off and jettison|
|+206 sec||Stage II ignition|
Stage I recovery system activated
|+250 sec||Stage III nuclear core warm up|
|+345 sec||Stage II throttle-down|
|+350 sec||676 km||+3,268 m/s||5,490 m/s||Stage II cut-off and jettison|
|+351 sec||Stage II recovery system activated|
|+355 sec||Stage III nuclear engine ignition|
|+360 sec||Radar and mercury boiler|
housing cones open
|2,092 km||+1,945 m/s||9,656 m/s||Stage III nuclear engine cut-off|
Entering ascent coast phase
Rest of Mission
- Lift-off / Trans-Lunar Insertion (TLI)
- Ascent Coast Phase
- Lunar Landing
- Lift-off / Trans-Earth Insertion (TEI)
- Descent Coast Phase
- Earth Orbit Insertion (EOI)
- Braking Ellipses and Landing
Braking ellipses is aerobraking on the installment plan. Each aerobraking pass slows you down a little more. In two days you will be slow enough to actually land at the airfield.
|+1,945 m/s||Stage III nuclear engine cut-off|
Entering (2) Ascent Coast Phase
|Within 58,050 of Luna|
(Lunar Hill Sphere)
|+4d, 16hr||Within 39,000 km of Luna|
Nuclear core warm up
|+5d||Altitude 3,058 km|
(4,795 from Lunar Center)
|+5d, 3hr||2,470 m/s||(3) Lunar Landing in crater Plato|
Nuclear engine cut-off
Transient lunar phenomena
|+7d, 12hr||3,298 m/s||Nuclear core warm up|
(4) Lift-off / Trans-Earth Insertion
Nuclear engine cut-off
Entering (5) Descent Coast Phase
|+8d||Leaves Luna Hill Sphere|
(58,050 km from Luna)
Orbital position checked with ground bases
|+12d||(6) Ship aimed at edge of atmosphere|
|+12d, 3hr||Aerobraking starts|
(7) Ship undergoes series of braking ellipses
|+14d||Ship slow enough to enter atmosphere|
Conical tanks jettisoned
|+14d, 3hr||Atmospheric entry|
|+14d, 4hr||Landing at base|
|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)
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 general rule 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.
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.
This is from Space rescue operations. Volume 2: technical discussion and Space rescue operations. Volume 3: Appendices (1971)
This is a crewed spacecraft designed for a rescue mission, to save astronauts in a disabled spacecraft.
Warning: don't be confused. In the documents are references to an Earth Orbit Shuttle (EOS) and a Space Shuttle (SS). The EOS is what we would call a Space Shuttle, and the Space Shuttle is what we would call a Reusable Nuclear Shuttle.
In the documents, focus on the rescue vehicle called the EOS/MCCM, and ignore any references to the "Space Shuttle." I became mightily perplexed while reading the documents before I figured this out. EOS is "Earth Orbit Shuttle", a reusable heavy lift vehicle. CCM is "Crew/Cargo Module", a standard module sized to fit inside the EOS. They took the CCM design and modified it into an orbit-to-orbit rescue vehicle, a "Modified Crew/Cargo Module" or MCCM.
The Space Rescue Vehicle (SRV) is a standard EOS crew/cargo module (which in our time-line was never created) modified into a space vehicle. It is called the Modified Crew/Cargo Module or MCCM. Rescue specific features include:
- Docking fixtures
- Air lock
- Special rescue equipment
- Rescue trained crew
For low delta-V missions (60 m/s) it relies upon its RCS for propulsion, if more delta-V is needed a large propulsive module (PM) can be attached (LOX/LH2 fuel). It is not capable of reentry, it has to return to an orbital safe haven (space station or reentry vehicle). It can be based in orbit, or based on Terra and boosted into orbit by an EOS.
