Inspired By Reality

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

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

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

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

A. C. Clark

A. C. Clark
SNRE-class Engine
Thrust73,000 N
900 s
367 MWt
0.89 m
0.6 g U
per cm3
U-235 wt
2,860 K
59.6 kg
Crew Size5
Length89.4 m
Engine Arrayx3 engines
Engine Mass100 t
Shield Mass
6 t
Tank Mass
95.8 t
Star Truss
& x4 drop tanks
197.5 t
Payload86.7 t
Inital Mass
480 t
62.4 t
71.6 t
Drop Tank
141.4 t
(35.4 @)
Hab Modules42.2 t
5 crew + suits1.0 t
Logistics Hub7.2 t
and braces
5.5 t
Consumables4.4 t
8.1 t
Orion MPCV13.5 t
RCS and
4.8 t

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

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

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

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

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

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

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

Habitat modules

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

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

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

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

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

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


Engine Mass
107,000 kg
Engine Mass
(mod oil)
91,000 kg
Reactor Power2.5 GW
Thrust4,651 N
Thrust Power730 MW
32,000 sec
313,900 m/s
Mass Flow
3.12×10-5 kg/s
Mass Flow
0.0179 kg/s
Mass Flow
0.018 kg/s
RCS925 kg
Propulsion268,961 kg
Structure5,899 kg
280,816 kg
Power6,200 kg
Avionics3,118 kg
INERT MASS565,865 kg
Payload170,000 kg
DRY MASS735,919 kg
Propellant345,599 kg
WET MASS1,081,518 kg
Mass Ratio1.47
ΔV120,900 m/s

Robert Werka has apparently figured out a new configuration for his fission-fragment rocket engine (FFRE).

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

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

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

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

Robert Werka and Thomas Percy took the standard Human Exploration of Mars Design Reference Architecture 5.0 and designed it using an AFFRE in Opening the Solar System: An Advanced Nuclear Spacecraft for Human Exploration Report, Slides.


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

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

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

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

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


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

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

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


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

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

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


The AFFRE has such a spectactular specific impulse that most designs have outrageous amounts of delta-V. Other engines are so weak that they must need to resort to staging (even entire NERVA engines jettisoned) and even then the remaining part of the spacecraft is about the size of the Apollo command module. Everything else is thrown away. The AFFRE ship on the other hand returns to Terra basically intact, so you can reuse the entire thing for multiple missions.

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


This is from Affordable In-Space Transportation (1996)

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

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

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

They found this is very hard to do.

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

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

Ground Rules:

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


EngineSolid core NTR
Thrust67,000 N
900 s
Mass Flow
7.6 kg/s
Engine Mass2,450 kg

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

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

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

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

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

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

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

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

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

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

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

NTR Transport Mass Budget
Passive thermal control202279
Propulsion subsystem7979
Nuclear Rocket Engine2,4542,454
Customer Payload1,3614,536
Apogee Kick motor91302
AKM Propellant1,2054,016
TOTAL DRY MASS7,56414,758
Stage Fuel4,8147,830
TOTAL WET MASS12,37722,588
Stage Delivery ASE weigh1,8573,388
Mass Ratio1.6361.531
Exhaust Velocity8,829 m/s8,829 m/s
delta V4,348 m/s3,758 m/s
Refuel/Resupply Package
Passive thermal control202279
Propulsion subsystem7979
Apogee Kick motor91302
AKM Propellant1,2054,016
Resupply Fuel Weight4,8147,830
Replacement Parts Weight78109
Customer Payload weight1,3614,536
TOTAL DRY MASS9,29119,421
Stage Delivery ASE weight1,3942,913

Antares Dawn Battlecruiser

Battlecruiser Discovery
EnginePhoton drive
(with gears)
ΔV10,500,000 m/s
(10,500 km/s)
2.36×1014 W
(236 terawatts)
(471 terawatts)
Fusion fuel
burn rate
0.73 kg/sec
Engine High Gear
9.81 m/s2
(1 g)
Thrust1,570,000 N
3×108 m/s
30,600,000 sec
Engine Low Gear
39 m/s2
(4 g)
Thrust6,240,000 N
75,500,000 m/s
7,690,000 sec
Length110 m
Body Dia14.5 m
22 m
4 m
6,950 m3
1.0 g: 6.4 RPM
0.5 g: 4.5 RPM
0.1 g: 2.0 RPM
Ship aprox
32,000 m3
Ship aprox
5 kg/m3
Ship aprox
wet mass
160,000 kg
x4 armed scouts
Weaponsantimatter proj
particle beam
FTL energy10% fuel/jump

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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


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

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

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

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

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

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

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

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

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

“Close outer doors. Stand by to repressurize.”

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

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

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

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

From ANTARES DAWN by Michael McCollum (1986)

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

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

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

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

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

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

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

Moliere had arrived.

From ANTARES PASSAGE by Michael McCollum (1998)

Asteroid Mining Crew Transport

This section has been moved here

Asteroid Survey Vehicle

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

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

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

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

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

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

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

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

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

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

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

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

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

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


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

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

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

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


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

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

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


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

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

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

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


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

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

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

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

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

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

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


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

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

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


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

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

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

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

Atomic V-2 Rocket

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

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

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

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

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

Aurora CDF

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

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

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

The spacecraft is composed of four parts:

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


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

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

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

Trans-Mars Injection Stages

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

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

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

Mars Orbit Insertion Stage

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

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

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

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

Trans Earth Injection Stage

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

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


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

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

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

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

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

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

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

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

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

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


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

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

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


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

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

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


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

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


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

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

Aurora CDF Project Troy

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

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

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

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

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

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

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

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

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

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

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

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

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

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

Austin Mars Mission

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

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

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

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

In the diagram at left:

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

The three habitable components are:

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

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

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

The six propulsive stages are:

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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


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

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

Basic Solid Core NTR


RocketCat sez

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

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

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

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

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

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

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

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

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

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

Rescue Ship

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

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


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

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

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

Shadow Shield

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

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

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

Plug Nozzle

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

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

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

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

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

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

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

Thrust Vectoring

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

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

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

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

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

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

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

Cascade Vanes

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

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

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

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

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

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

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

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

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

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

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

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

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

Bimodal NTR

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

This is from a NASA study TM-1998-208834-REV1. The idea was to take NASA's Mars Design Reference Mission (DRM) and update it. Specifically a throwaway stage with a nuclear thermal rocket (NTR) was to be replaced with a reusable stage using an NTR with the bimodal option.

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

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

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

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

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

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

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

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

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

Bimodal Hybrid NTR NEP

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

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

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

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

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

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

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

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

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

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

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

Bimodal Hybrid NTR NEP 2

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

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

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

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

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

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


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

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

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

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

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

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

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

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

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

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

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

Benton Spaceship Discovery

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

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

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

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

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

Four types of landers were designed:

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

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

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

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

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

Crew Module (CM)

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

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

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

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

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

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

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

Docking Module (DM)

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

Service Module (SM)

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

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

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

Engineering Module (EM)

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

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

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

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

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

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

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

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

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

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

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

Strap-On Boosters

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

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

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

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

Crew Reentry Module (RM) for Terra Return

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

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

Design Reference Mission 2 (DRM 2) – Mars Exploration

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

Design Reference Mission 5 (DRM 5) – Callisto Exploration

Design Reference Mission 6 (DRM 6) – Ganymede Exploration

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

Comparison of Design Reference Missions

Boeing IMIS

RocketCat sez

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

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

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

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

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

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

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


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

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

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


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

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

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

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

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

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

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

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

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

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

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

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

Propulsion Module

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

These look suspiciously like the NASA reusable nuclear shuttle.

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

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

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

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

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

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

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

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

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

Mission Module


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

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


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


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


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

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

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


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

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

Boeing STCAEM Mars NTR

(Nuclear Piloted version)
Thrust330,000 N
Engine Mass3,402 kg
925 sec
9,070 m/s
Habitat Module47,000 kg
(Mars Lander)
5,700 kg
Dry Mass260,360 kg
Propellant Mass554,520
Wet Mass814,880 kg
Mass Ratio3.1
ΔV10,350 m/s

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

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

Boeing STCAEM Cryo/Aerobrake

(Chemical Piloted version)
Height50 m
Span30 m
Mars Surface
25,000 kg
Dry Mass301,000 kg
Propellant500,000 kg
Wet Mass801,000 kg
Mass Ratio1.6
Specific Impulse475 sec
Exhaust Velocity4,660 m/s
ΔV2,190 m/s
Outbound Time350 days
Mars Stay Time30 days
Return Time200 days
Total Mission580 days

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

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

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

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

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

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

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

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

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

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


Aerobraking Shield

Transfer Vehicle

Terra Reentry Vehicle

Mars Lander

Spin Gravity Configuration

Aerobrake Shield Booster Vehicle

Bono Mars Glider

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

Blue Max Studio Liberty Bell

Liberty Bell
ParamLaunchLunar Run
6.18 MN2.2 MN
10.5 (4.2)*1.56
Accel10 m/s24.58 m/s2
1,132.4 kg/s1,000 kg/s
5,457 m/s2,200 m/s
556.2 s452.3 s
770 s105 s
≈2 hrs5.56 days
Δv7,842 m/s4,947 m/s

The Liberty Bell is a tramp freighter created by Ray McVay for his Black Desert universe

The Liberty Bell proper is a command module with a dry mass of 50 tons, and 50 tons of propellant. It has a power plant, life support, and thrusters. It can carry a crew of five plus up to 20 passengers from the surface into LEO.

On the nose is an airlock with an androgynous docking port and a maneuvering unit.

On the tail there are four couplers, each of which can hold one cargo container. The containers are cylinders 9.5 meters long and 5 meters in diameter. They are rated to carry a maximum of 62.5 tons of cargo each.

There are four remote manipulator arms used to handle cargo containers. The arms are not permanently attached, they can move like a giant inchworm over the spacecraft's surface just like the Canadarm 2 on the International Space Station.

The Liberty Bell is boosted into orbit with an L-Drive assembly. This is a laser launch system. At the spaceport, the launch pad has a huge stationary laser built into it. The L-Drive assembly is attached to the bottom of the Liberty Bell. The L-Drive is an air-breather, it scoops up atmosphere and sprays it into the mirrored dish-with-a-spike. The laser beam from the launch pad heats the air, creating the thrust to boost the spacecraft into orbit. The laser beam tracks the L-Drive as it climbs into the sky. When the L-Drive reaches an altitude where the air is too thin, it switches to its internal propellant tanks.

Typically the L-Drives are owned and maintained by the spaceport, they cost $1,250,000 Black Desert dollars. The spaceport will rent an L-Drive, laser boost time, plus fees and taxes to the captain of the Liberty Bell. This will cost the captain $100,000 total to boost the Liberty Bell into LEO.

Upon reaching LEO, a Liberty Bell generally makes a rendezvous with an orbital transport nexus, unloads its four cargo containers (250 tons of cargo total) and 20 passengers, loads new cargo and new passengers to be delivered to Terra's surface, pays the spaceport for laser landing services (including fresh propellant for the L-Drive), and rides the laser beam back down to the spaceport.

However, our Liberty Bell is heading to Luna.

The Liberty Bell jettisons the L-Drive, delivering the rental vehicle back into the hands of spaceport personnel (the orbital representatives). The captain knows that when they make the return trip, the spaceport will be more than happy to reserve them an L-Drive for the trip down.

On this trip, instead of carrying four cargo containers, the Liberty Bell only has two containers (125 tons), a translunar rocket engine (20 tons, thrust equivalent to a SSME), and a small cobbled together weapons package (105 tons). The total payload tonnage is 250 tons, same as four cargo containers.

The weapons package contains two Kinetic Kill Vehicles (KKV) at 40 tons each, two Caltrop space mines at ten tons each, and a laser turret with power supply at five tons.

The Liberty Bell then moves into a higher orbit, to make a rendezvous with a transfer space station. In the Black Desert universe, the orbits are patrolled by the astromilitaries of various nations, all looking for trouble and whatever they can get away with. This is the main reason for the Liberty Bell's weapons package.

At the transfer station, the Liberty Bell outfits itself for the Lunar trip. It leases four propellant tanks to feed the translunar rocket engine. It also leases or purchases a cupola.

Using the remote manipulator arms, the translunar rocket engine and the airlock/docking ring swap positions. The rocket engine is mounted on the nose and the four propellant tanks are attached. The docking ring is mounted next to the other cargo, and a cupola installed on top. For the rest of the trip, the cupola will serve as the Liberty Bell's cockpit.

As it turns out, one of the captain's business partners had three cargo containers waiting at the transfer station to be delivered to Luna. The remote manipulator arms install these as well.

The Liberty Bell is ready for the trip to Luna. The command module now faces opposite the direction of thrust it had at launch, with the cupola and the weapons package aimed at the new forwards that used to be backwards. It is carrying three hundred tons of cargo.

It has enough life support and consumables to haul five crew and twenty passengers on the five and a half day trip to Luna or one of the La Grange stations.

Discovery II

This section has been moved here

Douglas Mars

This is from Proceeding of the Symposium on Manned Planetary Missions 1963/1964 Status, part 10 "Manned Mars Exploration in the Unfavorable (1975-1985) Time Period". It is a report on a Mars mission study by the Douglas Aircraft Company.

The section about the spacecraft is interesting because they examine about 15 different options, and score them according to a variety of criteria. They went with option 5.

The missions was to be 460 days duration wih a 20-day Mars capture-orbit stay time. The unsurprising recommendations were to restrict crew selection to 20-percentile men (sexist!), have the crew cabins as close as possible to the drive-the-astronauts-to-psychotic-break mimimum size limit (31.15 m3 per crewperson), combine meteoroid and insulation with the load-carrying structure (oh, like any spacecraft design doesn't do that?), a crew of six, use fiberglas tanks, and gas core nuclear thermal rockets would be real nice if they could be man-rated (in your dreams...).

The spacecraft would have a wet mass of 979,000 kilograms, and a dry mass of 278,000 kilograms. It would have four stages, not counting the ROMBUS reusable chemical booster that lofts it into LEO. A separate ROMBUS flight lofts the propellant. After each burn the current stage is discarded along with the still-hot nuclear engines. This means the spacecraft does not have to carry along extra propellant to cool down the engines.

The report is a little vague on performance. If this was a single-stage rocket it would have a delta-V of about 11,000 m/s. Since it is a staged rocket it presumably has more than that.

The habitat module with consumables for the crew of six is 35,320 kg, which is the mass of the payload package less the mass of the Mars Excursion Module and the Earth Entry Module. The payload is packed around the fourth stage. The artificial gravity centrifuge is an enclosed ring containing two cable-driven carts riding on the inner surface of the cylindrical rails.

Stages one and two have 250K Phoebus nuclear engines, stages three and four have 30K metallic core nuclear engines (as opposed to graphite core). Each Phoebus engine has a thrust of 1.11×106N, each metallic engine has a thrust of 1.33×105N. Both have a specific impulse of 850 seconds (8,340 m/s exhaust velocity). It would be nice to use the metallic engines on the lower stages, but you'd need clusters of eight, and nuclear decoupling is a big challenge (neutrons from adjacent engines make the nuclear chain reaction in a given engine go out of control).

1250K Phoebus2
2250K Phoebus1
330K Metallic2
430K Metallic1

After leaving Mars, when approaching Terra, the fourth stage nuclear engine will slow the vehicle down to 12.2 km/s relative to Terra. The remaining velocity will be eliminated by aerobraking with the astronauts inside the Earth Entry Module. The rest of stage four will go sailing off into the wild black yonder.

Aeronutronic EMPIRE

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

The ship is now reconfigured into orbit mode.

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

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

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

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

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

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

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


PowerNuclear Reactor
Mission Duration1.34 yrs
Propellant Mass38,600 kg
Payload Mass39,500 kg
Radiator Mass25,800 kg
Reactor Mass18,600 kg
Ion Drive Mass4,500 kg
Wet Mass127,000 kg
Dry Mass88,400 kg
Mass Ratio1.44
Thrust125 N
Specific Impulse14,000 s
Exhaust Velocity137,000 m/s
ΔV49,600 m/s
Total Power10 MW
Power for Drive9.5 MW
Departure Orbit560 km
Venus Orbit104,000 km

Details on this design are sparse. Apparently it is a 1960 Convair/General Dynamics design for a Venus orbital mission. However my main source of information was from the 1966 Young People's Science Encyclopedia vol: Sp-Su. If anybody has more info please get in touch with me.

The spacecraft is a nuclear-electric ion drive ship with plenty of heat radiators and habitat modules on arms for spin gravity.

Having said that:

  • The radiators spray heat on each other, which is counter-productive. Angle between adjacent fins should be no less than 90 degrees or the efficiency goes down. This design has them spaced at 45°, for an miserable efficiency of 38%

  • The spin gravity centrifuges are parallel to the thrust axis, requiring two of them counter-rotating so it doesn't precess all over the place. Almost all other designs I've seen have one centrifuge placed normal to the thrust axis.

  • The radiators are properly trimmed at an angle so they stay inside the protective shadow cast by the anti-radiation shield. The ship's spine appears capable of telescoping out to increase the habitat module's distance from the radioactive reactor.
    Alas it appears that the hab modules are sticking out of the safe shadow into the deadly shine from the reactor. Either that or the radiators are trimmed back too far, which is a waste of radiator area.

Enzmann Starship

This section has been moved into the Slower Than Light page.

Exacting Class Starfighter

This section has been moved here

First Men to the Moon

This design is from a book called First Men to the Moon (1958) written by a certain Wernher von Braun, aka "The Father of Rocket Science" and the first director of NASA. The book came out shortly after the Sputnik Crisis.

Gasdynamic Mirror

This section has been moved here

GCNR Spacecraft

RocketCat sez

You want an atomic rocket? I'll give you an atomic rocket!

Yeah, yeah, this ain't an over-the-top torchship like an Orion Drive ship much less Zubrin's outrageous Nuclear Salt Water Drive. But it is a good working-man's atomic rocket that has the horsepower to Get The Job Done. Orion drives are for battleships, this one is a space trucker hauling cargo.

Bloated chemical drives can barely do a Mars mission in two years, this little atomic number can smoke the mission in 80 days flat! I know that saying the exhaust is radioactive is putting it mildly, but nobody is near enough to it to be harmed (well, except for the poor working-class slobs who are the ship's crew).

Bottom line:

  • It is undisputably an Atomic Rocket

  • It has both high thrust and high specific impulse, approaching torchship levels

  • The design does a clever end-run around the "melting reactor" problem with a solution both elegant and brute force

  • 80 day round-trip to Mars, man! How cool is that?

GCNR Spacecraft
Specific Impulse2,500 to 6,500 s
Exhaust Velocity24,500 to
63,800 m/s
Mass Flow0.8 to 6.7 kg/s
Thrust20,000 to
430,000 N
Fixed Thrust224,000 N
Thrust Power0.25 to 13.7 GW
Initial Accel0.01 to 0.05g
GCNR Spacecraft
Mars Courier
80 days
Wet Mass950,000 kg
Dry Mass290,000 kg
Mass Ratio3.28
Thrust150,000 N
Initial Accel0.016 g
Specific Impulse5,500 s
Exhaust Velocity53,955 m/s
ΔV64,100 m/s
H / 235U Ratio200:1
235U Fuel3,300 kg
660,000 kg

Data from Gas Core Rocket Reactors - A New Look.

This little hot-rod can do a round-trip mission to Mars in 80 days flat! That's only 2.7 months. Using Hohmann trajectories a round-trip Mars mission will take 32.3 months (2.7 years) when you take into account the wait for the Mars-Terra launch window to open.

The report starts off with the common complaint that most rocket propulsion is either high-thrust + low-specific-impulse or vice versa. The problem being that rocket designers want a high-thrust + high-specific-impulse engine. In other words they want a torchship.

The closest thing they can find that is actually feasible is a Gas-Core Nuclear Thermal Rocket. Open-cycle of course, closed-cycle has only half the exhaust velocity. So what if it spews still-fissioning uranium in an exhaust plume of glowing radioactive death?

The report examines the GCNTR's performance to see if it is a torch drive. It comes pretty close, actually.

The higher the specific impulse / exhaust velocity, the more waste heat the engine is going to deal with. They figure that a GCNTR can control waste heat with standard garden-variety regenerative cooling like any chemical rocket, but only up to a maximum of 3,000 seconds of specific impulse. Past that you are forced to install a dedicated heat radiator to prevent the engine from vaporizing. Otherwise the engine vaporizes, your spacecraft has no engine, and perhaps centuries from now your ship will come close enough so that space archaeologists can recover your mummified remains.

As everybody knows, thermal rockets use a heat source to heat the propellant (usually hydrogen) so that its frantic jetting through the exhaust nozzle creates thrust. Solid-core nuclear thermal rockets (NTR) use solid nuclear reactors. They are limited to a specific impulse (Isp) of about 825 seconds, since that corresponds to a propellant temperature of about 2,500 K. Any higher specific impulse raises the temperature high enough that the reactor starts to melt. And nobody likes an impromptu impression of the China Syndrome. If you want an Isp of 5,000 seconds you are talking about a propellant temperature of 22,000 K!

Also as everyone knows the gas-core NTR concept is the result of clever engineers thinking outside of the box and asking the question what if the reactor was already vaporized?

Instead of solid nuclear fuel elements it uses a super hot ball of uranium vapor which is dense enough and surrounded with enough moderator (neutron reflector) that it still undergoes nuclear fission. The fission produces huges amounts of thermal radiation, which heats the hydrogen propellant. The fissioning uranium is like a nuclear "sun" in the center of the engine. The reaction chamber directs a flow of propellant around the sun to be heated.

Since this is using the concentrated energy of fission there is no real limit to the thermal energy generated (think nuclear weapons). Unfortunately there is a limit to the hydrogen propellant's ability to absorb heat. Any heat that the hydrogen fails to sop up will hit the engine walls. If this unabsorbed heat is more than the heat radiator can cope with, bye-bye engine. This puts the upper limit on the engine's Isp capability.

Cavity Linergraphite +
5% niobium
Moderatorberyllium oxide
5.07×107 to
20.34×107 N/m2
10% by weight
0.46 m
Cavity Liner
0.0063 m
Engine Cavity
2.44 m
Plasma Dia
1.8 m
Plasma Vol
3.04 m3
Critical Mass
21 kg
Engine Mass
(including 235U)
40,000 to
210,000 kg

The engine is spherical. The outer layer is the pressure vessel (since both the propellant and uranium gas needs lots of pressure to make this thing work), a layer of beryllium oxide (BeO) moderator (a neutron reflector to help the uranium undergo nuclear fission), and an inner porous slotted cavity liner that injects the cold propellant to be heated. In the center is the furious blue-hot atomic vortex of uranium plasma.

Sadly, this structure does suffer from waste heat:

[1] a bit under 0.5% of the reactor power gets to the slotted cavity liner from thermal radiation emitted by the hot propellant. Which is a problem but not a major one. Most of the thermal radiation is soaked up and removed by the propellant.

[2] A whopping 7% of the reactor power hits all three layers of the engine, because part of the fission output is in the form of gamma-rays and neutrons, instead of useful thermal radiation. Hydrogen propellant does not do zippity-doo-dah to soak up gammas and neutrons, all of it sails right through the propellant to hit the engine structure. Deep inside the engine structure, gamma-rays and neutrons are more penetrating than x-rays.

This waste heat is managed by the engine heat radiator (and a bit managed by regenerative cooling, about as effectively as a 3-year-old helping Daddy wash the car). Most of the engine is the beryllium oxide moderator. It is designed to operate at 1,400 K, which is below the 1,700 K melting point of the BeO but above the 1,100 K radiator temperature (otherwise the radiator will refuse to remove the heat).

The hydrogen propellant is pumped into the engine at about 5.07×107 to 20.34×107 newtons per square meter (which is why the engine needs a pressure vessel).

As it turns out hydrogen propellant is transparent, which means it is lousy at absorbing thermal radiation. That's not good. To remedy this sad state of affairs, it is "seeded" by adding tiny metal bits about the size of particles of smoke, about 5% to 10% seeding material by weight. This is done right before the propellant exits the porous cavity liner into the flood of heat from the nuclear vortex. The seeding absorbs all the thermal radiation and passes the heat to the propellant by conduction. The seeding material will be something like graphite, tungsten, or non-fissionable uranium 238.

Around the exhaust nozzle the seeding concentration will have to be increased to 20% to protect the nozzle from propellant heat. The cold 20% seeded hydrogen will reduce the specific impulse a bit but it has to be done.