An MCCM boosted by an EOS has no propulsive module, if one is needed a second flight is need to boost it into orbit to be mated to the MCCM. The modules will probably be loosely based in the propulsion modues for the Boeing Space Tug. In the table below, different sizes of propulsive modules are shown with their different delta-V capabilities.
|ΔV||61 m/s||305 m/s||4,330 m/s||5,490 m/s|
|Life Support||4 days||4 days||4 days||14 days|
|PM dry mass||0 kg||358 kg||5,200 kg||7,400 kg|
|Propellant Mass||187 kg||1,000 kg||31,500 kg||51,700 kg|
|TOTAL||13,600 kg||15,000 kg||50,300 kg||72,600 kg|
The Space Rescue Vehicle must provide:
- A habitable haven for the rescued crew
- Medical aid (facilities and service) for ill or injured personnel
- Life support for extending crew survival
- Communication with the disabled crew during the rescue operation
- Emergency power during the rescue operation
- Transportation from the scene of the emergency to a final haven of safety
The Space Rescue Vehicle coming to the aid of a distressed vehicle (DV) may need the following capabilities:
- Collision avoidance with debris generated by the DV
- Protection from DV radiation sources
- Ability to dock with a disabled vehicle
- Ability to arrest the motion of a tumbling vehicle
- Ability to retrieve personnel from EVA and from a DV where docking is not possible
- Ill or Injured Crew (physical, chemical, disease, mental)
- Metabolic Deprivation
- Stranded or Entrapped Crew
- during EVA
- in vehicle
- Inability to Communicate
- Out-of-Control Spacecraft
- tumbling in safe orbit
- in decaying orbit
- on unsafe trajectory
- Debris in Vicinity
- Radiation in Vicinity
- Non-Habitable Spacecraft Environment
- lack of environmental control (pressure, temperature, humidity extremes)
- contamination (experiments, animals, insects, bacterial)
- Radiation (internal source)
- Abandonment (crew in EVA after bail-out in lifeboat)
- Inability to Reenter Earth's Atmosphere
Factors to consider in determining the rescue vehicles requirements:
- Hazards to the SRV (such as debris or radiation) caused by the distressed vehicle
- Problems of personnel and equipment transfer to and from the distressed vehicle under docked and undocked conditions; specialized equipment needs include:
- Means for establishing communication with a mute spacecraft after rendezvous
- Procedures for gaining emergency access to the interior of a disabled vehicle
- Equipment for assessing and controlling damage to the disabled vehicle
- Medical aid for the rescued crew
- Portable equipment and supplies to provide extended survival on an emergency basis for the crew of the disabled vehicle
- ΔV needs of the SRV for rendezvous and an external inspection of the disabled vehicle
|Communication and Survey Equipment||318|
|Soft Docking Fixture||113|
|Attachable Docking Fixture||363|
|Sampling and Analysis Kit||23|
|Damage Control Equipment||68|
|Extended Survival Kit||227|
|Miscellaneous and Spares||91|
As mentioned before, the Space Rescue Vehicle is a EOS crew/cargo module modifed into a spacecraft.
The foreward compartment is probably loosely based on the crew module for the Boeing Space Tug.
The center compartment is retrofitted with a sizable reaction control system (RCS). This can be the entirety of the spacecraft's propulsion system for missions with delta-V requirements under 60 meters per second. Otherwise a larger propulsion stage is mated to the "aft" end.
The aft cargo compartment is refitted to accomodate crew and passengers from the distressed vehicle, including incapacitated members transported by personnel carriers. The cargo section is also outfitted to allow medical aid to be provided, allowing the SRV to also act as an ambulance.
Since the SRV is based on an EOS crew/cargo module, it can be designed to fit into an EOS (NASA space shuttle) cargo bay. This will allow it to be boosted into orbit and recovered back to Terra's surface by an EOS. This allows the SRV to use off-the-shelf technology instead of the headache of designing some new technology from scratch.
The SRV has an estimated reaction time of one to two days, between the declaration of the emergency and the launch of the SRV. Estimated cost is $250 million US in research and development, and $70 million US per unit, in 1971 dollars (about $1.55 billion and $434 million US in 2019 dollars). Estimated service life of the SRV is 16 rescue missions.