The porous cavity liner (in some as yet to be defined manner) magically sets up flow patterns so that the propellant flows around the hot uranium and exits via the exhaust nozzle. Meanwhile miraculously the uranium is trapped in a stagnant cavity in the center so hideously radioactive fissioning uranium does not escape through said exhaust nozzle. Uranium escape not only exposes the crew to deadly radiation, it is also a criminal waste of uranium (that is, it lets get away uranium that is not contributing to the engine's thrust).

The interior of the engine (cavity diameter) is 2.44 meters in diameter (7.61 cubic meters), and the incandescent ball of violently fissioning uranium is planned to have a diameter of 1.80 meters and a volume of 3.04 cubic meters. This gives a fuel-to-cavity radius ratio of 0.74. The idea is for the uranium sphere to be 40% of the volume of the entire chamber. However since hydrogen propellant is going to diffuse into the atomic vortex, the uranium sphere might be up to 50% hydrogen. This means the effective volume of pure uranium will be closer to 20% to 30% of the entire chamber.

The uranium can be injected by pushing a very thin rod of solid uranium into the chamber. The uranium penetrates the BeO moderator inside a tunnel lined with a cadmium oxide neutron poison, because otherwise there would be a nuclear explosion once the uranium was surrounded by BeO. This is a bad thing. The engine was designed to have the nuclear reaction happen in the core of the chamber, not in the walls.

As the uranium rod enters the chamber, the heat of the fission ball vaporizes the rod so the fresh uranium atoms can join the party.

A problem is how to get the process started. At startup, there ain't no ball of fissioning uranium to heat up the rod. The report says that the engine will have to be started by first blowing in some hydrogen and somehow injecting some powered uranium metal into the stagnant cavity until it reaches critical mass. Sounds tricky to me.

Figures 2a through 2c above are for a reactor of the following characteristics:

  • Spherical geometry
  • Uranium-235 fuel
  • Beryllium-oxide (BeO) moderator
  • Fuel-to-cavity radius ratio 0.67
  • Cavity liner thickness 0.63 cm
  • Cavity liner graphite + 5% niobium

Figure 2a shows that the 235U critical mass ranges from 10 to 35 kilograms for the cavity diameters and moderator thicknesses considered (all the curved lines are more or less above the 10 kg line and below the 35 kg line). Now for a given cavity diameter, you can reduce the critical mass required by adding more BeO neutron reflector. This means the pressure inside the engine can be lowered, which means the mass of the pressure shell can be lowered. Alas the increased penalty mass of the BeO moderator more than offsets the mass saving on the pressure shell.

(Elsewhere the report notes an optimal thickness of BeO to be 0.46 m, and a cavity diameter of 2.44 m. Eyeballing the graph implies a critial mass of 21 kg of 235U.)

Figure 2b shows that if the BeO moderator thickness is fixed, increasing the cavity diameter will decrease the critical density (the curved line will be closer to the bottom of the graph). Not shown in the table is the unfortunate fact that increasing the cavity diameter also has the side effect of increasing the total BeO weight.

(Elsewhere the report notes an optimal thickness of BeO to be 0.46 m, and a cavity diameter of 2.44 m. Eyeballing the graph implies a critial density of 18 kg/m3. If the uranium plasma ball has a volume of 3.04 m3, at that density it will contain about 55 kg of uranium, which is more than the 21 kg (from eyeball value above) it needs for criticality. However, since propellant seepage will make the sphere about 50% hydrogen, this means it will have about half of 55 kg. Which is a reasonably close eyeball value to a second eyeball value. I'm just playing number games with the graphs, do not put too much credence to these speculations on my part.)

Figure 2c shows that there is an optimum BeO moderator thickness which gives a minimum critical density for a given BeO moderator weight.

Why is there an optimum BeO moderator thickness?

If the BeO is too thin there is excessive neutron leakage (the purpose of the BeO moderator is to reflect escaping neutrons back into the fissioning uranium, basically kicking the out-of-bounds neutrons back into play). Excessive neutron leakage means the blasted cavity diameter will have to be extremely large to avoid very high critical densities.

If the BeO is too thick, the total BeO weight becomes very large. Even though you can get away with smaller cavity diameters without the heartbreak of very high critical densities.

Figure 2c is telling you that the optimum BeO thickness is 0.46 meters (for a reactor of the specified characteristics). 2c goes on to tell you that above a moderator weight of 40,000 kg larger cavities only give a slight reduction in the critical density (the curved lines are almost horizontal).

So all the engine weight estimates below are assuming a BeO thickness of 0.46 meters.

Experiments show that an effective fuel volume is about 20% to 30% of the cavity volume, for a uranium flow rate less than 1% of the hydrogen flow rate.

The paper assumes the engine can accelerate at about 0.01 to 0.05g (0.098 to 0.491 m/s)

The idea is to get the maximum thermal radiation from the fissioning atomic fireball into the cold hydrogen propellant, and the minimum thermal energy escaping the hot hydrogen propellant (which reduces the specific impulse and scorches the heck out of the cavity wall).

Figure 5b shows experimental data for tungsten-seeded hot hydrogen. It says that adding just a few percent by weight of tungsten will increase the thermal absorption cross section to between 2,000 to 100,000 square centimeters per gram. The figure also shows the thermal absorption increase at elevated pressure, which is a good thing since the engine is a high-pressure rig.

These cross sections are high enough to protect the cavity wall from damage for Isp from 4,000 to 7,000 seconds.

Figure 6 is the straight dope on the gas-core NTR engine parameters. The critical density of uranium given cavity size and moderator is as per figure 2. Thermal absorption of seeded hydrogen is as per figure 5. Heat tranfer analysis is used to determine maximum specific impulse that will keep heat load on cavity wall below 1,000 K. Engine pressure is whatever is required to have a critical mass of uranium.

The engine weight is assumed to be the sum of the three major components: BeO Moderator, Pressure Shell, and Heat Radiator. Plus 4,000 kg or less for the uranium fuel.

Pressure Shell assumes a strength-to-density value of 1.7×105 N-meters/kg.

Heat Radiator assumes a unit weight of 140 kilogrmas per megawatt of radiated power. Heat depostion rate is assumed to be 7% of reactor power. Heat radiator operates at 1,100 K (instead of 945 K), which reduces the required radiator surface area by a factor of 2. This kind of radiator more than doubled the specific impulse without adding enough weight to offset the gain. Future radiator designs with even lower unit weights would give even more specific impulse gains.

  • Spherical geometry
  • Uranium-235 fuel
  • Beryllium-oxide (BeO) moderator
  • Beryllium-oxide (BeO) thickness 0.46 meters
  • Fuel-to-cavity radius ratio 0.67
  • Fuel volume 30% of cavity volume
  • Uranium loss rate is 1% or less of hydrogen flow rate
  • Cavity liner thickness 0.63 cm
  • Cavity liner graphite + 5% niobium

In figure 6, the abscissa for both charts is engine thrust. The charts are for thrust levels from 20,000 to over 400,000 Newtons.

The ordinate of the upper chart (Fig 6A) is specific impulse, engine weight for lower chart (Fig 6B). Specific impulse ranges from 2,500 to 6,500 seconds. Engine weight ranges from 40,000 to 210,000 kg.

The curved lines are Engine Pressure, for ranges between 0.5×108 to 2.0×108 N/m2. Note in Fig 6A the three curves are labeled "Low", "Nominal", and "High". These labels are used in the Mission Chart below.

A higher engine pressure allows higher specific impulse because higher pressure makes the hydrogen propellant more opaque. But higher pressure also makes the engine heavier.

Higher thrust increases the specific impulse because there is more propellant flow to cool the cavity wall (note this is the opposite of what occurs when shifting gears). But this also makes the engine heavier.

The two reason above are why it is impossible to chose the "best" engine. What you have to do is specify a specific mission in order to have enough determining factors to figure which engine would be best.

The spacecraft is composed of a gas-core engine (with heat radiator and uranium fuel), a command module, payload, various jettisonable liquid hydrogen propellant tanks , and interconnecting structure.

The engine provides four burns:

  1. Terra orbit escape/target planet trajectory insertion
  2. Target planet orbital capture
  3. Target planet orbit escape/Terra trajectory insertion
  4. Terra orbital capture

After each burn the associated empty propellant tanks are jettisoned, except for the last burn. This is because the command module is attached to the last tank, and the crew would object strongly to being cast off into deep space. The command module also relies upon the hydrogen in the last tank for extra engine-radiation shielding.

Initial Mass In Orbit
Command Module50,000 kg
Payload to Planet150,000 kg Science/Exploration
0 kg Courier
Expendables50 kg/day
Propellant Tankage20% of hydrogen mass
Interstage Structure2% of transmitted load
Thrust Frame5% of thrust
Gas-core Engineas per Figure 6, including uranium storage and supply
Parking orbits600 km circular at Terra
high ellipse at target planet
Propulsive Effortideal ΔV from ref. 19
gravity-loss corrections Cg from ref. 20
Propellant Fraction1 - exp(-((ΔVi * Cg) / (Isp * g0)))
Ref. 19. Fishbach, L. H., Giventer, L. L., and Willis, E. A., Jr., "Approximate Trajectory Data for Missions to the Major Planets," TN D-6141, 1971, NASA, Cleveland, Ohio.
Ref. 20. Willis , E. A., Jr., "Finite Thrust Escape from and Capture into Circular Elliptic Orbits," TN D-3606, 1966, NASA, Cleveland, Ohio.

Propellant Fraction equation comes from combining these four equations into one big equation:

Pf = 1 - (1/R)
R = ev/Ve)
1/ex = e-x
Ve = Isp * g0


Cg = gravity-loss corrections Cg from ref. 20
Δv = ship's total deltaV capability (m/s)
ex = antilog base e or inverse of natural logarithm of x.
g0 = acceleration due to gravity = 9.81 (m/s2)
Isp = specific impulse (seconds)
Pf = propellant fraction, that is, percent of total rocket mass M that is propellant: 1.0 = 100% , 0.25 = 25%, etc.
R = mass ratio (dimensionless number)
Ve = exhaust velocity (m/s)

In the charts below

  • SCNR: Solid-Core Nuclear Rocket (an old-fashioned NERVA)
  • REGEN GCNR: Gas-Core Nuclear Rocket cooled with Regenerative Cooling (choked down to avoid need for heat radiators)
  • RAD GCNR: Gas-Core Nuclear Rocket cooled with Heat Radiator (uses heat radiators so it can run full-bore)
  • FUSION: Fusion Rocket (for comparison purposes)
  • SCIENCE/EXPLORATION: A mission where you bring along tons of scientific payload, and stay on Mars for 40 days to do some science.
  • COURIER: A mission with no payload just a Very Important Person. And no staytime on Mars, just a quick unloading/loading and immediate return to Terra.

In many of the charts Initial Mass in Earth Orbit (IMEO) is used to measure efficiency. The lower the IMEO value, the more efficient. Usually because it means lower propellant requirments, and may allow more payload.

Figure 6 shows that the radiator-cooled gas-core nuclear rocket becomes more efficient (higher Isp and lower specific weight) as the thrust level is raised. So the GCNR is best for missions with large payloads and/or big thrust-to-weight requirements. The missions depicted in the charts below were chosen with this in mind.

This chart shows the effect of changing the duration of the mission on the Initial Mass in Earth Orbit (IMEO). You want IMEO to be as low as possible. The shorter the mission duration, the more propellant you have to pack to increase ΔV, so the higher IMEO becomes. Obviously you can lower IMEO by increasing the mission time, but who wants to spend years on a Mars mission?

The scientific missions assume a 40 day stay on Mars to do science stuff.

The patheticaly weak SCNR (NERVA style solid-core nuclear rocket, shown with yellow curved line) has minimum mass at around 500 days and 1.5×106 kg IMEO (very roughly). This wimp ain't gonna manage a trip time below 400 days, not with a practical IMEO it isn't.

The first gas-core nuclear rocket (green curved line) show an immediate performance improvement. This is the gas-core with no heat radiator, deliberately throttled down so it can make do with mere regenerative cooling. If it is given the SCNR's 1.5×106 kg IMEO, it can do the mission in half the time, only 250 days. Its lowest IMEO is about 0.7×106 kg (700 metric tons) with a mission time around 480 days.

But the other gas-core rocket is even more powerful.

The gas-core nuclear rocket with a heat radiator (blue curved line) lowest IMEO is 0.4×106 kg (450 metric tons). This is only twice the payload (150 tonne payload + 300 tonnes = 450 tonnes). If it is loaded at a IMEO of 0.7×106 kg (the regenerative GCNR's minimum) it will do the mission in 250 days flat instead of 480 days.

With performance this high, the 40 day stay on Mars becomes an appreciable fraction of the total mission time. However low transit times mean high ΔVs and high propellant fractions.

So we now present "courier mode." This has a zero day stay on Mars, instead it immediately turns around to return to Terra. No payload either, except for something way under 1 metric ton (like a Very Important Person or a box of serum to treat the Martian Anthrax-Leprosy Pi epidemic.). The entire mission is nothing but Terra/Mars transits.

A gas-core rocket with radiator on a courier mode mission (hot pink curved line) has truly jaw-dropping performance. It can do an entire mission in only 80 days!

Just for comparison sake, the report includes a fusion rocket with typical high specific impulse but miniscule thrust (orange curved line). The fusion ship has a power plant specific mass ("alpha" or "α") at a very advanced 1 kg/kW. It has extremely low IMEO's if the mission time is greater than 250 days. But below that mission time the fusion ship's performance is lackluster. This is because the fusion drive is low thrust and is power-limited. In order to accelerate up to cruising speed in sometime less than a decade it has to increase its thrust at the expense of the specific impulse. Which sends its IMEO skyrocketing.

Unlike the fusion drive, the radiator-cooled gas core nuclear rocket is not power-limited, it is specific impulse limited (as shown in Figure 6A, see how it rapidly reaches a plateau?). This means if it trades thrust for specific impulse, it isn't reducing the specific impulse very much at all. It can crank up the thrust so it gets up to cruising speed in only two or three days. Then it can drop down to high specific impulse fuel economy gear for the rest of the 80 day mission, at a vast savings in IMEO.

Actually one can calculate the functional equivalent of α for the gas-core drive by using Figure 6. Thrust power is:

Fp = (F * Isp * 9.81) / 2

where Fp is thrust power in watts and F is thrust in Newtons. Divide Fp by the engine weight We' to get the engine α. When you do that with Figure 6, all the engines have an α in the range of 0.01 to 0.1 kg/kW, which makes the fusion drive look like a hippo.

Since these rockets were designed to be reusable, it is important to look into the difficulty of refurbishing one for a new mission.

Insipid solid-core nuclear rockets are woefully weak, but at least their nuclear fuel elements don't go anywhere. They stay safe inside the reactor ready for the next trip. Gas-core on the other hand have the drawback that the nuclear fuel elements eventually spew out the exhaust nozzle. The gas-core rocket's uranium requirement for one mission may be considerably less than the solid-core. Unfortunately the solid-core can re-use its uranium several times before more has to be added, while the gas-core has to restore its entire supply with each mission.

In figure 9 the H/U numbers are Hydrogen-Uranium flow ratios. So for instance, a rocket with a H/U of 200 will expend 200 units of hydrogen propellant for each single unit of uranium. The green SCNR curved line has no H/U number, it is a solid core rocket so zero units of uranium are expended regardless of the hydrogen flow (unless there is a catastrophic engine malfunction).

The family of yellow lines of the scientific/exploration missions show several flow ratios. There is only one flow ratio for the courier mission (200), the one in orange.

Since these are ratios you can take the uranium fuel requirement, multiply by the flow ratio, and thus calculate the hydrogen propellant requirement. For example, the 80 day Mars courier mission requires 3,350 kilograms of weapons-grade uranium-235 (98% enrichment) at a H/U of 200. Therefore the hydrogen propellant requirement is 3,350×200 = 670,000 kg.

Due to the fact that solid-core rockets can re-use their uranium a few times, a gas-core needs a H/R of 200 or more to have a lower uranium fuel bill. In 1971 (when the report was written) uranium fuel was roughly $10,000US/kg. Which means the 150-day Mars courier mission, needing 1000 kg of the hot stuff, has a uranium bill of about ten million dollars.

Not that the hydrogen propellant is exactly cheap, mind you. The element is inexpensive but shipping it from the ground into LEO can make the price tag for the 200,000 kg of propellant somewhere between $44,000,000 and $440,000,000US. This is why space fans are so keen on things like space elevators and in-situ resource utilization, to reduce these outrageous costs.

The preceeding charts assume that the spacecraft uses the optimum thrust level given the mission time and engine. This is shown in figure 10.

If low IMEO missions are desired, the thrust should be within the range of 70,000 to 90,000 Newtons (green area, favoring the right side of each curve). For low mission durations ("fast" missions) the thrust should be within the range of 112,000 to 224,000 Newtons (gold area, favoring the left side of each curve).

This chart shows the effect of using a fixed, non-optimum thrust levels. Since both lines are virtually horizontal the chart is saying there is very effect at all. Over huge ranges of thrust the IMEO doesn't really change.

If you needed a fixed thrust spacecraft that can do both missions, 150,000 Newtons is a good compromise.

But not so fast on choosing 150,000 Newtons.

Remember how shifting gears to increase the thrust imposes a penalty on specific impulse? Well, gas-core rockets with heat radiators laugh at your puny Isp penalties (the technical phrase is "relatively insensitive to Isp penalties").

In the chart, look at the area between "Low" and "Nominal". Notice how the 112,000 Newton curve is far more steep than the 224,000 Newton curve. True a gas-core is relatively insensitive to Isp penalties, but the 112,000 N engine is the more senstive of the two. Lower its Isp and the IMEO penalty mass shoots up to ugly levels.

In light of this information, a fixed thrust spacecraft that can do both missions was given a compromise of 224,000 Newtons.

The paper decided to look beyond Mars to see how the gas-core rocket would handle outer solar system missions. These all use the 224,000 Newton engine.

The science/exploration missions have a 200 day stay time, courier is still 0 day stay time. The chart shows a family of missions for each planet of gradually increasing mission durations, with the first being the courier mission (obviously). The actual feasible missions only occur at 12 to 13 month intervals, so they are marked with squares or circles. There are no missions on the connecting lines, those are just to group the planets and to indicate trends.

The Jupiter courier mission is 1.67 years (600 days) round trip and only requires an IMEO of 1.3×106 kg. The very next mission is a scientific/exploration mission with a 2.75 (1000 day) round trip and an IMEO under 106 kg. This is almost as efficient as the Mars mission.

The Saturn mission IMEOs are almost as good. Of course the trip times are about a year longer (400 days) than the Jupiter missions.

The IMEOs for the Neptune and Uranus missions are very discouraging. This probably means they are better performed with a nuclear-electric, a fusion drive, or other propulsion with a much higher Isp.

GCNR Liberty Ship

RocketCat sez

Ho, ho! This brute kicks butt and takes names! You want to boost massive amounts of payload into orbit? Freaking monster rocket has eight times the payload of a Saturn V rocket. It can haul three entire International Space Stations into LEO all at once!

But to do this it packs seven honest-to-Heinlein nuclear lightbulb engines! The only rocket that could come close to this beast is a full blown Orion drive rising on a stream of nuclear explosions at about one Hertz.

Liberty Ship
ΔV15,000 m/s
Specific Power350 kW/kg
(350,430 W/kg)
Thrust Power560 gigawatts
Specific Impulse3060 s
Exhaust Velocity30,000 m/s
Wet Mass2,700,000 kg
Dry Mass1,600,000 kg
Mass Ratio1.6875
Mass Flow1246 kg/s
Thrust37,380,000 newtons
Initial Acceleration1.4 g
Payload900,000 kg
Length105 m
Diameter20 m wide

Anthony Tate has an interesting solution to the heavy lift problem, lofting massive payloads from the surface of Terra into low Earth orbit. In his essay, he says that if we can grow up and stop panicking when we hear the N-word a reusable closed-cycle gas-core nuclear thermal rocket can boost huge amounts of payload into orbit. He calls it a "Liberty Ship." His design has a cluster of seven nuclear engines, with 1,200,000 pounds of thrust (5,340,000 newtons) each, from a thermal output of approximately 80 gigawatts. Exhaust velocity of 30,000 meters per second, which is a specific impulse of about 3060 seconds. Thrust to weight ratio of 10. Engine with safety systems, fuel storage, etc. masses 120,000 pounds or 60 short tons (54 metric tons ).

Using a Saturn V rocket as a template, the Liberty Ship has a wet mass of six million pounds (2,700,000 kilograms). Mr. Tate designs a delta V of 15 km/s, so it can has powered descent. It can take off and land. This implies a propellant mass of 2,400,000 pounds (1,100,000 kilograms). Using liquid hydrogen as propellant, this will make the propellant volume 15,200 cubic meters, since hydrogen is inconveniently non-dense. Say 20 meters in diameter and 55 meters long. It will be plump compared to a Saturn V.

Design height of 105 meters: 15 meters to the engines, 55 meters for the hydrogen tank, 5 meters for shielding and crew space, and a modular cargo area which is 30 meters high and 20 meters in diameter (enough cargo space for a good sized office building).

A Saturn V has a dry mass of 414,000 pounds (188,000 kilograms).

The Liberty Ship has seven engines at 120,000 pounds each, for a total of 840,000 pounds. Mr. Tate splurges and gives it a structural mass of 760,000 pounds, so it has plenty of surplus strength and redundancy. Add 2,400,000 pounds for reaction mass, and the Liberty Ship has a non-payload wet mass of 4,000,000 pounds.

Since it is scaled as a Saturn V, it is intended to have a total mass of 6,000,000 pounds. Subtract the 4,000,000 pound non-payload wet mass, and we discover that this brute can boost into low earth orbit a payload of Two Million Pounds. Great galloping galaxies! That's about 1000 metric tons, or eight times the boost of the Saturn V.

The Space Shuttle can only boost about 25 metric tons into LEO. The Liberty Ship could carry three International Space Stations into orbit in one trip.

Having said all this, it is important to keep in mind that a closed-cycle gas-core nuclear thermal rocket is a hideously difficult engineering feat, and we are nowhere near possessing the abilty to make one. An open-cycle gas-core rocket is much easier, but there is no way it would be allowed as a surface to orbit vehicle. Spray charges of fissioning radioactive plutonium death out the exhaust nozzle at fifty kilometers per second? That's not a lift off rocket, that's a weapon of mass destruction.

There is an interesting analysis of the Liberty Ship on Next Big Future.

Hariven-class Free Trader

This is not actually "real", but the science is admirably hard.


     “It’s a steel box.”
     “It’s a fully functional – well, mostly functional, but all primary systems are functional – Hariven-class free trader. Just what you want when you’re starting out in this business.”
     “It’s a steel box with a plasma torch welded on the back.”
     “And a generous cargo capacity for its displacement, regenerative life support, ah – adequate crew quarters and food vats, and docking room for a single surface-orbit shuttle.”
     “And it’s –”
     “– a steel box, yes. If you wanted to pay for stylish, would you be shopping for starships in a wreckyard?”
     “Show me the contract again.”

– overheard in Kathar orbit, Cilmínár system


So, I got a request from a reader for a few specs on the Hariven-class free trader. Well, why not?

(Sadly, they were imagining something like Vaughan Ling’s Planetes-inspired debris collector with comparable dimensions, capacity, etc. Sorry to say it, but that ship? Had some style. The Hariven? Really doesn’t.)


Operated by: Desperate free traders, just starting-out bands on tour, your sketchy brother, refugees, space hobos, and anyone else who can’t afford a better ship.
Basic freighter.
Under open-source license; produced by multiple manufacturers, most of whom would prefer not to admit it, along with various backyard fab shops.

(And when I say “desperate free trader”, I don’t mean, say, the people who fly around in a Firefly-class in Firefly. Those people, in this verse, own something like a Kalantha-class. This is down from there at the true ass end of space travel.)

Length: 46m, of which 30m is the hold.
8m (not including radiators)

Gravity-well capable: No.

Personnel: 3, as follows:

Flight Commander
Flight Director
Flight Engineer

(This assumes you’re following the typical regulations which require – since the Hariven has no AI, and only dumb automation – that at least one qualified person be on watch at all times, hence a minimum of three. In practice, a Hariven can be flown by one and very often is, if they don’t mind violating the rules of navigation of every halfway sane polity in space.)

Drive (typical; may vary from build to build): Nucleodyne Thrust Applications “Putt-Putt” fusion pulse drive.
Deuterium pellets. (dirty D-D fusion)
Cruising (sustainable) thrust:
0.6 standard gravities (0.56 g)
Peak (unsustainable) thrust:
1.2 standard gravities (1.12 g)
Delta-v reserve:
(Not yet calculated, but limited; if you’re flying a Hariven, you ain’t going brachy unless you devote a lot of your hold space to extra tanks. Be prepared to spend much of your voyage time on the float.)
Maximum velocity:
0.02 c (based on particle shielding)


Not supplied as standard, but buy some. You’re gonna need ’em.


Orbital Positioning System sensors
Inertial tracking platform
Passive EM array
Short-range collision-avoidance and docking radar



Other systems:

Omnidirectional radio transceiver
Communications laser
Whipple shield (habitable area only)
Mechanical regenerative life support (atmosphere/water only)
Algi-prote vat
2 x information furnace data systems
Sodium droplet radiators

Small craft:

Not supplied as standard, but a common as-supplied variant adds a partition to convert part of the forward hold into a bay with docking clamps suitable for many surface-to-orbit vehicles.


It’s a classic tail-lander layout of the crudest form: a 30m steel box welded on top of an 8m steel cylinder welded on top of a cheap fusion pulse drive, the latter two surrounded by pellet containers. It couldn’t look more brutalist/functional if it tried. At least most Hariven owners try to give it a bright paint job.

The hold is up front, a big steel box roughly the size of eight standard shipping containers. (Indeed, sometimes it’s made from eight standard shipping containers.) Putting it right for’ard has the advantage of simplifying construction greatly – all the machinery is at one end – and giving Hariven captains the assurance that if they ram their junker into anything accidentally, at least there’s 30m of other stuff between them and whatever they hit.

The hold opens up along its entire length on the port side to permit access. Responsible captains who convert their Hariven for passenger transport (the aforementioned touring bands, refugees, and space hobos, for example) by attaching deck partitions inside the hold and adding canned air have these welded shut. Less responsible captains simply pray for a lack of wiring faults.

The habitable section (the cylinder at the back) is wrapped in auxiliary engineering machinery and fuel storage, to the point that it’s only 4m in internal diameter. (If you need to fiddle with most of the engineering systems, you’re going to need a drone, or to take a walk outside.) It’s divided into four decks, from the bow down:

The bridge, which shares space with most of the avionics;

A small living area, which contains the food vat, a tiny galley, the inner door of the airlock, and any luxuries you see fit to squeeze in there. Like chairs;

The crew quarters, which means four vertically-mounted sleep pods, and maybe room for another luxury or two if they’re small;

And a tiny workshop, for any repairs that need doing.

That all sits right on top of the shadow shield and the business end of the drive. If you need to adjust anything below that – well, hope you brought a drone.

But enough of this. You buy this ship, treat her proper, she’ll be with you the rest of your life.

Ain’t sayin’ how long that’ll be, mind.

Ru said:

     Not that long ago, I spent quite some time running the numbers on fusion-pulse torch drives, and working out various performance figures and limitations. It was all quite informative and interesting. You haven’t given nearly enough information about the ship (eg. dry mass) or the engine (exhaust velocity ranges, reaction mass) for me to hazard a guess at its parameters, so it’ll be interesting to see what else you have for us…
     There’s a lot of steel mention in its design. Sounds pretty heavy. Also, metal shells plus charge particle radiation equals bremsstrahlung delight (meaning crew will be constantly irradiated by deadly x-rays) (and it makes for a poor neutron shield, which this sort of drive badly needs). Carbon is probably easier to come by, and much easier to push around.
     You don’t mention reaction mass, but with those performance figures you won’t be using pure fusion for peak acceleration. Presumably the drive expends additional deuterium for that purpose (though lithium might be a better choice).
     You’re using pure deuterium fusion, but that’s a terrible choice for spacecraft fuel, really. For flights much less than the half-life of tritium, D-T offers easier ignition, lower neutron flux and more charged particles to thrust against. D-3He would be the fuel of choice, but if you want stable and conveniently mineable fuel p-6Li or p-11B would be a much better choice than pure deuterium.
     Delta-V reserves for even a fairly conservative fusion spacecraft design are pretty generous. You might not be tooling around at cruising speed for long, but it should be able to sustain a centigee for weeks (or even months if the drive is good enough) at a time with a 3 or 4:1 wet to dry mass ratio.
     (on reflection, I am of course wrong that D-T offers a lower neutron flux or a higher proportion of charged reaction products than D-D, but it does offer a significantly lower x-ray output and a higher exhaust velocity)

Alistair Young said:

     Ah, but you’ve got to bear in mind the target market, and therefore the design paradigm. If this were higher up the scale of starships, it’d have all the fancy carbon-composite hulls, high-efficiency fuel blends, etc., etc., one could possibly desire.
     It’s steel, though, because it’s designed to be repaired – and in some cases, even built – by a monkey with a wrench, a backyard welding kit, and duct tape, not by professional yard dogs with all the nanowhatsits in the catalog.
     (Same reason the neutron protection is a slug of paraffin in the lower hull space rather than proper formed HICAP.)
     Likewise, it uses D pellets to power an old-style fusion pulse drive rather than D-He3 slush to power a new-style fusion torch because that drive needs much less maintenance, any backplanet schmuck can separate deuterium from water, and the calibration is rough enough that in a pinch, you can stuff just about anything that’ll fuse in there and it’ll mostly work for a while.
     (Basically, you want to picture the spacegoing equivalent of the beat-to-hell jalopy that’s been driven around the rainforest for forty years, being fixed with banana peels and duct tape and occasionally run on rough home-cooked rum when gas was short. It’s a sh*tbox, but it’s a sh*tbox that’s hard to kill by design.)

Hedrick Fusion Spacecraft

This section has been moved here


HELIOS Stage One
Thrust12,000,000 newtons
Wet Mass700 metric tons
not including
Stage 2
Dry Mass32 metric tons
Body Diameter6 meters
Wingspan27 meters
HELIOS Stage Two
ΔV21,000 m/s
Specific Power57 MW/kg
(566,100 W/kg)
Thrust Power3.8 gigawatts
PropulsionSolid Core NTR
Thrust981,000 newtons
Exhaust Velocity7,800 m/s
Reactor Power2,600 MW
Wet Mass100 metric tons
Payload6.8 metric tons

HELIOS stands for Heteropowered Earth-Launched Inter-Orbital Spacecraft. Unfortunately "HELIOS" became a catch-all term for quite a few post-Saturn studies around 1963. This entry is about the 1959 version from Krafft Ehricke at Convair.

As you should recall, when dealing with a radioactive propulsion system the three anti-radiation protection methods are Time, Distance, and Shielding. A rocket cannot shorten the time, a burn for specific amount of delta V takes as long as it takes. Most designs use shielding, even though the regrettable density of shielding savagely cuts into payload mass.

But some designers wondered if distance could be substituted. The advantage is that distance has no mass. The disadvantage is it makes the spacecraft design quite unwieldy. You'd have to either put the propulsion system far behind the habitat module on a long boom, or more alarmingly have the propulsion system in front with the habitat module trailing on a cable. In theory the exhaust plume is not radioactive, so again in theory the habitat module can survive being hosed like that. The propulsion exhaust is poorly collimated so it is not like a spacecraft weapon is being directly aimed at the hab module.

There is no way this design would work as a warship. It would be like trying to run through a maze while carrying a ladder. If you made too tight a turn the tow cable will be subject to the "crack-the-whip" effect, the cable will snap, and the hab module will be shot into deep space like a stone from a shepherd's sling.

The break-even point is where the mass of the boom or cable is equal to the mass of the shadow shield. Past that point it is much less trouble just to use a standard shadow shield and deal with the mass.

This is the Waterskiing school of spacecraft design.

Dr. Ehricke design was two-staged. It has a liftoff mass of 800 metric tons, a diameter of 6 meters (omitting the delta wings) and a length of 60 meters.

The first stage was chemical powered since even in 1959 they knew nobody was going to allow a nuclear propulsion system to lift off from the ground. The lower stage has a delta wing, and will glide back to base after stage separation to be reused on future missions. The lower stage has a diameter of 6 meters, and a wingspan of 27 meters. Wet mass of 700 metric tons, dry mass of 32 metric tons, twin chemical engines with a combined thrust of 12,000,000 newtons. The first stage pilot rides in a little red break-away rocket in case the first stage has an accident. In which case it will just be too bad about the crew riding next to the nuclear reactor.

The first stage separates from the second at an altitude of about 50 kilometers when the velocity reaches 4.5 km/s. The corrugated coupler that held the two stages together falls away.

The second stage will use retrorockets to lower the habitat module on cables about 300 meters below the nuclear stage, then let'er rip. The second stage has a wet mass of 100 metric tons, the nuclear reactor has a power of 2,600 Megawatts, and a thrust of 981,000 newtons. Initial acceleration is 1 g.

When it comes to Lunar landing, the habitat module touches down, then the nuclear stage move down and sideways so it stays 300 meters away as it lands. HELIOS can deliver about 6.8 metric tons of payload to the Lunar surface, and stil carry enough propellant to make it back to LEO.

Dr. Ehricke does not give details above the return trip, but it would need to involve some sort of ferry rocket to retrieve the crew from Terra orbit. There is no way anybody would allow that radioactive doom rocket to actually land. Even if it could carry enough propellant. Dr. Ehricke Convair Space Shuttle would do nicely to retrieve the crew.

Nowadays most experts agree that a 300 meter separation from a 2,600 MW reactor is totally inadequate to protect the astronauts from a horrible radioactive death. I've heard estimates of a minimum 1,000 meter separation from a 1 MW reactor. For 2,600 MW you'd want a separation more like 14,000 meters, which probably has more mass than a conventional radiation shadow shield.


Cole Nuclear Pulse
Cole Mod I
Chamber Dia40 m
Chamber Mass454,000 kg
Chamber Wall1.27 cm
Height91 m
67 m
ΔV7,900 m/s
Pulse Rate1 per sec
Pulse Yield0.01 kt
Num Pulse2,400
Water Propel
per pulse
389 kg
Thrust @
1 pulse/sec
3,560,000 N
Isp931 sec
Exhaust Vel9,100 m/s
T/W @
1 pulse/sec
Wet Mass1,611,502 kg
934,400 kg
Dry Mass677,102 kg
Payload159,000 kg
Inert Mass518,102 kg
System Mass
454,000 kg
Structural Mass64,102 kg
Cole Mod II
Chamber Mass90,720 kg
Isp1,150 sec
Exhaust Vel11,280 m/s
Num Pulse5,800
Payload1,325,000 kg
Wet Mass3,048,000 kg
Cole Mod IIa
(x10 scaleup of II)
Chamber Dia86 m
Pulse Yield0.10 kt
Isp1,350 sec
Cole Nuclear Pulse Jet
Pulse Rate2 per sec
Thrust @
2 pulse/sec
42,970,000 N

This is mostly from Aviation Projects Review volume 1 number 3 which has more details than given here. Additional material from Helios pulsed nuclear propulsion concept (1965) which discusses the Lawrence Radiation Laboratory (LRL) Helios.

Again, there were several spacecraft designs that all wanted to use the name "Helios", which is confusing. Almost as many as the designs who all want to use the name "Orion."

This Helios is closely related to the Project Orion designs, in as much as they both used tiny nuclear bombs as propulsion. Sadly the Helios concept had some fundamental design problems that it never overcame.

The basic idea was created by visionary Dandridge Cole who was then working at the Martin corporation. Mr. Cole was unaware of the nuclear-shaped-charge innovation, so he thought the Project Orion design was wasting 90% of the bomb energy. He figured he could do better than that. The more you surround the bomb, the less energy you will waste. Since most material objects fare poorly when hit by a nuclear blast, Mr. Cole used three strategies:

  • The reaction chamber surrounding the bomb was given a huge radius. This spreads the ravening energy of the blast over more chamber wall area, so each square meter of wall has to deal with a smaller portion of the total blast. Keeping in mind that when he said "huge", he wasn't fooling. The first design had a reaction chamber diameter of a whopping 40 meters (130 feet).
  • The bombs were much weaker than the Project Orion pulse units, so the total blast was less. Project Orion units were 1 kiloton, Helios units were 0.01 kiloton, or one hundred times weaker.
  • 390 kilograms of water propellant was injected into the chamber prior to each bomb. The pious hope was that the water would soak up the blast and go shooting out the exhaust nozzle at high velocity, instead of the chamber walls. Hopefully the water would also cool off the chamber walls so they wouldn't melt.

The Cole model I had engine performance that can be charitably described as "disappointing". Specific impulse was 931 seconds, which is in the upper range of conventional solid core nuclear thermal rockets. At one pulse per second the engine had a thrust-to-weight ratio of only 0.25, enough to land on Luna but not enough to lift-off from Terra. By comparison small first generation Project Orion ships were expected to have a specific impulse of 2,500 seconds and a thrust-to-weight ratio of at least 4.0.

One little inconvenient detail that Mr. Cole glosses over is the problem with tiny nuclear bombs. You see, fission reactions have that tiresome "critical" mass requirement. Meaning that if you use less than the critical mass there will be no fission chain reaction. The problem is that a critical mass of uranium-235 or plutonium will ordinarily make a much bigger bang than 0.01 kiloton. Damping the bomb down to 0.01 kiloton means that most of the uranium or plutonium does not enter the reaction. Instead they are merely volatilized into glowing radioactive vapors of death and spread to the four winds at high velocity. This makes it difficult to get permission to launch this monster from Terra's surface.

Even ignoring the radioactive contamination the inefficient use of fissionables is unconscionable. Weapons-grade uranium and plutonium are monstrously expensive, and this design will use tons of the stuff.

A more ambitious (and utterly insane) version was Cole's nuclear pulse jet. This would be a titanic airbreathing version that utilizes Terra's atmosphere as propellant until the ship climbs into space. The radioactive fallout would be only slightly less horrific than that from Project Pluto. The difference was that Pluto was supposed to be a weapon.

Cole and the Martin corp stopped working on the concept in the early 1960s, because of the lack of interest on the part of the USAF, NASA, and Martin higher management. There were a few amusing "artist conceptions" of the concept created by the advertising department of other aerospace companies that wanted to appear new and trendy.



American Bosch Arma Corporation

This is the Atomic Pulse Rocket, a pot-bellied space ship nearly the size of the Empire State Building, propelled by a series of atomic blasts.

The enormous rocket (weighing 75,000 tons fully loaded) is designed to leave Earth with a thrust of 100,000 tons. Altogether a thousand atomic blasts—each equal to 1,000 tons of TNT—are fired from a low velocity gun into a heavy steel rocket engine at a rate of one per second until the vehicle leaves Earth's atmosphere. Then steam and vaporized steel maintain the thrust. After transit speed is reached, and the propulsion system shut off, power is provided by solar batteries plating the wing and body surfaces.

Inside the rocket. living quarters are situated in the rim of a pressurized wheel-like cabin which revolves to provide artificial gravity. Radio and radar antennae revolve with it. Tubular hydroponic "gardens" on either side of the rim grow algae to produce oxygen and high protein food.

The Atomic Pulse Rocket could transport payload to the Moon at $6.74 per lb., less than one quarter the prevailing air freight charges over equivalent distance. A similar project is past the pilot-study stage in the Defense Department

(ed note: This is vaguely based on the Cole study, but is more public relations than a real engineering design study)

Helios Nuclear Pulse
Wet Mass680,000 kg
Payload Mass91,000 kg
ΔV18,000 m/s
Chamber Dia9.2 n
per pulse
100 kg
Pulse Rate1 per 10 sec
Pulse Unit
32 kg

In 1963 the Lawrence Radiation Laboratory started working on their own version under the name of Project Helios. This was for a crewed mission to Mars. Mass in low Earth orbit (IMLEO) was to be 680,000 kg, delta-V of 18,000 m/s, delivering a 45,000 kg Mars lander into Mars orbit (total payload 91,000 kg).

The reaction chamber would have a diameter of 9.2 meters (radius 4.6 m); into which would be introduced 100 kg or so of hydrogen propellant, a small nuclear explosive charge, and a sacrificial positioning framework to hold the nuke at the center. This will be added with each detonation, at intervals of 10 seconds or longer. Of the hydrogen propellant, nuke, and framework mass; the fraction that is hydrogen propellant is called χ.

The nuclear pulse units were one meter in diameter. The core is a 2 kg sphere of weapons-grade uranium. It is coated by 5 kg of high density chemical explosive, and the entire clanking mess is jacketed by 15.7 kg of low density chemical explosive. The nuclear explosive yield is a miniscule 0.0051 kilotons (5.1 tons).


The nozzle sticking out of the chamber is conical with a 20° half-angle. The mass of the nozzle is approximately:

MN: mass of nozzle
k: a constant, report does not specify its value
ε: area expansion ratio of the nozzle
p0: initial pressure within the chamber
rt: radius of the nozzle throat
(ρ/σ)N: weight/strength ratio for nozzle material

Pressure Vessel

The minimum mass of a spherical pressure vessel that can withstand a steady internal pressure p without exploding into a zillion pieces is:

A factor of 4 is then included because the engine is NOT subject to a steady pressure, the pressure pulsates. Then an additional safey factor of 2 is added. So the equation becomes:

Ms: mass of pressure vessel
V: cavity volume in the shell
ρ: density of shell material
σs: shell material yield stress
p: steady pressure
p0: initial pressure within the chamber


The analysis used the payload mass MF to "hide a multitude of sins." It includes the mass of the nozzle throat valve, shock absorbers, shadow shields, life support, observational equipment, Mars excursion vehicle, and Terra atmospheric reentry vehicle. They figure that the sum of the nozzle throat valve, shock absorbers, and shadow shields will come to a total of less than 9,100 kg.


The analysis assumes that the liquid hydrogen propellant will require an additional 8% of propellant mass for tanks, insulation, and boil-off. The ratio of hydrogen tankage mass to useful hydrogen mass is α.

Nuclear Charges

Each nuclear charge and the sacrificial positioning framework is assumed to have a combined mass of 32 kilograms. There will be an additional 2.3 kg per unit for storage and handling in the pulse unit magazine. The ratio of the charge storage/handling mass to the total mass of the charges is β

Total Mass of Vehicle, Propellant, and Nuclear Charges

Remember that each pulse start with the pressure chamber containing hydrogen propellant, a nuclear pulse unit, and a sacrificial framework holding the nuke at the chamber center. The nuke and the framework will be volatilized in the explosion, and the volatilized gas plus the propellant will be heated and sent out the exhaust nozzle. Of the combined mass, the fraction that is hydrogen propellant is called χ.

δMH: mass of hydrogen propellant
δMc: mass of charge debris: volatilized nuclear charge and sacrificial framework
χ: propellant fraction

If it takes N pulses total to accelerate the vehicle to the mission delta-V, then the total amount of effluent mass is:

Remember that the ratio of hydrogen tankage mass to useful hydrogen mass is α and the ratio of the charge storage/handling mass to the total mass of the charges is β

Total Initial Vehicle Mass

Using the equation to determine mass ratio (μ) from delta-V and specific impulse (or exhaust velocity) we can make an equation that will spit out the total vehicle mass (M0) given the mission delta-V (ΔV) and engine specific impulse (Isp)

Combining the effluent mass equation and the total vehicle mass equation we can create three new equations:

Vehicle "Cost"

The cost of the vehicle is assumed to be $91 per kilogram (cost of delivering vehicle components into LEO) plus $50,000 per nuclear charge. Both in 1960 US dollars.

Above graph is number of pulse units (N) vs plenum chamber radius (r). Superimposed on top is a grid of chamber pressure (p) vs propellant-to-total-chamber-contents fraction (χ).

Plotted are the family of curves for vehicle cost COST (109$) in units of billion of 1960 US dollars.

For this chart the constants are:
  • Payload Mass (MF) = 9,100 kg
  • Nozzle expansion ratio (ε) = 200
  • Chamber temperature (T) = 6000 K
  • Delta-V (ΔV) = 18,000 m/s

The cost curves close on the left because the mass of the chamber increases rapidly with pressure, due to the thick-shell correction.

The cost curves close on the right because the enthalpy and specific impulse decrease with decreasing pressure for a fixed expansion ratio and initial temperature.

Hermes from The Martian

The Martian movie is based on the novel of the same name. Both have the Atomic Rockets Seal of Approval. Enough said.

Warning: this section contains spoilers for the novel and the movie.

Hermes (Novel)
PropulsionIon drive
Acceleration0.002 m/s2
Gravity0.4 g
Gravity TypeBola Spin

Hermes in the Novel

Author Andy Weir based the original mission on Robert Zubrin's Mars Direct proposal. Weir updated Zubrin's chemical rocket to a nuclear-reactor-powered ion drive using argon propellant. You see, a puny chemical rocket has to use Hohmann transfer orbits which have launch windows tied to the synodic period of Mars. That mission would have had a required stay time on Mars of a little over a year. For dramatic reasons, Weir needed the mission capable of being aborted at any time with a return to Terra. The ion drive allowed this.

In Andy Weir's original conception, the Hermes is cone-shaped so it can aerocapture at Mars and at Terra, saving precious propellant mass.

The Hermes has an acceleration of 0.002 m/s2 (2 millimeters per second, per second). Andy Weir said that the delta V budget for the return trip was about 5,000 m/s.

The spacecraft is split down the middle parallel to the long axis. This allows the two halves to separate, attached with cables, so they can spin like a bola to provide artificial gravity. The halves are called "Semicone-A" and "Semicone-B".

The main airlock/docking port is located in the customary place, on the nose.

Andy Weir mentioned that the movie version of the Hermes has quite a different design. But he also noted it was "way cooler-looking than the version I imagined."

flight plan
Terra to Mars124 days
Surface Ops31 days
Mars to Terra241 days
Sol-6 Abort
flight plan
Terra to Mars124 days
Surface Ops6 days
Mars to Terra236 days
flight plan
Terra to Mars124 days
Surface Ops6 days
Mars to Terra236 days
Terra Slingshot0 days
Terra to Mars Flyby322 days
Mars to Terra211 days
The Martian: A Technical Commentary

An Ares mission begins with 14 uncrewed launches (probably with an Atlas or similar sized booster) dropping airbag-cushioned payloads on Mars. These would each weigh about 1000kg on launch, with up to 600kg of payload to the surface. This includes parts of the Hab and supplies.

The crewed part of the mission is mediated by the Hermes, a large vehicle for deep space with a nuclear powered ion drive designed to fly between Earth and Mars and back. The Hermes is used by every mission and was assembled in Earth orbit at (no doubt) astronomical expense.

The six astronauts of Ares 3, together with their supplies are launched from Earth to Hermes. The Mars Descent Vehicle is launched separately towards Mars at about this time.

Hermes and the MDV travel to Mars, parking in Mars orbit after 124 days in deep space. Hermes remains in orbit, uncrewed, while the MDV flies the astronauts to the surface.

On the surface, the astronauts build their Hab from airbag cushioned cargo drops and perform their mission. After 30 days on the surface, they climb into the Mars Ascent Vehicle and fly back up to orbit, where they meet the Hermes and fly back to Earth, taking 208 days to return.

Hermes parks in Earth orbit and the crew return to Earth in some re-entry vehicle like Orion or Dragon.

The MAV was launched years before, made its return fuel on Mars using electricity and ambient atmosphere, then was used for about 6 hours to get back to the Hermes.

This mission architecture is very credible, given a nuclear powered ion drive, which is technically possible but politically problematic. IMO, the architecture is inefficient given most of the hardware is used only once, Hermes is not self sufficient, and the astronauts spend only 30 days on the surface.

Mars Ascent Vehicle (MAV) is quite large. It had to be soft landed, but even empty weighs much more than all 14 presupply missions combined. Given that NASA has (in the story) developed soft-landing capability for tens of tonnes, it's not clear why stuff as mission critical as the Hab is landed relatively inaccurately in lots of parts. It could be that this simply reduces mission cost and complexity, or that there was no practical way to land something as bulky as the hab (even disassembled) in one piece.

The MAV employs In-Situ Resource Utilization (or ISRU) to make fuel and oxidizer for the return flight. Two (Earth) years of power from a 100W Radioisotope Thermoelectric Generator (RTG) is enough to make 13kg of fuel (methane) and oxidizer (oxygen) from every 1kg of hydrogen (H2) precursor brought from Earth, for a total of nearly 20T of fuel.