It is possible to make an uncrewed version of the SRV, but of course the rescue will need more self-help on the part of the crew of the distressed vehicle. The SRV is required in rescues when the crew of the distressed vehicle are incapacitated or otherwise incapable of utilizing self-help.
|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|
|Specific Impulse||850 sec|
|Exhaust Velocity||8,340 m/s|
|Engine Mass||17,000 kg|
|Reactor Power||5.1 GW|
|STAGE I Leave Terra – Nuclear|
|STAGE I TOTAL||441,485 kg|
|STAGE I ΔV||3,810 m/s|
|STAGE II Outbound midcourse|
|Structure + engine||2,575 kg|
|Propellant (Impulse)||14,526 kg|
|Propellant (Attitude)||4,594 kg|
|STAGE II TOTAL||21,694 kg|
|STAGE III Arrive Mars – Nuclear|
|Propellant boiloff||5,302 kg|
|STAGE III TOTAL||196,736 kg|
|STAGE III ΔV||3,215 m/s|
|STAGE IV Leave Mars – Nuclear|
|Propellant boiloff||5,510 kg|
|STAGE IV TOTAL||146,191 kg|
|STAGE IV ΔV||5,327 m/s|
|STAGE V Inbound midcourse – Storable|
|Propellant (Impulse)||2,122 kg|
|Propellant (Attitude Stopover)||422 kg|
|Propellant (Attitude Inbound Leg)||674 kg|
|STAGE V TOTAL||3,784 kg|
|STAGE VI Earth Retro – LH2/LOX|
|Structure, Engine, Insulation||3,026 kg|
|Propellant boiloff||1,028 kg|
|STAGE VI TOTAL||13,627 kg|
|STAGE VI ΔV||17,967 m/s|
|Mars Lander||35,607 kg|
|Storm Cellar||10,405 kg|
|Habitat Module||31,177 kg|
|Life Support Consumables||10,319 kg|
|Reentry Capsule||6,271 kg|
|TOTAL PAYLOAD||93,780 kg|
|TOTAL VEHICLE||917,296 kg|
This study was performed under NASA contrat NAS8-5371, and was an incredible tour de force of comprehensive parametric mission analysis. The results were published in nine volumes, I've only found two. They used two trajectory optimization computer programs to generate over 20,000 different mission simulations, with an optimum trajectory and vehicle determined for each simulation. The results were:
- Optimum thrust range for the advanced nuclear engine
- Design characteristics of a compromise advanced nuclear engine
- Sensitivity of vehicle design to:
- variations in mission
- variations in engine
- variations in vehicle parameters
- variations in operating modes
As you can see this is yet another one of those insane designs with nuclear stages, jettisoning violently radioactive nuclear engines willy-nilly into eccentric Solar orbits as a deadly surprise for space explorers for centuries to come. Or maybe not. There is still a lot of valuable uranium-235 left in those engines, all it needs is some fuel rod reprocessing. Future scavengers will try to track the stages down and salvage them.
|Cryogenic Chemical (LOX/LH2)||440 sec|
|Storable Chemical||330 sec|
|Attitude Control||1 percent each leg|
|Optimum Cryogenic Insulation/Boiloff|
|MARS STOPOVER MISSION CRITERIA|
|Earth Recovered Payload||10,000 lb|
|Habitat Module (8 Crew)||68,734 lb plus storm cellar|
|Mars Lander (MEM)||80,000 lb|
|Weight Recovered from MEM||1,500 lb|
|Life Support Expendables||50 lb/day|
|Stopover Time||20 days|
|Midcourse Correction||100 m/sec each leg storable propellant|
|FLYBY MISSION CRITERIA|
|Earth Landed Payload||8,500 lb|
|Mission Module (3 Crew)||65,000 lb including storm cellar|
|Planet Probe||10,000 lb|
|Life Support Expendables||40 lb/day|
|Planet Passage Altitude||Mars 1,000 km (Rd = 1.3)|
Venus - 1,000 km (Rd = 1.16)
|Midcourse Correction||200 m/sec outbound leg|
300 m/sec inbound leg
|LUNAR TRANSFER MISSION CRITERIA|
|Payload in 100 nmi Lunar Orbit||100,000 to 400,000 lb|
|Midcourse Correction||30 m/sec storable propellant|
|Transfer Time||70 hr|
The data in the table to the right is a representative vehicle for a Mars mission. For a Lunar Cargo mission delivering payload into a 100 nm circular lunar orbit, the vehicle masses are in the table below. The delta V required is about 30 m/sec.
|Payload Mass||Vehicle Mass|
|90,700 kg||226,800 kg|
|136,100 kg||340,200 kg|
|181,500 kg||430,900 kg|
|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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