At various points of the novel, Weir describes the MAV as weighing 32 metric tons when fully fueled, and standing 27m tall. This implies that it is very long and skinny, which is unnecessary in the thin Martian atmosphere. Not only that, this means a lot of rocket mass relative to the amount of fuel it can carry (spherical rockets are vastly more efficient, absent significant atmosphere). Needless to say that's a bad thing. By comparison, the Falcon 9, a long and skinny rocket by usual standards, is about 70m tall and weighs about 600T on launch. The MAV could easily be a conical shape perhaps 5m wide and 10m tall.

Weir states that it has two stages, though one stage is perfectly adequate for the relatively low delta-V required to reach Low Mars Orbit (4.1km/s). Nevertheless, with a 325s Isp methane-oxygen engine, a two stage system would have a 16T first stage, a 8T second stage, and a 8T orbital module, with an implied mass fraction of 81% fuel vs 19% metal in each stage.

Towards the end of the novel, engineers at JPL describe the MAV as having an unrealistically low launch weight of 12,600kg (12.6T) — similar to a fully-loaded Dragon capsule. So we'll assume this is the dry mass. Let's assume, then, that the orbital module is 8T, the first stage is 3T, and the second stage is 1.6T, empty. The 19.397T of fuel is distributed accordingly, implying an engine Isp of 405s in order to reach 4.1km/s of Low Mars Orbit. This is low for H2/O2 engines, but extremely high for a methane-oxygen engine. Even SpaceX's planned monster Raptor engine has a notional vacuum Isp of 380s.

In order to get to 5.8km/s and intercept the Hermes, the mass of the orbital module needs to be reduced from 8000kg to 4280kg, a reduction of 3720kg. This takes into account adding 780kg of fuel, removing 500kg from the first stage (pulling off an engine), and so on. The accuracy of the numbers indicate that Andy Weir did the math, but it's not clear on what metrics he designed the MAV and its launch system.

More generally, given that the total delta-V needed to get from Mars to Earth is *only* 7.8km/s, a MAV that flies all the way back to Earth is completely possible, though it would probably need to be bigger than the MAV presented here to have adequate life support. But given that the fuel/delta-V is most easily obtained on the surface of Mars, rather than brought from Earth, a direct ascent architecture actually makes a lot of sense.

On Sol 68, Watney points out that NASA never used large RTGs on crewed missions before Ares, but during the Apollo program RTGs were deployed by astronauts to power lunar seismometers. On Sol 69, Watney states that Lewis had buried the RTG for safety reasons. A RTG stashed somewhere on the surface, however, is much less likely to overheat.

Mars Descent Vehicle (MDV). In the Sol 7 log entry, Watney mentions that the Mars Descent Vehicle (MDV) is useless to him for escaping, since its thrusters cannot even lift its own weight. This, of course, refers to its weight when fully fueled. Before landing, much of its fuel has burned off and it can achieve neutral thrust for a hovered landing. Nevertheless, it lacks (by far) the fuel capacity, thrust, and delta-V necessary to fly anything back to orbit!

In Chapter 8, Bruce and Teddy discuss potential MDV modifications. It is strongly implied, though not stated, that the design would not admit the addition of more engine clusters, and they don't have the time to invent a bigger engine. It is likely that this is a narrative device.

Orbital Mechanics

On the first page, Watney states that he was six days into the best two months of his life. Evidently he was confused, as later it's made clear that the surface operations of the mission were only 30 days long.

30 days on the surface is the extent of surface stay permitted under an "opposition" class mission, wherein the astronauts fly by Venus on either the outbound or inbound leg. While shortening the overall mission, opposition class missions significantly lengthen the time spent in space, and also bring the spacecraft much closer to the Sun, increasing the crew's exposure to radiation.

The alternative mission design is the "conjunction" class mission, wherein the crew takes a relatively short 4-6 month Hohmann transfer flight either way, with a ~560 day stay on the Mars surface in between launch windows. Obviously, if Watney had been stranded on a conjunction mission, he would have had no shortage of snacks!

One additional detail is that Weir's spaceship, the Hermes, employs ion thrusters throughout the mission, enabling a wider class of missions and trajectories than the traditional point-and-shoot orbital mechanics described in the previous paragraphs.

In Chapter 16, the Purnell maneuver is discussed, by which the crew can return to Mars fewer days than the 404 it would take Iris to get there. It's probably worth noting that there is a very similar delta-V requirement for Iris to get to Mars vs a resupply probe to reach the Hermes. The advantage of the Hermes approach is that Iris had to be able to do entry, descent, and landing. If this is the case, Iris could also get there faster by borrowing a basic ion thruster package from, say, the Asteroid Redirect Mission (ARM) spacecraft. It's also not clear why all the crew need to return to Mars (aside from narrative reasons) - most or all could return to Earth in the entry vehicle while Hermes takes the Purnell maneuver to Mars to pick up Watney before he starves. Although the remaining crew would then depend on a new entry vehicle being sent up to meet them on their eventual return to Earth.

In Chapter 20, Annie somewhat incongruously asks why Hermes can't wait at Mars for Watney to get there, when it seems he'll be slowed by the dust storm. Venkat points out that Hermes is on a fly by and can't slow down enough to be captured into orbit, but this is not entirely true. On Sol 505, Bruce says to Venkat that Hermes is flying by Mars at 5.8km/s. Mars escape velocity is only 5.5km/s (the Earth's is 11.2km/s by comparison) meaning that a delta-V of only 300m/s is needed to capture into orbit. Given that Hermes can accelerate at 2mm/s/s, a two day burn would be sufficient to capture into a big elliptical orbit, drastically increasing their margin of error. Perhaps, if Hermes slows enough for an orbital capture, its launch window to return to Earth will close too quickly to be useful.

Of course, the MAV was designed to reach Low Mars Orbit, with a delta-V of 4.1km/s. Getting to 5.8km/s is highly non-trivial, as discussed in the previous section describing the MAV. Of course the unmodified MAV has life support, so Watney could wait while Hermes spirals down to 4.1km/s to pick him up (~30 days, because Hermes can't exploit the Oberth effect), while in the modified MAV he gets close to 5.8km/s, making it much easier for Hermes to rendezvous. A MAV that got to, say, 5.2km/s would split the difference nicely. Either way, the most likely explanation is that maneuvering Hermes to do this would make them miss the Earth launch window.

On Sol 543, Beck mentions that the modified MAV will hit 12gs during launch. While they have lightened the primary payload by about 16%, Watney also removed a spare engine, suggesting that the unmodified MAV would hit at least 10gs during launch, which is unlikely for a rocket designed to fly humans! Later, Johanssen reads out a velocity of 741m/s at an altitude of 1350m, which is staggeringly fast, implying an acceleration of 20.7gs. Perhaps she dropped a zero?

When Johanssen and Vogel talk about getting Watney to orbit, what they mean is solar orbit, since Watney will have to escape Mars entirely in order to be intercepted by Hermes.

During the intercept procedure, Ares 3 crew have to think fast to find additional sources of delta-V to move the Hermes close enough to catch Watney as he flies by. The distances and velocities mentioned during this passage in Chapter 26 are correctly calculated and almost entirely realistic.

Watney suggests making a small breach in his suit and using the stored gas as a rocket to close the velocity mismatch of 42m/s. Assuming he has 5kg of gas on board (including reserve tanks) and an exhaust velocity of 400m/s (unlike rocket exhaust, it's not hot) this confers about 17m/s of delta-V, which is just not enough. This idea is transferred to the Hermes, which will spit out its atmosphere to slow down. Assuming Hermes weighs 100T, it would have to lose about 5T of air to make up the required 29m/s of delta-V. At sea-level atmospheric conditions, this implies that Hermes has a volume of 4000 cubic meters, or a floor area of 1300 square meters, or 13,000 square feet, which is the same as a very large house. Perhaps Hermes has large pressurized volumes that aren't used much for habitation? Martinez estimates that the air will take 4 seconds to leave, which implies a relatively small hole, since the shockwave would take about 0.1s to cross a Hermes-sized volume of air. A realistic concept is that Hermes is composed of two large Bigelow inflatable modules each with a diameter of about 12m, such as the BA 2100 habitats. Also worth mentioning that the process of blowing the "Vehicular Airlock" (VAL) will send lots of airlock fragments into space, hopefully missing Watney.

From The Martian: A Technical Commentary by Handmer, Jermyn, Paragano, Lommen, Nosanov (2015)

Hermes (francisdrakex)
Inert Mass60,000 kg
Crew+Payload Mass7,000 kg
RCS Propellant6,000 kg
Argon Propellant29,000 kg
Dry Mass67,000 kg
Wet Mass102,000 kg
Mass Ratio1.52
ΔV24,000 m/s
EngineIon drive
Exhaust Velocity50,000 m/s
Single Engine Thrust5 N
Engines in Arrayx40
Total Thrust200 N
Acceleration0.002 m/s
Length85 m
Span22 m
Reactor Power10 MWth
Reactor Fuel600 kg 239Pu
Gravity0.4 g
Gravity TypeTumbling Pigeon
Gravity Spin3 rpm
Gravity Radius40 m

francisdrakex's version of Hermes

francisdrakex is a talented space artist who took a stab at designing the Hermes. He did an outstanding job if I say so myself, and not just because he was assisted by some data from this website.

The entire spacecraft was designed to fit inside a 5 m payload fairing for easy boosting into LEO.

His design used the "tumbling pigeon" method of artificial gravity, which is always a good choice to reduce the spin rate below nausea levels.

The ion engine array is mounted at the spin center, a classic technique from Stuhlinger's Ion Rocket.

As per standard best practices, the dangerously radioactive nuclear reactor is mounted as far as possible from the habitat module and the crew. The reactor has a set of heat radiators to reject waste heat. The radiators are in a triangular pattern, to keep them inside the shadow cast by the anti-radiation shadow shield.

The habitat module is a standard TransHab inflatable module.

Hermes (Movie)
Length83 m ?
120 m ?
Gravity0.4 g
Gravity TypeCentrifuge
Gravity Spin4.97 to 6.30 rpm
Gravity Radius9.0 to 14.5 m

Movie version of Hermes

The movie Hermes has the engines mounted aft the reactor, and a centrifuge to provide artificial gravity.

Ship's power is apparently from a set of 12 solar cell arrays, straight off of the International Space Station (the brown elongated rectangles). Which seems a bit redundant if you already have a nuclear reactor.

Rhett Allain did some calculations about the gravity centrifuge on the movie version of Hermes, and not surprisingly discovered that there was a bit of artistic license involved.

The novel states that the artificial gravity is 0.4 g. Mr. Allain did some measurements from the movie and figured the centrifuge is spinning at about 1.08 rotations per minute (0.109 radians per second). Unfortunately to produce 0.4 g the radius of the centrifuge would have to be an outrageous 329 meters! According to one of the graphics in the flight center, the Hermes is only 80 meters long.

Mr. Allain made some further measurements from the movie and concluded the centrifuge was about 9.0 to 14.5 meters in radius. To produce 0.4 g it would have to spin at an angular speed of 4.97 to 6.30 rotations per minute (0.52 to 0.66 rad/second). Which is right at the nausea limit.

Anyway 1.08 rpm is six times slower than 6.30 rpm, which is where the artistic license comes in.

Using the movie figures of 14.5 meters in radius and a spin rate of 1.08 rpm, the artifical gravity would be a pathetic 0.02 g, not 0.4.

Another difficulty is that the spacecraft is supposed to slow down at Mars and Terra by using aerobraking. This will require something like the ballute from 2010 The Year We Make Contact. Two of them, one for each braking.


Human Outer Planet Exploration (HOPE) is from the NASA report TM-2003-212349 by Melissa McGuire, Stanley Borowski, Lee Mason, and James Gilland (2003). . Revolutionary Concepts for Human Outer Planet Exploration (HOPE) { slide show }.

This was a given as a design problem for rocket scientists.

The problem was to design a manned mission to the Jovian moon Callisto, transporting a given payload, and returning the crew and scientific samples back to Terra. The payload included an In-situ resource utilization (ISRU) plant capable of cracking Callistonian water ice into hydrogen and oxygen rocket fuel. They assumed that space probe precursor missions had mapped Callisto's surface so that landing sites could be selected in advance, with due respect toward safety, operations, and scientific gain.

Calllisto was chosen as a destination because it is outside of Jupiter's radiation belts, and it has water ice on the surface for propellant production. The purpose of the mission was to establish an outpost and propellant production facility near the Asgard impact site on Callisto.

Several design teams entered the challenge, each basing their spacecraft around a different propulsion system for comparison purposes. The idea was to promote apples-to-apples comparison, as opposed to the sad proliferation of apple-to-oranges comparisons.

Transportation of the specified payload is left up to the mission designers.

Some designs use several unmanned spacecraft to deliver all the payload except the crew and TransHab module. Those arrive on a separate manned spacecraft, which is only dispatched upon successful arrival of the unmanned spacecraft.

Other designs using more potent propulsion systems have a single spacecraft carrying all the payload.


Mass Breakdown
Crew Quarters
40,000 kg
3,933 kg
3-Person Crew Pod
40,000 kg
Surface Habitat40,000 kg
ISRU Plant40,000 kg

The standard HOPE payload is a TransHab crew quarter for six (including consumables and the crew), a Lander to ferry three crew and supplies to and from the surface of Callisto, a surface habitat module to house the three surface explorer crew members, and an In-situ resource utilization (ISRU) plant. The ISRU plant package includes an ISRU factory to crack Callistian ice into fuel for the lander, a reactor to power the ISRU plant and surface hab, and two rovers.

Some of the designs that use weaker propulsion systems and thus have longer mission lengths use two TransHab modules to reduce risk and increase available storage for the increased consumables required.

Payload: TransHab Module

TransHab Mass
Rad shield
6 crew-year
& Spares
TOTAL w/Cont46,268

This is pretty much a bog-standard TransHab habitat module, right off the shelf.

Crew quarters for six crew. Pressurized volume is about 333 cubic meters. Typically includes 15 metric tons of consumables, but varies according to mission length of the particular design.

Payload: Lander

Lander Mass
Tanks &
Life Support2,025
TOTAL w/Cont25,009

The common base section carrying a three person crew pod. Can transport three crew to Callisto surface and back. It can carry down 40 tons to the surface.

Payload: Surface Habitat Module

Surface Module Mass
Tanks &
Life Support10,779
TOTAL w/Cont36,616

The common base section carrying the inflatable surface habitat module. Can house three crew members on the surface of Callisto. It provides shelter and serves as a laboratory.

Payload: In-Situ Resource Utilization Plant

ISRU Plant Mass
ISRU Plant1,782
TOTAL w/Cont37,909

The common base section carrying the ISRU kit. This lands on Callisto the nuclear reactor, two rovers, and the ISRU plant to crack Callistian ice into LH2/O2 fuel for the lander.

This is a 1 MW-thermal reactor using a Brayton power converter to produce 250 kilowatts of electricity. It supplies power to ISRU plant and surface habitat module. Reactor is sited one kilometer away from habitat due to radiation. Alternatively tractors can be used to create hills out of local material to act as radiation shielding and reduce the mass required for reactor shielding and long cables.

Material on Callisto's surface is about 55% water ice and 45% rock. The ISRU plant will consume 215 kW of electrical power while processing 21 kilograms of water per hour into liquid hydrogen and liquid oxygen fuel for the lander. This will produce enough lander fuel for one lander sortie mission between the base and the orbiting ship every 30 days. The created fuel is stored in the fuel tanks of the common base sections of the ISRU plant and surface habitat. The engines of those common base sections are used as spares in case the lander's engines need repairs.

The rovers are equipped with bulldozer shovels in order to scoop and transport ice to the ISRU plant.


ΔV138,000 m/s
Specific Power38 kW/kg
(37,700 W/kg)
Thrust Power111 megawatts
PropulsionFission Fragment
Payload60,000 kg
Wet Mass303,000 kg
Dry Mass295,000 kg
Propellant Mass4,000 kg
Length120 m
Span62 m
Radiator area6,076 m2
Total Power1 GW
Thrust43 N
Isp527,000 s
Exhaust Velocity5,170,000 m/s

Final Report: Concept Assessment of a Fission Fragment Rocket Engine (FFRE) Propelled Spacecraft . This HOPE spacecraft was designed using a Fission Fragment Rocket Engine.


This HOPE mission concept was based around Magnetoplasmadynamic (MPD) Nuclear Electric Propulsion (NEP).

There are three spacecraft: a one-way tanker, a one-way cargo ship, and a round-trip manned ship (the Piloted Callisto Transfer Vehicle or PCTV).

The tanker is unmanned. It transports to Callisto orbit propellant tanks full of propellant that the PCTV will need for the return trip back to Terra.

The cargo vehicle is unmanned. It transports part of the payload to Callisto: the lander, the surface habitat, and the ISRU plant. Both spacecraft will be dispatched on a slow low-energy trajectory to Calliso.

Only after the unmanned vessels successfully arrive at Callisto (especially the tanker) will the PCTV be dispatched. It transports the rest of the payload to Callisto: the 6 crew, life-support consumables, and the TransHab crew quarters. It will use a fast high-energy trajectory to Callisto (in order to minimize consumables and crew radiation exposure) thus arriving with most of its propellant expended. It will replenish its propellant from the tanker for the return trip.

The habitat module is surrounded by tanks for radiation shielding. The tail radiators are cut in a triangular shape, and the outer heat radiators are arc shaped to keep them inside the shadow shield's radiation free zone, to prevent them from scattering radiation into the ship.

The crew will explore Callisto for 120 days, then depart back home to Terra.

HOPE Cargo vehicle

HOPE Cargo vehicle
ΔV20,600 m/s
Specific Power2 W/kg
Thrust Power430 kW
PropulsionMPD thrusters
Specific Impulse8,000 s
Exhaust Velocity78,500 m/s
Wet Mass242,000 kg
Dry Mass182,000 kg
Mass Ratio1.3
Mass Flow1.4 x 10-4 kg/s
Thrust11 n
Initial Acceleration4.6 x 10-6 g
Payload120,000 kg
Length130 m
Diameter55 m

The purpose of this unmanned vehicle is to transport the HOPE payload elements: lander, surface habitat, and ISRU plant. And a couple of propellant tanks for the benefit of the manned spacecraft.

HOPE Tanker

HOPE Tanker
ΔV20,600 m/s
Specific Power2 W/kg
Thrust Power430 kW
PropulsionMPD thrusters
Specific Impulse8,000 s
Exhaust Velocity78,500 m/s
Wet Mass244,000 kg
Dry Mass184,000 kg
Mass Ratio1.3
Mass Flow1.4 x 10-4 kg/s
Thrust11 n
Initial Acceleration4.6 x 10-6 g
Payload103,000 kg
Length135 m

The purpose of this unmanned vehicle is to transport propellant tanks so that the crew vehicle can refuel at Callisto for the trip home.

HOPE Crew vehicle

Piloted Callisto Transfer Vehicle
ΔV26,400 m/s
Specific Power6 W/kg
Thrust Power1.5 MW
PropulsionMPD thrusters
Specific Impulse8,000 s
Exhaust Velocity78,500 m/s
Wet Mass262,000 kg
Dry Mass188,000 kg
Mass Ratio1.4
Mass Flow3.6 x 10-4 kg/s
Thrust28 n
Initial Acceleration1.1 x 10-5 g
Payload79,000 kg
Length117 m
Diameter52 m

Stuhlinger Ion Rocket

Stuhlinger Ion Rocket
Length150 m
Wet mass360 metric tons
Dry massLander ship: 240 metric tons
Cargo ship:170 metric tons
Lander ship: 120 metric tons
Cargo ship: 190 metric tons
Mass of
Mars Lander
70 metric ton
Storm cellar
50 metric tons
Storm cellar
1.9 m
Storm cellar
Rotation rate1.3 rpm
0.14 g
115 MWt
40 MWe
4,300 m2
75 MWt
thrust98 N

Note the similarity of the HOPE MPD Crew Vehicle to this 1962 Ernst Stuhlinger design for a Mars ion-drive rocket. In both cases the engine are at the ship's middle, with triangular heat radiators.

In the mission plan, the expedition would have three spacecraft carrying a Mars lander, and two without. The astronauts would live in the storm cellars for the 20 days it would take to pass through the Van Allen radiation belts. Earth-to-Mars transfer would span mission days 57 through 204. On day 130 the thrust would be changed 180°, brachistochrone style.


This HOPE mission concept was based around Magnetized Target Fusion engines.

This section has been moved here


This HOPE mission concept was based around Variable Specific Impulse Magnetoplasma Rocket (VASIMR) propulsion.

Revolutionary Concepts for Human Outer Planet Exploration (HOPE) .

The third option utilizes Variable Specific Impulse Magnetoplasma Rocket (VASIMR) propulsion for all vehicles. VASIMR systems heat hydrogen plasma by RF energy to exhaust velocities up to 300 km/s producing low thrust with a specific impulse ranging from 3,000 to 30,000 seconds.

There is significant debate in the advanced propulsion community with respect to the complexity of the engineering challenges associated with the VASIMR system and hence for the purposes of the HOPE study, VASIMR was viewed at a lower state of TRL than MPD thrusters.

VASIMR performance potential was utilized in this option to improve upon the previous option. A single VASIMR propelled vehicle is used to transport the surface systems and return propellant to Callisto as opposed to two. As in the previous scenarios, the tanked/cargo vehicle remains in orbit around Callisto to be used a future propellant depot.

The piloted VASIMR vehicle was fitted with a second TransHab and configured with its main tanks clustered around the rotation axis. The two TransHabs balance each other and are connected by a pressurized tunnel so that the crew can move between them. Like the previous option, there are hydrogen tanks protecting the crew but they do not begin to empty till the last few months of the return mission. The resulting configuration reduces risk by having two crew habitats, the ability to generate artificial gravity throughout the entire mission plus significantly improved radiation protection.

The down side is that the payload masses have gone up due to combining the cargo and tanker vehicles and the piloted vehicle enhancements. The 10 MW that was used for the MPD option is not enough power for the VASIMR option to meet mission requirements. The VASIMR option does close assuming 30 MW on each vehicle resulting in a piloted mission round trip time of around 4.9 years with 32 days at Callisto. The total mission mass is between the previous two options with the benefits of increased safety and robustness.

From the report

HOPE (Z-Pinch Fusion)

This HOPE mission concept was based around Z-pinch fusion propulsion.

This section has been moved here

Hyde Fusion Rocket

This section has been moved here


Source [1]
(Booster + Sustainer)
(to orbit)
145,000 kg
(to Terra escape)
82,000 kg
Stage 1 enginechemical
Stage 1 thrust10,700,000 N
Stage 1 Isp
(sea level)
365 sec
Stage 1
exhaust vel sea
3,580 m/s
Stage 1 Isp
457 sec
Stage 1
exhaust vel vac
4,480 m/s
Wet Mass850,000 kg
Height85.4 m
Diameter8.54 m
Source [2]
(Booster only)
Stage 1 enginechemical
Stage 1
num engine
Stage 1 thrust13,700,000 N
Stage 1 Isp
(sea level)
365 sec
Stage 1
exhaust vel sea
3,580 m/s
Stage 1 Isp
457 sec
Stage 1
exhaust vel vac
4,480 m/s
Wet Mass
394,625 kg
Dry Mass
18,144 kg
Stage 1
Burn Time
70 sec
Height12 m
Diameter8.45 m
Span13 m
Source [3]
(Sustainer only)
Stage 2 engineNTR LH2
Stage 2 thrust5,782,680 N
Stage 2 Isp800 sec
Stage 2
exhaust vel
7,900 m/s
Wet Mass
453,592 kg
Dry Mass
110,000 kg
Height51 m
Diameter8.54 m
Span8.54 m
Source [4]
(Sustainer only)
Num Crew4
Stage 2
39 m
Stage 3
101 m
Total Height140 m
Stage 2 engineNTR LH2
Stage 2 thrust2,700,000 N
Stage 2 power10,000 MW
Stage 3 engineNTR LH2
Stage 3 thrust44,500 N
Stage 3 power170 MW
Leave Terra
propellant burnt
293,300 kg
Arrive Mars
propellant burnt
232,000 kg
Leave Mars
propellant burnt
100,000 kg
Arrive Terra
propellant burnt
21,000 kg
Source [5]
(Sustainer only)
Num Crew4
Trip Time347 days
Height100 m
Leave Terra
Orbital Altitude560 km
Payload26,400 kg
Propellant burnt293,000 kg
Initial Mass721,000 kg
Final Mass428,000 kg
Thrust2,900,000 N
Arrive Mars
Orbital Altitude1,900 km
Payload24,400 kg
Propellant burnt232,000 kg
Initial Mass414,000 kg
Final Mass181,700 kg
Thrust45,000 N
Leave Mars
Orbital Altitude1,900 km
Payload23,200 kg
Propellant burnt100,000 kg
Initial Mass161,000 kg
Final Mass61,010 kg
Thrust45,000 N
Arrive Terra
Orbital Altitude560 km
Payload9,100 kg
Propellant burnt21,000 kg
Initial Mass39,000 kg
Final Mass17,000 kg
Thrust45,000 N

This is from a 1959 study by Krafft Ehricke for Convair. Alas, details are sketchy, and some sources disagree with each other. Indeed some source disagree with themselves. In the table I separate the data as per the sources, so you can be as confused as I am.

The concept is a solid-core nuclear thermal rocket (the "Sustainer") that would do fast reconnaissance to Mars and Venus. A chemical booster lofts it into orbit because even back then NASA was skittish about a nuclear-powered surface-to-orbit stage. The nuclear section is two-staged, with the first stage discarded after trans-Martian insertion. The second stage is used for the three remaining mission segments.

The mission envisioned a fleet of three to four spacecraft, for mutual support.


This section has been moved here

Kuck Mosquito

RocketCat sez

This thing looks really stupid, but it could be the key to opening up the entire freaking solar system. Orbital propellant depots will make space travel affordable, and these water Mosquitos are just the thing to keep the depots topped off.

Kuck Mosquito
ΔV5,600 m/s
Specific Power4.8 kW/kg
(4,840 W/kg)
Thrust Power484 megawatts
PropulsionH2-O2 Chemical
Specific Impulse450 s
Exhaust Velocity4,400 m/s
Wet Mass350,000 kg
Dry Mass100,000 kg
Mass Ratio3.5
Mass Flow49 kg/s
Thrust220,000 newtons
Initial Acceleration0.06 g
Payload100,000 kg
Length12.4 m
Diameter12.4 m

Kuck Mosquitoes were invented by David Kuck. They are robot mining/tanker vehicles designed to mine valuable water from icy dormant comets or D-type asteroids and deliver it to an orbital propellant depot.

They arrive at the target body and use thermal lances to anchor themselves. They drill through the rocky outer layer, inject steam to melt the ice, and suck out the water. The drill can cope with rocky layers of 20 meters or less of thickness.

When the 1,000 cubic meter collection bag is full, some of the water is electrolyzed into hydrogen and oxygen fuel for the rocket engine (in an ideal world the bag would only have to be 350 cubic meters, but the water is going to have lots of mud, cuttings, and other non-water debris).

The 5,600 m/s delta-V is enough to travel between the surface of Deimos and LEO in 270 days, either way. 250 metric tons of H2-O2 fuel, 100 metric tons of water payload, about 0.3 metric tons of drills and pumping equipment, and an unknown amount of mass for the chemical motor and power source (probably solar cells or an RTG).

100 metric tons of water in LEO is like money in the bank. Water is one of the most useful substance in space. And even though it is coming 227,000,000 kilometers from Deimo instead of 160 kilometers from Terra, it is a heck of a lot cheaper.

Naturally pressuring the interior of an asteroid with live steam runs the risk of catastrophic fracture or explosion, but that's why this is being done by a robot instead of by human beings.

In the first image, ignore the "40 tonne water bag" label. That image is from a wargame where 40 metric tons was the arbitrary modular tank size.

There are more details here.


# Engines1
2,700 kg
607s @MR 4
545s @MR 6
5,960 m/s @MR 4
5,350 m/s @MR 6
T/W  9.8 @MR 4
12.1 @MR 6
ΔVExpend: 8,030 m/s
Reuse: 7,170 m/s
Expend: 3.9
Reuse: 3.8
Expend: 19,000 kg
Reuse 9,600 kg
Expend: no ship
Resuse: 8,200 kg
Length18 to 20 m
Width4.6 m

The name of the spaceship is LOX-augmented Nuclear Thermal Rocket Lunar Transfer Vehicle, mercifully abbreviated to LANTR LTV.

This is from Human Lunar Mission Capabilities Using SSTO, ISRU and LOX-Augmented NTR Technologies: A Preliminary Assessment (1995). Stanley Borowski of NASA Lewis Research Center was looking for a way to economically send manned expeditions to Luna, that is, with something cheaper than a non-reusable Saturn V rocket.

First off, he outlined the parameters for a true reusuable single-stage-to-orbit booster, instead of that Rube Goldberg Space Shuttle contraption.

Secondly, a more powerful engine that puny chemical rockets was indicated for the spacecraft. A NERVA-like solid core nuclear thermal rocket would be nice. Unfortunately while their specific impulse was a vast improvement over chemical engines, scaling up the blasted things from the putt-putt NERVA prototype so they had halfway decent thrust levels was a problem. It was a pity, since it only needed high thrust at certain parts of the mission. For the rest of the mission it could get by with already achieved levels of thrust. It's too bad there wasn't any way to make the engine shift gears... waitaminute!

There is a way to shift gears, the old LANTR trick! Just inject some supersonic oxygen into the exhaust nozzle like an afterburner and you could increase the thrust by up to 440%. That is good enough, and sure is easier than designing a monstrously huge reactor. Of course this degrades the specific impulse by a drastic 45% but you can't get something for nothing. You only need the afterburner for small parts of the mission, the rest of the time you can have normal specific impulse.

As it turns out, while Borowski didn't actually invent the LANTR concept, he helped develop it and has promoted it for lunar applications.

Thirdly there was that perennial problem of The Tyranny of the Rocket Equation. You can't have a reasonably sized payload as long as you are lugging along all your propellant. This looks like a job for In-situ Resource Utilization. That always gives the Tyranny a swift kick in the gonads.

Luna has a scarcity of hydrogen, but it has oxygen coming out of its ears. Which is just what a LANTR needs. As it turns out, there was a 1993 study looking into this, called LUNOX.

There had been earlier grandiose plans for huge lunar bases with titanic mining and refining installations to exploit Lunar resources. But LUNOX was trying to do this on the cheap, on a more modest scale. Read: on a scale that would NOT give NASA's funders in Congress an acute case of sticker-shock.

An initial lander delivers an oxygen production plant, storage tanks, and a nuclear reactor. A second lander delivers six remote-controlled tractors. Two of them are "loaders", which operators on Terra use to scoop up ilmenite-rich lunar soil and deliver it to the oxygen production plant. They figure that one plant could produce about 24 metric tons of LOX per year, which is enough for three manned missions.

Before a LUNOX site had been established, the project would have make do with ordinary NTR using no LANTR. Due to the state of the art, these would be use-once-and-throw-away spacecraft. But once LUNOX was up and running they could switch to reusable LANTRs and enjoy a much more economical trip to Luna. Using LUNOX would cut the mass to be boosted into LEO in half! And a resuable LANTR LTV can perform up to 20 missions before its nuclear fuel rods become spent.

Expendable LANTR LTV would transport as payload a Lunar landing / Earth return vehicle (LERV) with a crew module. After the LANTR LTV was disposed of, the LERV would land the astronauts on Luna, and after the mission was over transport the crew back into Lunar orbit and send them on their way back home to Terra.

Resuable LANTR LTV would assume there were a supply of LERVs at the lunar site. These would have their name changed to Lunar Landing Vehicles (LLV) since they would not be used as Earth Return Vehicles. Instead, the LANTR LTV would just transport the naked crew modules and surface payload modules. The LLVs would ferry the crew and surface payload from orbit down to the surface, and later ferry crew modules and a tank of refueling LUNOX back up to the orbiting LANTR LTV. One of the surface payload module types would be liquid hydrogen from Terra. Remember that Luna has vast supplies of oxygen, but precious little hydrogen. The LLVs need the stuff.

The expendable LANTR LTV would operate their engines at an oxygen-to-fuel (O/H) mixture ratio (MR) of 4.0, while the resuables would use an MR of 6.0

Design Comparison
Thrust66,700 N66,700 NBRAVO66,700 NCHARLIE
Exhaust Vel9,230 m/s9,230 m/sBRAVO9,230 m/sCHARLIE
MR Thrustn/a221,000 NBRAVO273,000 NCHARLIE
MR Exhaust Veln/a5,960 m/sBRAVO5,350 m/sCHARLIE
# Missions11BRAVO20CHARLIE
INERT MASS7,000 kg8,300 kgBRAVO8,300 kgCHARLIE
Payload20,000 kg
18,500 kg
15,500 kg
8,800 kg
(crew mod
surface pl)
12,000 kg
(surface pl
Terran LH2)
DRY MASS27,000 kg26,800 kg23,800 kg17,700 kg20,300 kg
LH2 Propellant13,000 kg6,500 kgBRAVO6,500 kgCHARLIE
LOX Propellantn/a24,900 kgBRAVO24,900 kgCHARLIE
Refuel LUNOXn/an/aBRAVO+17,100 kg+9,700 kg
RCS Propellant300 kgBRAVO300 kgCHARLIE
Total Propellant13,000 kg31,700 kgBRAVO31,700 kgCHARLIE
WET MASS40,000 kg58,500 kg55,500 kg49,400 kg52,000 kg
Mass Ratio1.482.182.332.792.56
ΔV LUNOX3,6000 m/s7,190 m/s7,810 m/s9,470 m/s8,680 m/s
MR ΔV LUNOXn/a4,640 m/s5,040 m/s5,490 m/s5,030 m/s

Design Bravo and Design Delta have almost identical designs (Delta entries that say "BRAVO" are identical to Bravo entries). Design Charlie is one design with two columns, for two different payload mixes (2nd column Charlie entries that say "CHARLIE" are identical to 1st column Charlie entries).

ΔV LUNOX means delta-V without using LANTR afterburner, and not taking into account refueling at Luna with LUNOX lunar oxygen. MR ΔV LUNOX means delta-V with LANTR, but still not taking into account LUNOX.

DESIGN ALFA (expendable)

This is a pathetic Lunar Transfer Vehicle using only a putt-putt NERVA engine with no LANTR afterburner. It is presented for comparison purposes, so the LANTR designs can point at it and laugh. It was designed to have its components boosted into orbit by a conventional Space Shuttle or a Titan IV rocket.

Alfa has an inadequate delta V of 3,6000 m/s. This means it does not have enough ΔV to do a Lunar Orbit Insertion burn. Instead, unlike the other designs, the poor LERV has to separate and do the burn itself. The extra propellant required really cuts into the LERV payload mass. Alfa has a higher listed payload mass than the other designs, but more of it is LERV fuel and less of it is LERV hardware and payload.

DESIGN BRAVO (expendable)

This was designed to be boosted into orbit by a hypothetical new single-stage-to-orbit (SSTO) rocket with a 9.2 m (30 foot) cargo bay. So the design is split into three 9.2 m long parts (9 m propulsion module, 9 m propellant module, and the LERV).

Unfortunately the only way to make everything fit was to put the propellant module liquid oxygen tank inside the liquid hydrogen tank. This is a bad idea. In September of of 2016 a SpaceX Falcon 9 rocket blew up on the launch pad because a liquid hydrogen tank immersed in a liquid oxygen tank froze the oxygen into solid oxygen. Design Bravo has extra insulation around the oxygen tank (incidentally cutting into the payload mass) but it is still a matter of concern.

Also in a desperate attempt to make everything fit into the booster, they had to use a methane (CH4) fueled LERV instead of the more efficient hydrogen (LH2) fueled LERV (methane tanks are smaller because methane is more dense).

DESIGN DELTA (expendable)

This is basically Design Charlie with the assumption that NASA can get Congress to approve funding for a more spacious SSTO booster with a 13.7 m (45 foot) cargo bay. This allows relocating the liquid oxygen tank outside of the liquid hydrogen tank, so it is no longer a ticking time bomb. It also allows using the more efficient LH2 LERV.


There are two columns for Design Charlie, but they are the same spacecraft with two different payloads. It assumes that there is a supply of reusable LLVs on site at the LUNOX base to ferry crew and cargo back and forth (delivered by prior expendable missions). So Charlie just carries the naked payloads, it does not lug along the LLVs as well.

The "crew delivery" payload has 6.8 metric tons of crew and crew module, plus 2 metric tons of surface supplies for the LUNOX base. The "cargo delivery" payload is just 12 metric tons of surface supplies. Among the surface supplies are liquid hydrogen fuel for the LLVs and other equipment at LUNOX base. Remember the Lunar soil is jam-packed with oxygen but hydrogen is very hard to come by.

The design shown is based on Design Bravo, with the oxygen tank ticking time bomb.

Since the LLVs at LUNOX base can also refuel Charlie with oxygen, Charlie is reusable. It can perform 20 missions before the nuclear fuel runs out. The other designs are disposed of, criminally wasting 95% of their costly nuclear fuel rods. Actually even Charlie is wasting 85% of its fuel rods, but NASA figures attempting to remove the rods for reprocessing is just begging for a nuclear disaster in space.

Charlie just runs its engines at 6.0 MR for the entire mission, to stay within the LH2 and LOX propellant limits.


Normal Growth LCOTV
PropulsionIon Drive
Specific Impulse8,000 s
Exhaust Velocity78,480 m/s
Input power
per Engine
46 kW @
beam voltage
Engine Efficiency82%
Engine Life6,000 hours @
beam current
16 amps
per Engine
20 kg
Thrust per Engine0.7 N
Number of Engines206
Total Thrust145 N
Thrust to Weight5×10-5
Trip Time180 days
Payload Bay227 metric tons
at 100 kg/m3
Power PlantSolar Cell
Solar Cell Area54,416 m2
Structural Mass4,057 kg
Power Plant Mass26,831 kg
System Mass
11,671 kg
System Mass
2,217 kg
Thermal Control377 kg
Avionics520 kg
Growth Margin9,339 kg
Inert Mass55,012 kg
Payload Mass227,000 kg
Dry Mass282,012 kg
Propellant Mass29,744 kg
Propellant Reserves892 kg
Wet Mass312,648 kg
Mass Ratio1.11
ΔV8,190 m/s
Accelerated Technology LCOTV
PropulsionIon Drive
Specific Impulse8,000 s
Exhaust Velocity78,480 m/s
Thrust per Engine5.2 N
Number of Engines26
Total Thrust135 N
Thrust to Weight4.76×10-5 g's
Payload Bay227 metric tons
at 100 kg/m3
Power PlantSolar Cell
Solar Cell Area41,495 m2
Structural Mass2,880 kg
Power Plant Mass22,212 kg
System Mass
1,979 kg
System Mass
2,005 kg
Thermal Control68 kg
Avionics520 kg
Growth Margin5,075 kg
Inert Mass34,739 kg
Payload Mass227,000 kg
Dry Mass261,739 kg
Propellant Mass26,901 kg
Propellant Reserves784 kg
Wet Mass289,424 kg
Mass Ratio1.11
ΔV8,190 m/s

This is from Technology Requirements For Future Earth-To-Geosynchronous Orbit Transport Systems (1979). The unmanned Large Cargo Orbital Transfer Vehicle (LCOTV) transports cargo from Low Earth Orbit to Geosynchronous Orbit. The report also described a chemically powered LEO to GEO transport for priority cargo.

The report was trying to put a price tag on an transport system capable of handling the construction of a large solar power station.

According to a later report, for various reasons, the report concluded the project would require a fleet of 13 LCOTVs. Not all of the fleet would be on line, some would be undergoing maintenance. The ships in the fleet would be allocated such that the SPS project would receive a total of 56 flights, transport a total payload of 29,860 metric tons, make from 1 to 13 flights a year, and transport from 33 to 2010 metric tons per year.

Two designs were considered. The "Normal Growth" design used conservative extrapolations of the state-of-the-art, the "Accelerated Technology" assumed additional money was invested to increase the state of the art.

Lewis Research Center GCNR

This is from Mission performance potential of regeneratively cooled gas core nuclear rockets (1971)

The report notes that while solid-core nuclear thermal rockets have twice the specific impulse of chemical rockets, this isn't enough of an increase for high-energy trajectories with very high payloads. On the other end of the spectrum, ion drives have superb specfic impulse, but their pathetic thrust lead to undesirably long mission times.

The report decided to stop playing around and look at a rocket engine that is more in the middle between solid NTR and ion drives in both specific impulse and thrust: open-cycle gas-core nuclear thermal rocket. Gas-core has about four to twelve times as much specific impulse as solid-core, about sixty times as much thrust, and a hideously deadly radioactive exhaust plume. About as radioactive as if a solid-core rocket had a total nuclear meltdown at a rate of one solid-core rocket per second.

The report ignores the radiation, figuring it is not their department to worry about it. They wanted to analyze the performance of the gas-core NTR to find the circumstances where the performance was so superior that it warranted a more detailed study (including how to deal with the lethal exhaust).

They didn't bother analyzing the safer closed-cycle NTR because it was a typical compromise that wound up with the disadvantages of both and the advantages of neither. For one thing the specific impulse was half that of an open-cycle engine.

Spoiler Alert: the open-cycle gas-core nuclear rocket is so superior to the solid-core that it isn't even a contest. The real eye-popping gains show up at 3,000 seconds of specific impulse, but even 1,500 seconds is impressive. It is worth it to develop the GCNR further.

Their baseline was a pretty standard regeneratively-cooled open-cycle gas-core engine. They assume a specific impulse in the range of 1,000 to 3,000 seconds (exhaust velocity of about 10,000 to 30,000 m/s) which is about seven times that of chemical rockets and about four times that of solid-core NTRs.

But unlike prior reports, they looked into optimizing the thrust levels for a given mission.

Engine Mass

The mass of the gas-core NTR engine is obviously the sum of its parts:

Me = Mmod + Mps + Mtp + Mn

Me: mass of engine (kg)
Mmod: mass of nuclear moderator (kg)
Mps: mass of pressure shell (kg)
Mtp: mass of turbopump (kg)
Mn: mass of nozzle (kg)

They assumed the following relationships:

MH2: hydrogen propellant flow rate (kg/sec) Note the "M" should have a dot over it (because m-dot means mass flow). Unfortunately many computers do not have unicode fonts capable of rendering the character "Ṁ".
F: thrust (N)
Isp: specific impulse (sec)
g: standard value of gravity acceleration (m/sec2) = 9.80665 m/sec2
Isp × g: exhaust velocity (m/s)

MU = MH2 × ( 1 / H2/U)

MU: uranium fuel flow rate (kg/sec)
MH2: hydrogen propellant flow rate (kg/sec)
H2/U: hydrogen-to-uranium flow-rate ratio = 100
1 / H2/U: reciprocal hydrogen-to-uranium flow-rate ratio = 0.01

VU: volume of uranium in core (m3)
Vc: volume of core (m3)
MH2: hydrogen propellant flow rate (kg/sec)
MU: uranium fuel flow rate (kg/sec)

Moderator Mass

Moderator is assumed to be 0.762 meters thick.

Mmod: mass of nuclear moderator (kg)
D: outer diameter of the core/moderator (m) = for this study 3.66 m
ρmod: density of moderator material (kg/m3) = 1,150 kg/m3

Pressure Shell Mass

P: pressure (atm)
Mcr: critical mass in reactor (kg) = 48 kg
F: thrust (N)
Isp: specific impulse (sec)
D: outer diameter of the core/moderator (m) = for this study 3.66 m
MU: uranium fuel flow rate (kg/sec)
MH2: hydrogen propellant flow rate (kg/sec)

t: thickness of pressure shell (m)
P: pressure (atm)
D: outer diameter of the core/moderator (m) = for this study 3.66 m
σ: allowable stress in pressure shell (atm) = 13,600 atm

Mps: mass of pressure shell (kg)
D: outer diameter of the core/moderator (m) = for this study 3.66 m
t: thickness of pressure shell (m)
ρps: density of pressure-shell material (kg/m3) = 8,000 kg/m3

Turbopump Mass

Mtp: mass of turbopump (kg)
MH2: hydrogen propellant flow rate (kg/sec)
P: pressure (atm)
ρH2: density of hydrogen (kg/m3) = 72 kg/m3

Nozzle Mass

Mn: mass of nozzle (kg)
ε: area ratio of nozzle = 300
F: thrust (N)
P: pressure (atm)

Hydrogen Propellant Temperature

This does not help calculate the mass of the engine, but it is needed with other parts of the design.

T: propellant temperature (°C? K?)
P: pressure (atm)
F: thrust (N)
Isp: specific impulse (sec)
D: outer diameter of the core/moderator (m) = for this study 3.66 m
MU: uranium fuel flow rate (kg/sec)
MH2: hydrogen propellant flow rate (kg/sec)


Note that the thrust level has been optimized for each mission.

D: outer diameter of the core/moderator = for this study 3.66 m
Mcr: critical mass in reactor = 48 kg
H2/U: hydrogen-to-uranium flow-rate ratio = 100
σ: allowable stress in pressure shell (atm) = 13,600 atm
ρps: density of pressure-shell material (kg/m3) = 8,000 kg/m3
ρH2: density of hydrogen (kg/m3) = 72 kg/m3
ε: area ratio of nozzle = 300
ρmod: density of moderator material (kg/m3) = 1,150 kg/m3

Mission delta-Vs

The report took the simplistic ideal mission delta-Vs and made a function to account for gravity losses, since in the real world rocket impulse burns are not instantaneous (indeed, with ion drive a burn can take weeks). The function figured in acceleration levels, parking orbit eccentricity, and final hyperbolic excess velocity.

This has to be done iteratively, since once you calculate the real delta-V, you optimize the thrust level to the real delta-V, which means you have to recalculate the real delta-V, which means you have to reoptimize the thrust level… You keep iterating until the function converges on a value.

(Mp)i: propellant mass of ith maneuver (kg)
(M0)i: mass at beginning of ith maneuver (kg)
ex: antilog base e or inverse of natural logarithm of x
ΔV: delta-V (km/sec)
Isp: specific impulse (sec)
g: standard value of gravity acceleration (m/sec2) = 9.80665 m/sec2
Ispg: exhaust velocity (m/s)

(M0)i+1: mass at beginning of next maneuver after ith maneuver (kg)
(M0)i: mass at beginning of ith maneuver (kg)
(Mp)i: propellant mass of ith maneuver (kg)
(Mpstr)i: propellant-structure mass of ith maneuver (kg)
Mjettison: mass jettisoned, such as Mars Lander (kg)
(Mis)i: interstage structure mass of ith maneuver (kg)

(M0)imax: mass at beginning of last maneuver (kg)
(Mp)imax: propellant mass of last maneuver (kg)
Mpstr: propellant-structure mass (kg)
Mts: thrust-structure mass (kg)
Mpay: payload mass (kg)
Me: engine mass (kg)


Mts: thrust-structure mass (kg)
F: thrust (N)


(Mpstr)i: propellant-structure mass of ith maneuver (kg)
(Mp)i: propellant mass of ith maneuver (kg)


(Mis)i: interstage structure mass of ith maneuver (kg)
(M0)i: mass at beginning of ith maneuver (kg)


Me: engine mass OF A SOLID-CORE NUCLEAR ROCKET as an approximation. For GCNR use this equation. (kg)
F: thrust (N)

For Mars and Jupiter missions Mjettison = 136,100 kg

Uncrewed Lunar Ferry Mission

In the uncrewed lunar ferry mission the GCNR starts parked in LEO, does a Hohmann transfer to deliver various amounts of payload (Mjettison) to be placed into a lunar orbit. It still has 50,000 kg of Terra return payload. It then does a Hohmann back into LEO parking orbit. After refueling it is ready to deliver another payload.

The graph in figure 5 compares the performance of the gas core rocket with a 825 sec Isp solid-core Nerva-II, and a theoretical maximum 825 Isp Nerva-II (meaning it has zero engine mass). A specific impulse of 2,000 seconds was chosen for the GCNR. The GCNR leaves both NERVAs in the dust. It needs much less initial mass in LEO for a given payload. Actualy for the 50,000 kg payload to Luna and 50K kg back (blue line in graph) the GCNR can carry close to twice the payload (red line in graph).

The graph in figure 6 shows fuel/propellant consumption, where the GCNR shames the NERVAs even more. With a 50,000 kg payload the GCNR uses about half the the propellant needed by the NERVAs.

Figure 7's graph shows, among other things, that raising the specific impulse from 2,000 seconds to 3,000 seconds only reduces the initial mass in LEO (IMLEO) requirement by 11%. Simply because the propellant requirement goes down with higher specific impulse.

Figure 8a shows how the moderator mass varies with the core diameter. You probably should use the "best guess" curve for required critical mass.

Figure 8b shows that changing the hydrogen-propellant to uranium-fuel ratio (H2/U) doesn't change the IMLEO very much. The report uses H2/U = 100. The IMLEO goes up as H2/U rises because the pressure shell weight increases (see equation), of course as H2/U rises the amount of expensive uranium fuel needed goes down.

Uncrewed Slingshot Mission

In the uncrewed slingshot mission the GCNR starts parked in LEO, the spacecraft accelerates into some amount of hyperbolic excess velocity (V), releases a payload of mass Mjettison, waits until it comes back to Terra, and then burns like heck to circularize into LEO. It still has 50,000 kg of Terra return payload. It can be reused after being refueled.

Figure 9 shows how various factors affect the IMLEO. Mass increases with increase in amount of V, mass decreases with increase of specific impulse. At V=0 if you double the GCNR Isp from 1,500 to 3,000 seconds you will save 20% on IMLEO. At V=5 km/s doubling Isp will save 30% on IMLEO.

The GCNR can point its finger and laugh at the optimistic solid-core NTR. SCRN cannot reach V=5.5 km/s at all, not without staging at any rate. Which coincidentally is the delta-V requirement for a 300 day Mars round trip. A single-stage 1,500 Isp GCNR can manage it. A 3,000 Isp GCNR can do that without even working up a sweat.

Crewed Interplanetary Missions

For crewed interplanetary missions, it assumed that 136,100 kg (Mjettison) of payload is delivered into a 0.9-eccentricity planetary parking ellipse with a periapsis at 1.1 planet radii. An additional 90,700 kg is the command module / habitat module plus the Terra reentry vehicle. The GCNR does not brake into orbit to be reused, it goes streaking by Terra into a far Solar orbit. The crew bails out in the reentry vehicle that aerobrakes to the surface. The analysis did not set a limit on entry velocity, though in reality if it is much over 11 km/sec it is difficult to make a reentry vehicle that won't melt into a fiery blob of molten impure aluminum.

Figure 10 below shows the results for the Mars round-trip mission.

The analysis simplified things by assuming circular coplanar orbits, so this is more of the average performance regardless of synodic period.

In figure 10a, if you fix the IMLEO at 106 kg for all cases, 3,000 sec Isp GCNR can do the mission in 225 days, 1,500 sec GCNR needs 360 days, and the weakling 825 sec SCNR needs a whopping 430 days.

If you fix the trip time at 200 days, the 3,000 sec GCNR needs only one-fifth the IMLEO of the 1,500 sec GCNR. The IMLEOs are all higher than 106 kg, but it might be worth it if you are trying to spare the crew from excessive radiation doses.

Figure 10b shows among other things that the optimum thrust is in the range of 400,00 to 5 million Newtons (4×105 to 50×105 N).

Figure 11 below shows the results for the Jupiter mission.

The possible missions are the circles, ignore the lines connecting them (those are just to label circles with their specific impulse). A mission can be performed in 580 days, or 1,020 days or 1,420 days but not at any other intermediate time values. The circle values are when Jupiter and Terra are in opposition or conjuction at mid-say. At other times the delta-V becomes impossibly high. If you really want the scary mathematical details they can be found in Approximate trajectory data for missions to the major planets.

Because of the circle limits, for the Jupiter mission raising the specific impulse can only reduce the IMLEO. For the Martian mission raising the specific impulse can both reduce the IMLEO and the trip time. But the GCNR can perform the mission faster than the SCNR because the latter is too weak to have access to the columns of circles with lower trip times. It can only do the 1,420 day trip.

In the 1,420 day column, the 3,000 sec Isp GCNR requires only 15% the IMLEO of the SCNR, and 60% of the 1,500 sec GCNR.

Figure 11b shows among other things that the optimum thrust is in the range of 106 to 2×107 Newtons, about twice that of the Mars mission.

Lewis Research Center Ion Rocket

Lewis Ion Rocket
Propulsionion drive
Exhaust velocity20,000 to
100,000 m/s
Return payload
(cabin, etc.)
23,000 kg
Crew supplies
(4.5 kg/day/man)
18,000 kg
Exploration Rocket
(Mars lander)
18,000 kg
Powerplant41,000 kg
Propellant104,000 kg
Wet Mass204,000 kg
Dry Mass100,000 kg
Mass Ratio2.04
ΔV14,000 to
71,000 m/s
Mars mission
500 days
Length120 m

This is from a 1965 study by the Lewis Research Center entitled Space Flight Beyond The Moon.

This is a standard ion rocket powered by a nuclear reactor. The reactor is at the nose, behind a shadow shield and separated from the crew compartment by a 120 meter long boom in order to use distance as extra radiation shielding. The heat radiators are trimmed in order to not extend outside of the radiation shadow, their narrow aspect indicates the shadow shield is minimal in order to save on mass.

Higher ion drive exhaust velocities come at the expense of higher power requirements, which means higher power plant mass, which means lower acceleration. The report mentions that one can calculate the optimal exhaust velocity/power plant mass, but does not go into details. As a rule of thumb a rocket's acceleration should not be below 0.05 m/s2 (5 milligee) or it will take years to change orbits.

Some of the designs depict the crew cabin as two modules, which probably means they spin around the ship's spine for artificial gravity.

Lewis Research Center Mars Landing

This is from Manned Mars landing mission by means of high-thrust rockets (1966).

Yet another nuclear engine design that tries to squeeze out enough delta V so it can actually land astronauts on Mars, instead of attempting to do months of science in five minutes as they go whizzing by. Since they were designing with underpowered NERVA engines with only 850 seconds of specific impulse, they were forced into the irresponsibly cuckoo "nuclear staging" solution just like the Boeing IMIS. Except the Lewis design does not jettison five dangerously radioactive nuclear reactors like used beer cans. It jettisons seven.

And like the IMIS, it cannot afford a NERVA for braking into Terra orbit. They come in hot at full velocity, trusting to the miracle of aerobraking. The crew of seven enters a winged reentry craft and gets to dice with death, gambling that they will run out of delta V before they run out of ablative heat shield. If they lose, perhaps some kid will make a wish on the shooting star which is all that is left of their incinerated bodies.

The mission had a crew of seven. Depending upon the types of optimization, either four or all seven made excursions to the Martian surface. A minimum of three crew was needed for emergency operation of the spacecraft. Normal operation required six crew, operating in pairs. 1/3rd of the day was for spacecraft operational duty, 1/3 was for sleep, and 1/3 was off-duty (recreation, personal chores, scientific duties, and study).

The propulsion is by a series of NERVA style solid-core nuclear thermal rockets using liquid hydrogen as propellant. They have a specific impulse of 850 seconds and a thrust-to-weight ratio of 10. Well, actually T/W of 8 because the liquid hydrogen tanks need thermal protection from the engines or they will boil-off dry in no time. And T/W of 3 if you must have biological anti-radiation shields. The crew is protected from engine radiation by a combination of propellant tanks, separation distance, and command module storm cellar but no biological shields (but offhand it looks like that will not be enough shielding). Somewhat arbitrarily the study authors assumed each engine would have a minimum mass of 3,630 kg (7,260 kg with thermal and biological shields.) The initial acceleration of each engine with a full propellant tank is 0.2 gs.

There was a variant with chemical engines, but it was pretty pathetic.

The report has zillions of trajectories, each optimized for one factor or another. For mission with durations of 600 days and below, the total delta V requirement vary from 26,070 to 41,360 m/s.


As with most designs this is a combination of the spacecraft control room and the anti-radiation storm cellar. Because nobody wants to die horribly by manning an unshielded control room during a solar proton storm, and leaving the control room unmanned is a Really Bad Idea. During normal operations a crew of two or three occupies the command module. All seven can be contained during radiation events for short periods (1 day), four of the crew have to stand.

The command module has a mass of 4080 kg (9,000 pounds), not counting the radiation shielding. It has a volume of 12.7 cubic meters, of which 1.4 m3 are radiation sensitive operating equipment.

The radiation shielding is the chemical rocket fuel and oxidizer used by the Earth reentry vehicle. The fuel is pumped into the vehicle at the end of the mission.


The life support system is wrapped around the command module. The food and oxygen cycles are open, but the water cycle is closed (totally recycled). It assumes each crew consumes per day 1 kg of food, 1.8 kg of water, and 0.95 kg of oxygen (which is a little skimpy on the water). Plus each crew has 9 kg of recycled water per day for washing and other utility purposes.

Cabin air leakage is assumed to be about 0.68 kg per day, and complete air changes are made at three month intervals.

The total life support requirements for 7 crewmembers per day is 1608 kg plus 260 watts of power.


The living module has a mass of 4080 kg (9,000 pounds) not including radiation/meteor shielding and a volume of 156 cubic meters. The breathing mix pressure is 48 kPa (7 psi). To improve reliability the module is divided into two pressure independant units.

The walls hold about 29 kilograms per square meter of reentry vehicle chemical fuel as meteor and mild radiation protection. For full-blown radiation storms the crew retires to the storm cellar. The chemical fuel is held in multiple independant loops. Since the sun only heats the sun-side of the module, temperature is equalized by rotating the meteor shield or circulating the fuel. The desired level of temperature is maintained by adjusting the angle of the module with respect to Sol and by surface coatings.

The living module is connected to the Earth deceleration system by a long boom. The boom rotates fast enough to provide the living module with 0.3 g of artificial gravity (close to the Martian surface gravity of 0.376g). A smaller counter-rotating centrifuge balances the angular momentum. It provides up to 10 g's for the crew to exercise because it would be real nice for the crew to be able to walk when they return to Terra.

The scientific crew might spend their entire duty cycle (all day) in the living module. The control crew have to spend part of their duty cycle in the command module.


The spacecraft requires 7.5 kilowatts of power. A solar Rankine system was selected because nuclear power reactors have too much mass and are too radioactive to be repairable. The Rankine was assumed to have an alpha of 91 kilograms per kilowatt for a mass of 680 kg. 1,360 kg since they carry along a spare.

During the mission the generator dish does its best to shield the cryogentic hydrogen propellant tanks from the burning rays of the sun. Otherwise all the hydrogen will boil off.

Solar photovolatic panels were considered, but in 1966 they were not exactly "mature" technologies.


As previously mentioned the design cannot afford the additional NERVA engines and propellant to brake the spacecraft into LEO. So the designers took the cheap way out and used aerobraking.

The system has a mass of 24,490 kilograms, including the 16,780 kg of chemical fuel which is stored during the mission around the command and living modules as radiation shielding. It carries a payload of 1,090 kg (the "real mission payload") consisting of the crew, the Mars surface samples, and the data.

The system is a winged reentry vehicle with a thick ablative heat shield on its belly (much like NASA's Space Shuttle), a retro-rocket with the chemical fuel, and a meteor shield to protect everything during the long mission.

The winged reentry vehicle has a mass of 4,990 to 7,480 kg for atmospheric entry velocities of 7,930 to 19,810 m/s. It has 1.7 meters of unobstructed interior depth to accomodate the seated crew. Internal volume of 40 cubic meters. Leading edges swept 60°, small radius nose, and maximum attack angle of 23°. Maximum G load of 10 g's, entry corridor depth is 48 kilometers. The planned entry maneuver calls for 8 g deceleration at supercircular speeds and 4 g at subcircular speeds.


Each landing vehicle can carry two crew.

Since the landers use aerobraking most of their mass is the propellant needed to lift off from Mars and return to the orbiting spacecraft.

While the spacecraft is in Martian orbit, the landers separate and shed their meteor/thermal shields. The retrorocket fires to put the lander on entry trajectory, then is jettisoned so it does not obstruct the aerobraking heat shield. The shield burns its little heart out because the Martian atmosphere is like making love in an airlock exceedingly close to vacuum (1/100th of Terra's atmospheric density does not help much with braking). The landing/hover rocket extends from the top and burns at 2 gs of deceleration. It has enough fuel to hover and "translate" for only two minutes so the pilot has to pick the landing spot quickly. Hopefully the shock absorbers are up to the task of absorbing the landing impact.

The takeoff tanks are kept cool enough to avoid boil-off by foils in a vacuum jacket. Ordinarily you couldn't do this on a planet with an atmosphere, but as previously mentioned the Martian atmosphere is not that far from being a vacuum already. The fuel has a specfic impulse of 430 seconds, they suggest Diborane fuel with Oxygen difluoride oxidizer (B2H6 + OF2). Well, at least they didn't do a jackass maneuver like try to use FLOX for oxidizer.

The crew does as much science as they can possibly cram into 40 days. Life support and electrical power have 40 days worth of consumable. If you need a longer mission stay time, you'll have to replace the power and life support with a system as massive as the spacecraft's.

Just before takeoff the lander is stripped of every possible gram of excess mass, because it is not carrying much in the way of extra fuel. This includes ripping off the tank insulation, detaching the landing/hover rocket, ditto the remains of the heat shield and landing system. The lander takes off on a bare minimum delta V trajectory into orbit. Two more tiny boosts allow rendezvous with the spacecraft.

The takeoff section has a mass of 1,540 kg, including two crew plus 227 kg of Mars samples and data. If one lander fails, the other can carry all four crew but no Mars samples.


These are two unmanned landers on a one-way trip to land exploration equipment. Each lander has a mass of 1,360 kg and carries 1,810 kg of payload. The total mass for both loaded landers is 6,340 kg. The payload includes scientific equipment, land roving vehicles and their fuel.

Reference Source

     7. Anon: Manned Mars Exploration in the Unfavorable (1975-1985) Time Period. Vol. H. Summary. NASA CR-53911, 1964.
     8. Ehricke, K. A. : A Study of Early Manned Interplanetary Missions (Empire Follow- On). NASA CR-60375, 1964.
     9. Anon. : A Study of Manned Mars Exploration in the Unfavorable Time Period (1957- 1985). Vol. IH. NASA CR-53668, 1964.
     11. Widmer, Thomas F. : Application of Nuclear Rocket Propulsion to Manned Mars Spacecraft. Proc. AIAA and NASA Conf. on Eng. Problems of Manned Interplanetary Exploration, Palo Alto (Calif.), Sept. 30-Oct. 1, 1963, AIAA, pp. 85-101.
     12. Ragsac, R. V., et al. : Manned Interplanetary Missions. Follow-on Study of Final Report. Vol. 1. Summary. NASA CR-56762, 1964.
     14. Shapland, D. J. : Preliminary Design of a Mars-Mission Earth Reentry Module. NASA CR-56209, 1964.
     15. Dixon, Franklin P., and Neuman, Temple W. : Study of a Manned Mars Excursion Module. Vol. I of III - Pt. I. NASA CR-56182, 1963.

Lewis Research Center Mars Ref

This is from Nuclear thermal rocket workshop reference system Rover/NERVA (1991). The author is Dr. Stanley K. Borowski, who helped design several realistic spacecraft in this section of the website.

The point of this paper was to present a "reference design" to use to measure other design proposals presented at the conference. There was a Mars mission reference and a Lunar mission reference, both using solid-core nuclear thermal rocket propulsion based on NERVA technology.

The standard NERVA is basically liquid hydrogen heated in a nuclear reactor then emitted through a converging/diverging exhaust nozzle to create thrust. The hydrogen is compressed to high pressures by a turbopump. It is then "preheated" by cooling the nozzle, reflector, control rods, peripheral shield, and core support structure. Finally it is injected into the reactor.

One of the problems is where to get the energy to run the turbopump. There are two solutions: the "hot-bleed" cycle and the "full flow topping" or "expander" cycle.

In Hot-bleed, about 3% of the hot hydrogen exhaust emitted from the reactor is diverted (left green arrow) to run the turbopump. It is then either used for roll control or reintroduced into the exhaust nozzle. So it is called "hot-bleed" because it is bleeding off some of the hot stuff and using it to make the turbopumps spin.

In Full Flow Topping, all the preheated hydrogen is diverted (right green arrow) to run the turbopump. Then it is injected into the reactor to create thrust. Full Flow Topping has superior specific impulse compared to Hot-bleed, I presume it is much more difficult to engineer.

The Lewis Mars Reference mission was developed by Borowski in 1991, aimed at the Mars launch opportunity in the far-future year of 2016 (heh). It examined both old-school 1972 and more modern 1991 NERVA engines.

In 1969, Werner von Braun described a Mars mission where the spacecraft had triple NERVA engines, using a 640-day opposition class mission with an 80-day stay at Mars and inbound Venus swingby. The Lewis Mars Reference mission uses a spacecraft with a single NERVA, using a of 434-day opposition class mission with a 30-day stay at Mars and an inbound Venus swingby. The Lewis mission is much easier on spacecraft stress and astronaut exposure to galactic cosmic rays.

The Lewis mission came in two options. The "all propulsive" profile uses extra propellant so at the end of the mission the spacecraft can be braked into Terra orbit for future re-use. The other option is to forgoe the extra propellant, carry seven metric tons of Earth Crew Capture Vehicle (ECCV), and at end of mission have the crew use the ECCV to do a aerobraking landing while the abandoned spacecraft sails off into an eccentric heliocentric orbit, never to be used again. Though presumably in future decades the authorities would want to capture and properly dispose of derelict spacecraft with still radioactive engines littering the solar system.

The base assumptions and ground rules for designing the mission are as follows:





  • 34.94t MTV


  • 500 km x 24 hr2 (EARTH ARRIVAL)



  • 340 m/s (dia > 28.5°)
  • 100 m/s (dia < 28.5°)


  • 560 m/s



1 250 km x 33,852 km = 1 SOL ORBIT = 24.66 HOURS
2 500 km x 77,604 km = 24 HOUR ORBIT


Engine Mass3
Ext Shield Mass4
'90 GRAPHITE NERVA850334/758.004.519.4
'90 COMPOSITE NERVA925334/758.824.520.2
'90 CARBIDE NERVA1020334/759.314.520.7
'90 COMPOSITE PHOEBUS9251112/25021.769.037.65



     VARIES WITH TANK SETS: TMI (~ 13%), MOC (~ 15%), COMMON TEI/EOC (~ 16%)

The reference mission's optimization is focused on reducing the Initial Mass in Low Earth Orbit (IMLEO) {which can be thought of as the wet mass}. This is the reason the spacecraft and mission is built around using a single NERVA engine, since those things are heavy. The engine has a thrust of 334 kilo-newtons (75 klbf).

One engine instead of three means the thrust-to-weight (T/W) ratio goes way down. The spacecraft has a lower acceleration, which means it takes longer to escape Terra's gravity, which means the gravity losses become larger, which means more delta V is needed, which means more propellant is needed.

The way to avoid this vicious cycle is to use the magic of the Oberth Effect. By doing the Terra departure burn at perigee (at the point in the orbit when closest to Terra, more generally periapsis) you actually get some delta V for free (actually the extra delta V comes from the potential energy from the mass of the propellant expended).

In this case, you want to do three burns at periapsis to minimize gravity loss. Refer to the graph above. If the spacecraft has a thrust-to-weight ratio of 0.05, escaping Terra with a single burn at periapsis will cost you 1,500 m/s of gravity loss. But if you do the escape with three separate burns at periapsis, the gravity loss is only 350 m/s.

Naturally if you increase the engine thrust in such a way that the T/W ratio goes up, this will also lower the gravity loss penalty. This is tricky since higher thrust engines generally also have a higher mass. But in this case it is almost impossible since the optimization is focused on lowering IMLEO. You'd somehow have to increase the thrust while keeping the mass the same. Maybe by shifting gears.

The report points out if you swap the 334 kilo-newton NERVA engine for a honking monsterous 1,112 kilo-newton Phoebus engine the spacecraft could do a single burn escape with gravity loss of only 400 m/s (the T/W ratio rises to 0.15). The price is the IMLEO rises from 615 metric tons to 750 metric tons.

Boeing Ref
Lewis Ref
EARTH DEPARTURE2/25/20162/25/20163/15/2016
MARS ARRIVAL7/31/20167/31/20168/19/2016
MARS DEPARTURE8/31/20168/31/20169/19/2016
VENUS FLYBY3/10/20173/10/20173/16/2017
EARTH ARRIVAL5/04/20175/04/20175/23/2017
EARTH DEPARTURE C3 (KM2/SEC2)10.3410.3414.07
MARS ARRIVAL VH (KM/SEC)6.826.825.31


The table above compares the IMLEO spacecraft masses of two older Mars reference missions, and the Lewis all-propulsive (resuable) reference mission with a NERVA operating at a specific impulse of 925 seconds. The impressive part is how the optimized trajectory of the Lewis mission saves about 150 metric tons of IMLEO.

On the left is the Lewis all-propulsive optimized reference ship. On the right is the NASA reference ship. The differences are in the sizes of the various propellant tanks, and the IMLEO. The Trans-Mars Injection Drop Tanks are limited by the payload shroud dimesions of anticipated heavy launch vehicles. The report assumes the limit is 10 meters in diameter by 30 meters in length.

For the four NERVA engine types, the initial mass in low Earth orbit (IMLEO) and the total engine burn time was calculated for the mission. The carbide core has the lowest mass, but the 1,112 kN composite core has the shortest burn time.

The 2016 propulsion-optimized 434-day mission was assumed, along with engine thrust of 334 kilo-newtons (75 klbf) or 1,112 kN (250 klbf), 1000 psia chamber pressure, 500-to-1 nozzle expansion ratio, 3 perigee burn Terra departure (1 burn for 1,112 kN composite), and in reuse mode (i.e., spending extra delta V to avoid discarding the ship).

NERVA Engines
EngineTempIspIMLEOBurn time
GRAPHITE CORE2,350 K850 s725 mt202.8 min
(334 kN)
2,700 K925 s613 mt179.4 min
(1,112 kN)
2,700 K925 s750 mt65.3 min
CARBIDE CORE3,100 K1,020 s518 mt158.4 min
Burn Durations
334 kN
(75 klbf)
1,112 kN
(250 klbf)
122.1 min104 min87.8 min38.2
(# perigee burns)
MOC40.0 min36.8 min33.8 min13.4 min
TEI30.0 min28.0 min26.1 min11.0 min
EOC7.1 min6.9 min6.7 min2.7 min
TOTAL199.2 min
(202.8 min)
175.7 min
(179.4 min)
154.4 min
(158.4 min)
65.3 min

TMI = Trans Mars Injection, MOC = Mars Orbital Capture, TEI = Trans Earth Injection, EOC = Earth Orbital Capture

The report looked at other missions across the synodic period. The chart below assumes the spacecraft uses the 334 kN (75 klbf) composite engine with an Isp of 925 seconds. The chart shows how increasing the initial mass in low Earth orbit (IMLEO) shortens the trip time.

"All Prop" is the all-propulsive mission where extra delta-V is spent to capture the spacecraft into Terra orbit for reuse.

"ECCV" is the mission where the extra delta-V is NOT spent, the crew abandons the spacecraft in the Earth Crew Capture Vehicle (ECCV) and aerobrakes to Terra landing, but the spacecraft goes sailing off into the wild black yonder.

"Split-sprint" is where the cargo is sent in an unmanned spacecraft on a Hohmann conjunction-class trajectory, while the crew goes in a manned spacecraft on a faster high-energy opposition-class trajectory.

Not shown is the dangerous "Hohmann tanker/dual vehicle" mission. This is where the unmanned cargo ship also carries the manned spacecraft's return propellant. Which means if the manned ship arrives only to discover that all the return propellant has leaked out, the crew is doomed.

As you can see, the 2018 All Prop mission has an IMLEO of about 700 metric tons at a 434 day mission. If you wanted to decrease the mission time to 365 days (1 year) you'll have to almost double the IMLEO to about 1350 metric tons.

Luna from Destination Moon

PropulsionLiquid Core NTR
Specific Impulse1,050 s
Exhaust Velocity10,300 m/s
Wet Mass226,000 kg
Dry Mass45,000 kg
Mass Ratio5.0
ΔV16,600 m/s
Thrust11,000,000 N
Thrust Power56 gigawatts
Initial Acceleration5g
49 m/s2
Engine Mass9,000 kg
Structure Mass27,000 kg
Payload Mass9,000 kg
Propellant Mass181,000 kg
Dry Volume70 m3
Total Volume100 m3
Length46 m
Ladder Length25 m
Body Diameter5.6 m
Wing Span21 m

Inspired by a post by Retro Rockets I took a look at the classic spaceship Luna from the movie Destination Moon (1950). With Robert Heinlien as technical consultant, this movie was the most scientifically acurate one since Frau im Mond (1929). It held the throne for 18 years, until it was supplanted by the movie 2001: A Space Odyssey (1968).

For the specifications I used data from Spaceship Handbook and the Retro Rockets article. I then massaged the figures until they were internally consistent.

Spaceship Handbook calculated that a round trip mission to the surface of Luna would take about 16,480 m/s of delta V. So that's our performance limit for the mission. In addition, it will have to have a thrust-to-weight ratio greater than 1.0, since it has to lift off from Terra's surface. The movie specifies 5 gs, which translates to 11,000,000 newtons.

The movie specified that the reaction mass was water, not liquid hydrogen. While this does simplify the tankage, it does cut the exhaust velocity/specific impulse in half.

A solid-core nuclear thermal rocket engine is not going to be able to crank out enough delta V, not at the specifed mass-ratio it ain't. But the liquid-core Liquid Annular Reactor System (LARS) will do nicely. It can jet out liquid-hydrogen propellant at 20,000 m/s or better, so it can probably manage to hurl water at 10,300 m/s. That will give the Luna a delta-V of 16,600 m/s, just a tad larger than the required 16,480 m/s for the Lunar mission. More than enough, assuming you don't waste a lot of delta V during the landing.

The movie says the structural mass is 27 metric tons, which makes it 60% of dry mass. Nowadays NASA vessels typically have a structural mass of 21.7% of structual mass. 60% is a bit extravagant but believable with 1950's technology. If you made the structure NASA-light, you could add about 17 metric tons to the payload. The payload is the crew, equipment, life support, acceleration couches, and controls.

Lunar Transportation System

Lunar Transfer Vehicle
Module Core
Inert Mass8.1 t
Propellant Mass7.0 t
Crew Module
(incl. crew)
8.4 t
Height14.4 m
(incl. drop tanks
less cargo)
15.2 m
Lunar Transfer Vehicle
Drop Tanks
Isp481 s
Num Engines4
Thrust89,000 N
Total Thrust356,000 N
Inert Mass5.8 t
Propellant Mass129.8 t
Lunar Excursion Vehicle
Isp465 s
Num Engines4
Thrust89,000 N
Total Thrust356,000 N
Inert Mass5.8 t
Propellant Mass22.4 t
Crew Module
(incl. crew)
4.4 t
Cargo Mass
15 t
Cargo Mass
33 t
Height8.5 m
Body Width
(less cargo)
7.5 m
Landing Gear
11.3 m

This is from Report of the 90-Day Study on Human Exploration of the Moon and Mars (1989) and from

This transport system has two components: the Lunar Transfer Vehicle (LTV) and the Lunar Excursion Vehicle (LEV). The LTV transports crew and cargo between Low Earth Orbit (LEO) and Low Lunar Orbit (LLO), part of the cargo could be a LEV. The LEV transports crew and cargo between LLO and the lunar surface.

Lunar Transfer Vehicle

The LTV is a "one and one-half" stage design, with a reusable core surrounded by expendable propellant tanks. This reduces the propellant load by about 10% compared to a single-stage reusuable vehicle. The core contains the propulsion/avionics module, the main propellant tanks, the aerobraking shield, the crew module (if any), and other assorted subsystems.

The LTV and LEV are boosted into orbit in a single heavy-lift launch vehicle. The LTV will be boosted with the core fully fueled, but the LEV will only be partially fueled due to the payload limit of the launch vehicle. The four fully loaded drop tanks will be boosted into orbit by two subsequent heavy-lift vehicles. The crew and any cargo modules would be boosted by the space shuttle.

Some in orbit assembly will be required: adding drop tanks to LTV core, the eight peripheral aerobrake segments attached to the LTV aerobrake shield core, and the cargo modules added to the LEV.

The LTV does a trans-lunar injection burn, and jettisons two empty drop tanks. It brakes into LLO and drops the two remaining empty drop tanks. It then acts as a staging base in LLO for the LEV.

If there is already an empty LEV in LLO parking orbit waiting to be reused, the LTV loads it with propellant, consumables, and attaches new cargo modules.

When the LEV has performed its mission, the LTV does a trans-Earth burn using the core propellant tanks.

It circularizes itself into LEO using aerobraking instead of propellant, at a considerable savings in initial mass required in LEO at mission start. After each mission the aerobrake shield is refurbished and verified at the International Space Station. The aerobrake shield can be reused for five missions.

The optional LTV crew module provides habitable support for the crew for the 4 day translunar trip and up to 7 days for the return to the space station. Naturally the crew can override the automatic rendezvous and docking system. Crew module obtains electricity from the LTV, has a two-gas open-loop environmental control and life support system, has a galley, zero-gravity toilet, and a personnel hygiene station.

The crew module has docking ports fore and aft, passing through one is an intravehicular activity (no space suit required). There is no airlock, so extravehicular activity requires all the crew to don space suits and depressure the entire module. There is enough repressurization gas carried for 2 EVAs.

The crew module carries a storm cellar with walls filled with water radiation shielding. The water is vented before aerobraking to save wear and tear on the aerobrake shield.

Lunar Excursion Vehicle

In "reusable" mode, the LEV can transport 15 metric tons of payload to the lunar surface (along with a crew and crew module), and return to LLO. It can be reused up to five missions.

In "expendable" mode the LEV can transport 33 metric tons of payload to the lunar surface (with neither crew nor crew module) and stays on the surface forever after.

If the cargo load is small enough, an unmanned LEV with no crew module has enough of an automatic pilot to be able to land, discharge cargo, return to orbit, and rendezvous with the orbiting LTV.

The LEV and LTV shares a lot of systems designs to reduce development and testing time (such as engines, cryogenic RCS, avionics, software, communiciation equipment, fuel cells, etc.).

When the LEV is parked in lunar orbit and abandoned, it is powered by solar arrays. On the lunar surface, the propellant system is designed for 30 days. For longer stays it will require surface support (from in-situ resource utilization).

The LEV's crew module is related to the LTV crew module, but with some differences. It has no storm cellar. It transports four crew members between the LTV and the lunar surface. During landing operations two crew members have landing control panels and windows, the other two are in shock webing and just have to be patient and stare at the walls.

The LEV's crew module's systems are in a quiescent state, except for 4 days during descent/ascent missions (2 days during descent and initial surface operations, 2 days for preparation and ascent to orbit). While quiescent the crew module has no interal power, thermal control, or propellant conditioning. Bottom line is either the descent/ascent mission only lasts 4 days, or there has to be support systems available on the lunar surface (a lunar base in other words).

Just like the LTV crew module, the LEV crew module has no airlock and only enough repressurization gas for 2 EVAs.

Lighter and Tanker

Specific Impulse450 s
Exhaust Velocity4,410 m/s
Wet Mass56,300 kg
Dry Mass25,898 kg
Inert Mass898 kg
Payload25,000 kg
Mass Ratio2.17
ΔV3,410 m/s
Mass Flow31.8 kg/s
Thrust140 kiloNewtons
Initial Acceleration0.25 g
Length18.3 m+engine
Diameter≈4.57 m
gas core NTR
Specific Impulse3,600 s
Exhaust Velocity35,000 m/s
Wet Mass433,000 kg
Dry Mass268,000 kg
Mass Ratio1.61
ΔV16,730 m/s
Inert Mass
(dry mass - payload)
108,000 kg
Payload Mass total160,000 kg
Payload Mass Hydrogen
(less tankage)
139,000 kg
Mass Flow100 kg/s
Thrust3,500 kiloNewtons
Initial Acceleration0.6 g
Length37 m+engine
Diameter≈18 m

These two designs are from The Resources of the Solar System by Dr. R. C. Parkinson (Spaceflight, 17, p.124 (1975)). The Lighter ferries tanks of liquid hydrogen from an electrolyzing station on Callisto into orbit where waits the Tanker. Once the Tanker has a full load of tanks it transports them to LEO. All the ships are drones or robot controlled, there are no humans aboard. The paper makes a good case that shipping hydrogen from Callisto to LEO would eventually be more economically effective than shipping from the surface of Terra to LEO, with the break-even point occurring at 7.8 years. Please note that this study was done in 1975, before the Lunar polar ice was discovered, and probably before the ice of Deimos was suspected.

Warning: most of the figures in the table are my extrapolations from the scanty data in the report. Figures in yellow are sort of in the report. Use at your own risk.

The tanker uses a freaking open-cycle gas-core nuclear thermal rocket. This is an incredibly powerful true atomic rocket, but it is only fractionally more environmentally safe that an Orion nuclear bomb rocket. The report says it should be possible to design it so the amount of deadly fissioning uranium escaping out the exhaust is kept down to as low as one part per 350 of the propellant flow (about 300 grams per second), but I'll believe it when I see it. Since it is used only in deep space we can allow it, this time. The report gives it an exhaust velocity of 35,000 m/s, which is about midway to the theoretical maximum.

The lighter can get by with a more conventional hydrogen-oxygen chemical rocket. It will need an acceleration greater than Callisto's surface gravity of 1.235 m/s2, for safety make it 1.5x the surface gravity, or about 1.9 m/s (0.6g).

The four major Galilean moons are within Jupiter's lethal radiation belt, except for Callisto. The black monolith from 2010 The Year We Make Contact only told us puny humans to stay away from Europa, so Callisto is allowed. If you want ice that isn't radioactive, you've come to the right place. It is almost 50% ice, and remember this is a moon the size of planet Mercury. That's enough ice to supply propellant to the rest of the solar system for the next million years or so. Europa has more, but it is so deep in the radiation belt it glows blue. Callisto is also conveniently positioned for a gravitational sling shot maneuver around Jupiter to reduce the delta-V required for the return trip to Terra.

The report says that the requirements for an economically exploitable resource are:

  1. It is not available in the Terra-Luna system
  2. It must provide more of it than the mass originally required to be assembled in Terra orbit at the outset of the expedition
  3. It must be done within a reasonably short time (the break-even time)

Hydrogen fits [1], or at least it did until the Lunar ice was discovered. [2] and [3] depend upon the performance of the vehicle.

There are three parts. First is the Tanker, which is an orbit-to-orbit spacecraft to transport the hydrogen back to LEO and brings the expedition to Callisto in the first place. Next is an electrolysis plant capable of mining ice, melting it into water, cracking it into oxygen and hydrogen, and liquefying the hydrogen. Last is a Lighter which is an airless lander that ferries liquid hydrogen from the plant on Callisto to the orbiting Tanker.

The report decided to use modular cryogenic hydrogen tanks that would fit in the Space Shuttle's cargo bay. They would have to be about 18.3 meters x 4.57 meters, about 300 cubic meters capacity. The report has a filled tank massing at 26,000 kg, with 22,000 kg being liquid hydrogen and 4,000 kg being tank structural mass. Examining the drawing of the tanker, the front cluster is composed of four tanks while the rear has nine, for a total of thirteen. The tanker will have a length of two tanks plus the length of the rocket engine, 37 meters plus rocket. The rear has tanks arranged in a triangular array about four tanks high. So a diameter roughly 18 meters or so.

The lighter carries a single tank, so it is roughly one tank in diameter, and one tank long plus the fuel tanks+engine length. It will need a large enough liquid hydrogen/liquid oxygen chemical fuel capacity to lift off from Callisto to the tanker and land back on Callisto.

The report figures that the electrolysis plant can produce hydrogen for about 39 kW-h/kg, that is, each kilogram of hydrogen in the plant requires 39 kilowatt-hours. Figure it needs more electricity to liquefy the hydrogen, and more to produce the liquid oxygen needed by the lighter, for a total cost of 50 kW-hr/kg for liquid hydrogen delivered to the orbiting tanker. So a 2 megawatt nuclear reactor could produce 350 metric tons of hydrogen per year. Launch windows back to Terra occur every 398.9 days.

Once the lighter has made enough trip to fully load the tanker, the tanker departs for LEO. It will use some of the hydrogen for propellant, some will be the payload off-loaded at LEO, and enough will be left to return the tanker to Callisto. The amount of payload is specified to have a mass equal to 37% of the fully loaded mass of the tanker. It also specifies that the inert mass fraction of the tanker is 25% of the tankers fully loaded mass.

The report had an esoteric equation that calculated the mass of the lighter and electrolysis plant as a percentage of the tanker mass in order to be economically viable. It turns out to be 13% of the fully loaded mass of the tanker. When the expedition is launched the tanker will carry the lighter, the electrolysis plant, and enough propellant so that the total mass is 52.9% of the fully loaded mass (i.e., it departs half empty). The lighter will have its tanks full.

Five years later, upon arrival at Callisto, the lighter lands the electrolysis plant on a prime patch of ice. It then starts the cycle: patiently waiting for the plant to fill the payload tank and the fuel tanks, boost the payload to the tanker, then land back at the plant to start again.

In context: this was one of a series or articles I wrote for Spaceflight at the encouragement of the then editor, Ken Gatland, triggered off in the dark days following abandonment of the Apollo programme by a discussion at the BIS as to what would be needed to make spaceflight self-supporting. The first article was published in Spaceflight 1974 p.322 under the title "Take-Off Point for a Lunar Colony." There was then a second on "The Colonization of Space" (S/F 1975 p.88) and a couple of subsequent ones on Lunar Colonies (S/F 1977 p.42/103). Later, when they invented the first spreadsheets, I did some speculation on how the economics of everything might fit together economically in a big input-output model which got published as "The Space Economy of 2050 AD" in JBIS v.44, p.111 (1991) which also appeared in my book Citizens of the Sky (1989) later. It is unlikely that I was consistent through all of this — my opinions develop with time — and by the 1991 period I was heavily in to the economics or reusable launchers and what would happen if the models were pushed to very high flight rates.

Going back to "The Resources of the Solar System", I'm not sure how much detail I managed at the time. I remember that there were a couple of things influencing me at the time. One was the concept of a gas core nuclear engine (GCR) which might have a specific impulse of about 3500 sec (35 km/s). To really move around the Solar System you need a high thrust-to-mass engine with this sort of specific impulse, and GCR had the interesting property of using hydrogen as propellant. (Ion motors can meet the specific impulse, but to do a similar job would require a power-mass ratio several orders better than anything we could consider then or even today — VASIMIR suffers the same problem). Nowadays I might put my money more on a pulsed-fusion system (see “Using Daedalus for Local Transport,” JBIS, 62, p. 422-426 (2009)) — note NOT using helium-3, which would change the model significantly.

The second thing at the time that influenced me — at a time when the Space Shuttle was still a paper vehicle — was that the Space Shuttle payload bay was just about the right size (15 ft × 60 ft) to carry a full liquid hydrogen tank (there are reasons now why it wouldn't which led to the abandonment of design work using Centaur as an upper stage) — so my modular design was based around using that as a standard tank. For use in long duration space missions the tank would have to have some sort of active cooling system to keep the hydrogen from boiling away, but given that you could then ship LH2 around the Solar System on slow, economical trajectories like modern oil tankers on Earth. Once you have rerfuelling stations at either end interplanetary flight becomes a lot easier and you can think of using higher speed trajectories for special cargo like human beings.

From personal email from Dr. Parkinson (2014)

All the other figures in the table are ones I've extrapolated from the few figures given in the report.

A plausible figure for nuclear power generation is 0.12 Megawatts per ton of generator. This would make the electrolysis 2 MW power reactor have a mass of 16,000. This is close to the 25,000 kg mass of a payload tank. So to simplify, assume the electrolysis rig with liquefaction gear and all masses a total of 25,000. This also ensures that the lighter is capable of landing it.

The tanker's inert mass fraction is 25%, and hydrogen payload is 37%. This means the dry mass is 62%, which means the mass ratio is 1.61. With an exhaust velocity of 35,000 m/s, this yields a total delta-V of 16,730 m/s. I am unsure if this is enough for a Callisto orbit-LEO mission followed by a LEO-Callisto orbit mission. Not without a heck of a gravitational sling-shot it isn't. Or I could have made a mistake in math.

Note both the payload and the propellant is hydrogen, stored in the same array of tanks. If the inert mass fraction is 25%, then the payload+propellant mass fraction is 75%. If there are 13 tanks each of 25,000 kg, then the total is 325,000 kg. If this is 75% of the wet mass, the actual wet mass is 433,000 kg. If the payload is 37% of the wet mass, it is 160,000 kg. If a hydrogen tank is 87% hydrogen and 13% tankage, the amount of hydrogen payload is 139,000 kg.

On the initial trip, the tanker carries the electrolysis plant and the lighter (with no payload, but with full fuel tanks). This is 13% of the wet mass or 56,300 kg. If the electrolysis plant is 25,000 kg, the lighter (with no payload) must be 31,290 kg. The lighter payload is one payload tank at 25,000 kg. So the lighter wet mass is 56,290 kg.

The lighter needs a delta-V of 3,414 m/s (Callisto-surface-to-orbit + orbit-to-Callisto-surface). Chemical fuel has exhaust velocity of 4,410 m/s. This means the mass ratio has to be 2.17. This implies the dry mass is 25,898 kg. Subtract the 25,000 kg payload, and there is 898 kg for the structure and the engine. Seems a little flimsy to me, perhaps 25,000 kg is a bit to generous for the payload tank.

Tank is scaled to fit in Space Shuttle cargo bay. At least the the proposed size of the bay in 1975 when the report was written, it was later reduced in size.

  • 4.55 m wide × 18.2 m long
  • Mass 26 metric tons
  • LH2 Mass 22 metric tons

Notes, from left to right:

  1. Docking Ring
  2. Limited amount of pressurization equipment round head end, also radar transponder
  3. Strong ring with attachment points
  4. Recirculating and pressurization pipes
  5. Strong ring with attachment points
  6. Docking Ring
  7. Fill valve connect. Associated propellant management equipment
  8. Radiator

Manned Mars Explorer

Mass Schedule
98,000 kg
70,000 kg
Tether System13,000 kg
Power System20,700 kg
32,000 kg
Comm Satellite3,500 kg
Crew Command
50,000 kg
(Scenario 1)
3,412,800 kg
(Scenario 2)
3,812,800 kg
(Scenario 1)
3,700,000 kg
(Scenario 2)
4,100,000 kg

This is from a student project for Master of Architecture candidates at the University of Houston: Manned Mars Explorer project: Guidelines for a manned mission to the vicinity of Mars using Phobos as a staging outpost.

After weighting all the options, chemical propulsion was chosen. Nuclear electric had too many drawbacks.

The tyranny of the rocket equation led them to go with reliability over redundancy. Equipping the spacecraft with back-up units for all critical systems cuts too much into payload mass. Instead they went with single units that were super-duper fault tolerant.

Medical issues dictated supplying the crew with a full one-Terran-gravity. An elaborate bola system was designed. The system resists twisting via a unique spreader system and four tether configuration. Spin grav is used for the trans-Mars coast and the trans-Terra coast. The tether is reeled in before each propulsive manuever to prevent the spacecraft from destroying itself by the crack-the-whip effect.


An opposition class Venus inbound swingby was used for the trajectory. About 300 days are spent travelling to Mars. It spends 60 days in Mars orbit. It uses the Venus inbound swingby leg to travel to Terra LEO which takes 210 days. This trajectory was chosen due to relatively short overall mission and Mars stay time. It does however require more delta V than conjunction class trajectories.

The sixty day exploration period is mostly focused on Phobos and Deimos, but there is a segment where a crew of three is sent to the surface of Mars for seven days. Staging bases will be set up on the moons, and they will be assesed for deposits of water ice and other valuable in-situ resource utilization goodies.

MPV is the Manned Planetary Vehicle, the spacecraft. CCV is the Crew Command Vehicle, a small auxiliary spacecraft carried as payload.

Mission Phases:

  1. Low Earth Orbit construction
    1. Vehicle assembly
    2. Crew training
  2. Trans-Mars injection
    1. Propulsive maneuver
    2. Communication satellite deployment
    3. Spin-up
    4. Power system deployment
    5. Tether system deployment
    6. Trans-Mars coast
    7. De-spin
    8. Power system retrieval
    9. Tether system retrieval
    10. Communication satellite retrieval
  3. Mars circularization
    1. Propulsive maneuver
    2. CCV surface operations
    3. CCV return to MPV
  4. Trans-Earth injection
    1. Propulsive maneuver
    2. Communication satellite deployment
    3. Spin-up
    4. Power system deployment
    5. Tether system deployment
    6. Trans-Earth coast
    7. De-spin
    8. Power system retrieval
    9. Tether system retrieval
    10. Communication satellite retrieval
  5. Earth orbit capture
    1. Propulsion stage, CCV, and MPV separation
    2. Propulsion stage remains in hyperbolic orbit
    3. CCV propulsively circularizes at LEO with crew
    4. MPV aerobrakes into Space Operations Center orbit


The mission spacecraft is composed of the following components:

  • Pressurized Environment System: the habitat module
  • Power System: provides electricity
  • Structural System: the ship's spine
  • Folding Aerobrake System
  • Four-Tether System: provides 1 gee of spin gravity
  • Staged Propulsion System

It carries the following payload:

  • Crew Command Vehicle (CCV): transports explorers to Phobos, Deimos, and Martian Surface
  • Communication Satellite

Pressurized Environment System

This is the MPV's habitat module. It is composed of a habitation module, a laboratory module, a safe haven (storm cellar) and connecting tunnels. The large modules will have three airlock section, each with two means of egress: one to another pressurized airlock section, and the other to either the exterior or another airlock section. Space suits will be stored next to each exterior egress for use in a planned EVA or emergency escape to the CCV. The storm cellar will accomodate the 6 member crew for 12 days. The anti-radiation walls are 10 cm thick aluminum.

The mass budget for the system is 98,000 kg.

Power System

The power system mass is budgeted at 20,700 kg. It is specified to provide 150 kW constant power. It uses a set of solar dynamic power systems rated at 128 kg/kW. This system cannot be used during propulsive maneuvers or planetary eclipses. During these periods power is supplied by fuel cells.

Structural System

The structural system is rated to withstand forces of 3.5 g under propulsion and 1 g under spin gravity. The layout with pressurized environment allows optimal thrusting through the center of gravity. The system budget is 32,000 kg.

Folding Aerobrake System

The aerobrake system is used at the end of mission, to brake the MPV into the orbit of the space station. This saves a whopping 50% of propellant. The system assumes that the crew has already departed in the CCV, to remove the mass of the CCV and to spare the crew a fiery death in case aerobraking fails.

The system consists of the aerobrake (in two folding sections), a transferrable strutural pallet, folding mechanism, and fuel for Terra orbit circularization and rendezvous with the space station.

It is also used as a movable counterweight mass while in spin-grav mode.

If the aerobrake is non-functional, the MPV would enter a hyperbolic trajectory and enter a wild solar orbit.

The mass budget is 70,000 kg.

Four-Tether System

This is a bola type of spin gravity, providing 1 gee of gravity to the habitat module. Bolas are much more lightweight than centrifuges using girders or something. Much easier to boost into orbit as well. However it is prone to dangerous oscillations and perturbations. If the cables snap, the habitat module will be separated from the propulsion system, and fly into the big dark heading for a lonely doom for the unfortunate crew.

The students designed a four-tether system to deal with oscillations.

The rotation speed was limited to 2 rpm. For 1 gee his means the spin radius at the habitat module will have to be about 224 meters. The spin center will of course shift as propellant is burnt and the center of gravity changes.


While the students were designing they realized that aerobraking the MPV at Earth left little counterweight mass for the return rotation cycle. So they created two scenarios with different arrangements, and different mass budgets.

Scenario 1

Total delta V: 10,475 m/s

Trans-Terra injection propulsion stage containment mass is retained for counterweight mass on return leg rotation cycle.

Aerobrake is transferred from MPV to propulsion stage for counterweight mass on return leg rotation cycle.

Trans-Mars injection separated into two propulsive maneuvers.

Scenario 2

Total delta V: 12,599 m/s

Trans-Earth injection propulsion stage containment mass is retained for counterweight mass on return leg rotation cycle.

Aerobrake remains attached to MPV and is not transferred as in S1

Aerobraking maneuver has a propulsive assist = 610 m/sec which requires more fuel than S2

Trans-Mars injection is one propulsive maneuver.

Falure analysis showed that up to two tether could snap simultaneously without catastrophic failure. In this case the habitat module would be reattached to the propulsion system, and the crew would just have to endure zero gee for the rest of the mission.

The mass budget for the tether system is 13,000 kg.

Propulsion System

The student's analysis showed that a nuclear-electric propulsion system was unworkable, so they used a plain vanilla liquid hydrogen / liquid oxygen chemical rocket. Unsurprisingly the failure analysis revealed that the propulsion system fails anytime after trans-Martian injection the crew faces a death sentence. Even with a free return flyby, the ship will run out of consumables years before it return to Terra.


The CCV is a little auxiliary spacecraft carried as payload. The mission re-uses the heck out of it, to get their money's worth out of the mass it eats up. The crew occupies it during all propulsion burns made by the MPV. It ferries explorers to Phobos and Deimos. It lands a team of three explorers on the Martian surface. And at mission's end it transports the crew to the space station while the unmanned MPV aerobrakes into parking orbit.

The CCV is used by the crew when the MPV does burns because the habitat module does not have any acceleration couches (every gram counts!). Having said that, the couches can be rotated 180° between two configurations, since the direction of "down" is different between MPV burns and CCV landing on Mars. Without rotation, MPV burns would feel like the couches were attacked to the ceiling with the astronauts in a most uncomfortable "eyeballs-out" position.

CCV mass is budgeted at 50,000 kg.

Mars Base Camp

This is from MARS BASE CAMP UPDATES AND NEW CONCEPTS by Lockheed-Martin (2017).

This Mars base mission concept was released about one hour before SpaceX released their Mars mission concept. The Lockheed mission relies heavily upon NASA's Space Launch System (SLS). A person of suspicious mind would find the timing questionable. It would seem to be for the purposes of stealing SpaceX's thunder.

The background is that the SLS has suffered development delays, cost overruns, and criticism that it is not really needed so should be cancelled. On the other hand it provides lots of jobs in states controlled by powerful senators. Meanwhile the United Launch Alliance (ULA) {which just so happens to include Lockheed-Martin} had a monopoly on boosting USAF payloads even though their boost price kept rising. SpaceX filed a lawsuit which they won, and then proceeded to boost the USAF payloads at a mere 20% of ULA's price tag ($90 million as opposed to $460 million). ULA's VP of engineering made public comments that ULA was quote "resentful of SpaceX" unquote. He latter resigned.

The point being that Lockheed-Martin does not like SpaceX very much.

In a related development NASA made an announcement it was looking into a Deep Space Gateway in cis-Lunar orbit (EML-3). The interesting facts are that the Gateway's components are carefully sized so they cannot be boosted by SpaceX's rockets {only by the as-yet nonexistent NASA SLS}, and that the project has been criticized as having no purpose. Well, no purpose other than giving the SLS a reason to exist, that is. And trying to sabotage SpaceX.

But I digress.

The Mars Base Camp is a crewed vehicle established in Mars orbit. From it a crew of 6 astronauts can perform excursions to Deimos and Phobos, perform telerobotic exploration of the Martian surface (including sample returns), produce LH2/LOX fuel via solar-powered electrolysis from water (either delivered from Terra by unmanned Water Delivery Vehicles {WDV} or from ISRU ice from Martian Moons), and allow astronaut sorties to the surface via reusable Mars Ascent/Descent Vehicles (MADV).

Elements of the camp are pre-placed before the arrival of astronauts, such as the lab, the center node, two excursion modules, two MADV and one or more Mars-orbiting cryogenic fuel depots. These are transported by unmanned dual-mode stages. These have solar electric propulsion for long-period delta V, and chemical thrusters for RCS.

The crew transfer vehicle stack consists of a habitat module, two cryo-stage propulsion systems, and two Orion spacecraft (NASA's proposed Multi-Purpose Crew Vehicle, not the nuclear bomb powered kind. There are too many freaking spacecraft named "Orion").

The crew transfer vehicle carries 6 crew. After the mission is over the transfer vehicle is docked to the Deep Space Gateway while the crew returns to Terra by Orion reentry. The transfer vehicle can be serviced and reused.

The idea is NOT to perform a stupid "flags and footprints" stunt like Apollo, spending billions of dollars to let some clown walk on the moon and then nothing. The Mars Base Camp is intended to establish reusable infrastructure to sustain long-termed crewed operations at Mars. Each mission should lay the groundwork for the next, otherwise the intermittent funding of NASA can lead to large post-mission gaps. As of this writing the post-mission gap after the last Apollo lunar mission has grown to 45 years with no end in sight. In addition every piece of equipment should be reusable because US lawmakers frown on multi-million dollar pieces of gear that are used once then thrown away like high-tech toilet paper.

As much as possible existing technologies should be used, because the long development times for non-existing technologies is just begging to get the entire project axed by penny-pinching congress-critters.

It goes without saying that crew safety is paramount. Astronauts dying a low agonizing death in space will be a public-relations nightmare that NASA is unlikely to survive. No single point of failure is to be allowed, which means everything should be redundant. Two Orions, two crew quarters, two MADV landers, etc. The movie The Martian was nice science fiction, but unlikely to have a happy ending in the real world.

The report also specifically states the Mars Base Camp must be used to prove that the Deep Space Gateway is not a gigantic boondoggle, even if it is. About two pages of the entire report is devoted to listing various ways the DSG can be used to assist the development of the Mars Base Camp.

The MADV landers can each make multiple sorties to the Martian surface, provided there is enough orbital propellant depots to refuel them. The fuel is generated as needed from water by electrolysis. This is because liquid hydrogen and liquid oxygen has an annoying habit of boiling, with the need to waste fuel by venting the vapor to space or the freaking fuel tanks explode. The longer the liquid fuel sits in the tanks, the more you lose to boil-off. The water is supplied by unmanned water delivery vehicles, from NASA if need be but the report has the pious hope that commercial suppliers of water will spring into being. They can produce water by mining various in-situ sources (asteroids, Martian surface, but ideally from the moons of Mars) and deliver it to Mars Base Camp propellant depots. And give NASA the bill for their services.

In between manned mission the unmanned central part of the Mars Base Camp remains in Mars orbit.

  • Lab Module
  • Center Node
  • Deimos/Phobos Excursion Module
  • x2 solar electric propulsion stages and related solar panels

Since the crew transport uses Hohmann transfers, the crew will have to stay at Mars for one synodic period (11 months) before they can head for home. The report notes that due to the availability of the MADV landers, the crew spends much of that time in orbit, inside the Mars Base Camp instead of spending the entire time on the surface as in most all other proposed Mars missions. This lowers the mission cost since you do not have to land the entire crew habitat module and all the support equipment. It also eliminates the failure mode where the habitat module is pre-positioned but the crew lander accidentally planets at a distance from the hab mod which is too far to walk.

The Crew transfer vehicle carries enough fuel for the round trip to Mars and back, and enough surplus fuel for either [a] Two sortie missions to the Martian moons or [b] One sortie mission to the Martian surface. If any more sorties are desired, water will have to be delivered by water delivery vehicles and electrolyzed into fuel. Missions to the moons are performed with an Orion-Excursion System-Cryogenic Propulsion Stage stack split off from the crew transfer vehicle. Missions to the surface are performed by a Mars Ascent/Descent Vehicle. Surface sorties will transport 4 crew to the surface while 2 crew remain in orbit in the base camp.

The Crew transfer vehicle is designed to have zero boil-off, but the MADV and WDV are not.


Water delivery vehicles can theoretically be of any size, but the report assumes they carry 52 metric tons of water. The report calls this a 50 MT Class WDV. Two of these carry enough water to fuel one MADV sortie. They have a solar-electric propulsion stage that doubles as a 375 kW solar-powered electrolysis plant. 375 kW at Terra orbit, it drops to 160 kW at Mars orbit due to the inverse-square law. WDV carries:

  • Tankage for 52,000 kg of water (full when launched)
  • Tankage for 40,000 kg of LOX/LH2 (6:1 ratio, empty when launched)
  • Solar-electric propulsion stage (375 kW at Terra orbit, 160 kW at Mars orbit)
  • Water electrolysis system (powered by SEP stage)
  • Navigation and communication systems

Upon arrival the WDV is captured by one of the Mars Base Camp's robotic arms, the WDV then starts electrolyzing the water. Two WDV working in parallel can create enough fuel for one MADV sortie in about 2.5 months. The base camp can hang on to two WDV at a time. If multiple sorties are planned, it will be better to have the WDV to cluster slightly ahead or behind of the base camp orbit to create an orbital propellant depot. This will mean the MADV will have to detach from the base camp and make a short trip to the propellant depot to fuel up.


The Mars Ascent/Descent Vehicles are sized for a 4.7 km/s entry velocity and have a mid-L/D profile. As previously mentioned they carry 4 crew, while 2 crew remain in orbit. They are totally reusable, which means no inflatable aerodynamic decelerators or other device that cannot be repackages and refurbished. Parachutes could be used were it not for the regrettable fact that the thin Martian atmosphere would require a prohibitively large chute to land a 100+ metric ton spacecraft. Retro thrust must be used.

The MADV has six RL-10 engines for propulsion. Wikipedia says they have 110,000 Newtons of thrust each for presumably a total of 660,000 Newtons.

Once landed the MADV will be home for the crew of 4 for the next 10 days (actually sols), though it has a 50% contingency margin for a total of 15 sols in case of emergency. It has a payload of 2,500 kg of scientific experiments. A 2 person airlock is adjacent to the 2 person mechanical lift on the spacecraft side which transport the crew to the planet's surface. It has enough consumables for two 2-person EVAs per sol.

The fuel tanks contain more fuel than is needed for a descent and ascent. The remaining fuel is used in fuel cells for power generation. It requires 780 m/s delta-V to land and 5,200 m/s delta-V to ascend, for a total of 6,000 m/s delta-V. The MADV has a propellant mass fraction of 74% (which I calculate to imply a mass ratio of 3.85).

Two water delivery vehicles (with 40,000 kg of LOX/LH2 each) can give the MADV enough fuel for one sortie. The report is not specific as to exactly how much is required, so the MADV requires something between 40,001 kg and 80,000 kg fuel for a sortie (probably the full 80,000). If so with the propellant mass fraction I calculate the wet mass of the MADV is about 108,000 kg. If true this is approximately the same mass as the Shuttle Orbiter (without the external tank and solid rocket boosters).

I did some pixel measuring on the cut-away view. Assuming that the human figures were the standard 1.77 meters tall (and the diagram is accurate), the MADV is approximately 25 meters tall (about 31 meters shorter than the Space Shuttle Stack).

Mars Expedition Spacecraft

This is from a NASA Manned Spaceflight Center (renamed the Johnson Space Center in 1973). The study was done in 1963. I have not been able to find lots of hard details, but there is some information in David Portree's monograph Humans to Mars on pages 15 to 18, available here.

It travels in a Hohmann transfer to Mars, separated into two parts spinning like a bola for artificial gravity. In Mars orbit, the heat shield, laboratory, and rendezvous ship separate and land. After a forty day stay, the astronauts use the rendezvous ship to climb back into orbit and travel to the mother ship. After the journey back to Terra, the astronauts land via the re-entry module.

Mars NEP with Artificial Gravity

This is from a document entitled Human Exploration Of Mars: Artificial-Gravity Nuclear Electric Propulsion Option, 15 July, 2003 which apparently has vanished from the face of the web so throughly that I cannot find it any more (but thoughtfully labeled "Internal NASA Use Only"). But you can find much of the details in this earlier report. Actually both are not so much reports as they are series of slides.

The Mars Crew Transfer Vehicle is basically a "tumbling pigeon" spacecraft with an ion drive powered by a nuclear reactor (nuclear-electric propulsion or "NEP"). The report mentions Ion thrusters, MPD thrusters, and VASIMR. At this point in the study they are assuming a specific impulse of 4,000 seconds (exhaust velocity of 39,200 m/s)


There are two basic mission classes: Short-Stay (opposition class) and Long-Stay (conjunction class). The important difference is once the astronauts have landed, how long is it until the launch window arrives? "Launch window" translates as "the date when the spacecraft had better depart for home or the astronauts will all die a lingering death from running out of oxygen". The report suggests that the first few mission will be short-stay because they are less risky. The longer you stay the bigger the chance that something will go wrong (vital equipment wears out, astronaut doing a surface excursion breaks their leg, somebody develops appendicitis, things like that).

The report decided the mission needed the following characteristics:

  • Initial missions limited to 18-24 month round trip (allows lesser performance engine to be used. Also minimizes steering requirements, which is a problem with huge fragile spinning artificial gravity ship)
  • Three months stay in Mars system
  • “Split mission” –no “Mars-specific” cargo sent out with crew (meaning the Lander is sent on an unmanned mission to High Mars Orbit first, it is not carried along with the manned mission. Means spacecraft can be generalized, not forced to be optimized to different types of Landers)
  • Assembly Orbit: Low Earth orbit 700 km (easier to assemble)
  • Departure/return point: High Earth orbit 90,000 km (requires less delta V and less propellant, assumes the presence of local orbital shuttles)
  • Destination: High Mars orbit 90,000 km (requires less delta V and less propellant)
  • Piloted vehicle stack less than 200 tons initial mass (less stress on spacecraft spine, less delta V, less propellant)

Prior to the Mars Crew Transfer Vehicle (MCTV) making the journey to Mars, the Mars Lander travels to High Mars Orbit unmanned and under remote control. Naturally if the lander fails to make it or arrives damaged the manned mission is scrubbed.

The components of the MCTV are boosted into 700 km LEO by a series of launches, and the components are assembled in orbit. It is then boosted into the 90,000 km HEO.

Crew is delivered to the MCTV by Earth Neighborhood transport infrastructure (XTV), which is some system of orbital transports. It then departs on the 10-odd month transit to High Mars Orbit.

At Mars Orbit, the MCTV rendezvouses with the Mars Lander and is loaded with the surface exploration crew. The poor Mikeys stay in orbit while being forgotten from the pages of history. The Lander lands, and the surface crew starts their 30 day surface stay. The Mikeys send them a constant stream of Mark Watney jokes.

At the end of 30 days the surface crew rides the Lander up to rendezvous with the MCTV. If something catastrophic happens during the lift off, the Mikeys will never forgive themselves for the jokes. Assuming all goes well the Lander is jettisoned and the MCTV departs on its ten-odd month trip home. In HEO it is met by an XTV craft and the crew is returned to Terra.

The mass returned to Terra is 89 metric tons.

Trajectory Sensitivity Analysis

The purpose of the trajectory sensitivity analysis is to determine tradeoffs and sensitivities of key trajectory parameters including:

  • Earth departure altitude
  • Mars parking orbit altitude (and ultimately Mars lander size)
  • Stay time in Mars vicinity
  • Useful time on Mars surface
  • Total trip time
  • Earth return altitude

Trajectory Assumptions:

  • Earth Departure Orbit: 700 km altitude
  • Earth Return Orbit: vary from 30,000 to 90,000 km altitude
  • Mars Parking Orbit: vary from 500 to 17,200 km altitude
  • Stay Time in Mars Orbit: calculated to sum time in Mars vicinityto approximately 90 days (Resulted in stay times at Mars in orbit from 37 to 77 days)
  • Total Trip time includes spiral time from LEO to high Earth orbit

Nuclear Electric Propulsion Vehicle System Assumptions:

  • Power: 6 MW
  • Specific Impulse (Isp): 4000 sec
  • Thruster efficiency: 60%
  • Tankage Fraction: 5% (metal tank mass as percentage of propellant mass)

Mission Assumptions:

  • Mass returned to Earth: 89 metric tons
  • Launch Date: 2026
  • Stay time in Mars space: approx 90 days (Resulted in stay times at Mars in orbit from 37 to 77 days)
  • Mission Total Trip time goal: 700 days

Limiting Orbit Assumptions (for sensitivity trade)

  • Earth departure orbit altitude : LEO of 700 km
  • Earth return orbit altitude: vary between 30,000 - 90,000 km
  • Mars parking orbit altitude: vary between LMO of 500 km and aerosynchronous


  • Missions of 700-days round trip are possible with limits on Earth and Mars orbit altitude choices
  • Total trip time does not equal total crew time (Note: The astronauts will ascend to the NEP vehicle once it’s in the high earth altitude via a the XTV)
  • Trade studies needed to evaluate choice of Mars parking orbit with respect to Ascent/Descent vehicle versus NEP vehicle performance
  • Further analysis needed to evaluate proximity to Sun on return leg

Mars Crew Transfer Vehicle Engine

The ion drive may be either electromagnetic or electrostatic.

Mars Crew Transfer Vehicle Artificial Gravity

Lander and Aerobrake

Mars Umbrella Ship

RocketCat sez

There is just something about this surreal design that gets to you. People who briefly saw the deep space umbrella in 1957 still remember it. Totally unlike any other spacecraft you've ever seen. That is, except for science fiction ships from artist who also were haunted by the blasted thing.

Not a bad ship either. Except that pathetic one-lung ion drive is so weak that it takes a third of a year to reach orbit halfway between Terra and Luna. I'm sure we can do better than that today. Swap it out for a VASIMR or something and you'll have a ship that can go places and do things!

Stuhlinger Umbrella Ship
Total Travel
6.30×107 sec2 years
Payload1.50×108 g150 tons
Ionic Mass
2.20×10-22 g
Specific Power5.00×104 erg/sec g0.95 kW/kg
6.70×10-2 cm/sec26.7×10-5 G
Initial Mass
7.30×108 g730 tons
3.65×108 g367 tons
Power Planet
2.15×108 g215 tons
Total Power
1.15×1015 erg/sec114.5 MW
Total Electrical
2.29×1014 erg/sec22.9 MW
in Jet
2.06×1014 erg/sec20.9 MW
Driving Voltage4.88×103 volts
Total Ion
4.22×103 amperes
1.15×104 cm115 metres
Total Length8.50×103 cm85 metres
8.40×106 cm84 km/s
Thrust4.85×107 dyne48.5 kgf
*dm/dt5.77×100 g/sec0.00577 kg/s
*delta-vee5.82×106 cm/sec58.2 km/s
Values from paper. * values derived by Adam Crowl

Unusual spacecraft designed by Ernst Stuhlinger in 1957, based on a US Army Ballistic Missile Agency study. It made an appearance in a Walt Disney presentation "Mars and Beyond". 4 December 1957. David S. F. Portree, noted space history researcher and author of Wired's Beyond Apollo blog, managed to uncover the identity of Dr. Stuhlinger's report for me, it is NASA TMX-57089 Electrical Propulsion System for Space Ships with Nuclear Power Source by Ernst Stuhlinger, 1 July 1955. Thanks, David!

Detailed blueprints of this spacecraft can be found in the indispensable Spaceship Handbook by Jack Hagerty and Jon C. Rogers, or are available separately.

The spacecraft resembled a huge umbrella, with the parasol part being an enormous heat radiator.

At the very bottom is a 100 megawatt (thermal power) fast neutron nuclear reactor, mounted on a 100 meter boom to reduce the radiation impact on the crew habitat. A fast neutron reactor design was chosen because they can be built will a smaller mass and smaller size (reducing the size of the shadow shield). The reactor is capped with a shadow shield broad enough to cast a shadow over the entire heat radiator array. The part of the shadow shield closest to the reactor is 1.8 meters of beryllium. This stops most of the gamma rays, and slows down the neutrons enough that they can be stopped by an outer layer of boron. The shadow shield has a mass of 30 metric tons, and coupled with the boom distance it reduces the radiation flux at the habitat ring to 10 fast neutrons per second per cm2 and 100 gamma rays per second per cm2.

The liquid sodium will be carried in pipes constructed of molybdenum. The reactor will have a specific power around 0.1 kW per kg. It contains 0.6 cubic meters of uranium enriched 1.7%, and has a mass of 12 metric tons. No moderator or reflector is required. "Cool" liquid sodium (500° C) enters the reactor and leaves the reactor hot (800° C) at the rate of 300 kg/sec. The reactor contains 600 molybdenum pipes with an inner diameter of 1.8 cm and a length of 1 meter. Electromagnetic pumps move the liquid sodium, since it is metallic. Such pumps are used since the only way to make pumps that will operate continuously for over a year with high reliability is to have no moving parts. The pumps will consume about 100 kW.

The hot sodium enters the heat exchanger, where it heats up the cool silicon oil working fluid. The now cool liquid sodium goes back to the reactor to complete the cycle. The heat exchanger is used because silicon oil is more convenient as a working fluid, and because the liquid sodium becomes more radioactive with each pass through the reactor. The heat exchanger contains 3000 tubes for liquid sodium, with a total length of 1,800 meters and an inner diameter of 1.3 cm. The silicon oil is boiled into a vapor at 500° C under 20 atmospheres of pressure.

The hot oil vapor travels up the boom to a point just below the umbrella. There it runs a turbine which runs a generator creating electricity. The turbine is a low-pressure, multi-stage turbine with a high expansion ratio. Silicon oil was selected since it can carry heat and simultaneously lubricate the turbine, since this has to run continuously for over a year. Silicon oil is also liquid at 10° C, allowing the power plant to be started in space with no preheating equipment. The oil has a specific heat of about 0.4 cal per g per degree C, a heat of vaporization of 100 cal per g, a density of 1 g per cm3. If the umbrella heat radiator is at a temperature of 280° C, this implies that about 100 kg/sec must flow through the turbine. The feed pumps will consume about 200 kW. The total mass of the working fluid in the entire system will be about 8 metric tons.

Newton's third law in the turbine causes the section of the spacecraft from turbine upwards to rotate, including the ring habitat module and the umbrella heat radiator. The spin rate is about 1.5 rotations per minute. The generator is cooled by small square heat radiators mounted on the habitat ring.

The boom below the turbine is counter-rotated so it remains stationary. This is because the boom has the ion engine. If the boom was not counter-rotated, the ion engine would also rotate. The result would be a stationary ship behaving like a merry-go-round, spinning in place while spraying ions everywhere like an electric Catherine wheel.

The hot silicon oil vapor is injected into the central part of the rotating umbrella heat radiator (the radiator feed), and centrifugal force draws it through the radiator. The cooled oil is collected at the rim of the radiator, and pumped back to the reactor to complete the cycle. The rotation of the ring habitat module provides artificial gravity for the crew. The habitat ring is in the central part of the umbrella.

The umbrella heat radiator will have a temperature of 280° C. The silicon oil vapor will be reduced to the low pressure of 0.1 atmosphere, to reduce the required mass of heat radiator. The ship will be oriented so that the umbrella is always edge on to the Sun, for efficiency. The diameter of the umbrella will be about 100 meters, constructed of titanium. The wall thickness is 0.5 mm, the thickness of the disk is 6 cm near the center and 1 cm near the rim. The umbrella is composed of sectors, each with an inlet valve near the center and an outlet valve at the rim. If any sector is punctured by a meteorite, the valve will automatically shut until repairs can be made. The other sectors will have to take up the slack.

The electricity runs an ion drive, mounted on the lower boom at the ship's center of gravity. The ion drive uses cesium as propellant since that element is very easy to ionize. Cesium jets have a purplish-blue color. The umbrella section and the reactor have about the same mass, since the reactor is composed of uranium. The habitat ring has a bit more mass, this is why the ion drive is a bit above the midpoint of the boom.

Cesium has a density of 1930 kilograms per cubic meter. The spacecraft carries 332,000 kilograms of cesium reaction mass. This works out to 172 cubic meters of reaction mass, which would fit in a cube 5.6 meter on a side. Which is about the size of the block in between the ion drive and the landing boat, the one with the boom stuck through it. (ah, as it turns out my deduction was correct, now that I have the original report to read)

However, cesium propellant is now considered obsolete, nowadays ion thrusters instead use inert gases like xenon. Cesium and related propellants are admittedly easy to ionize, but they have a nasty habit of eroding away the ion drive accelerating electrodes. Xenon is inert and far less erosive, it is now the propellant of choice for ion drives.

Mounted opposite the ion drive is the Mars landing boat. It is attached so its center of gravity is along the thrust axis. This ensures that the umbrella ship's center of gravity does not change when the landing boat detaches. The landing boat uses a combination of rockets and parachutes to reach the surface of Mars. The upper half lift off to return to the orbiting umbrella ship.

The habitat ring has an outer radius of 19.5 meters, an inner radius of 15 meters, and a height of 6 meters (according to the blueprints). If I am doing my math properly, this implies an internal volume of 2,900 cubic meters, less the thickness of the walls. At a spin rate of 1.5 rotation per minute, that would give an artificial gravity of about 0.05g.

Above the umbrella and habitat ring is an airlock module containing two "bottle suit" space pods. Above that is a rack of four sounding rockets with instruments to probe the Martian atmosphere. At the top is the large rectangular antenna array.

The spacecraft is much lighter than an equivalent ship using chemical propulsion, and has a jaw-droppingly good mass ratio of 2.0, instead of 5.0 or more. However, the spacecraft's minuscule acceleration is close to making the ship unusable. It takes almost 100 days to reach an orbit only halfway between Terra and Luna. At day 124 it finally breaks free of Terra's gravity and enters Mars transfer orbit. It does not reach Mars capture orbit until day 367, but it takes an additional 45 to lower its orbit enough so that the landing boat can reach Mars. All in all, the umbrella ship takes about 142 days longer than a chemical ship for a Mars mission, due to the low acceleration. Which is bad news if you are trying to minimize the crew's exposure to cosmic radiation and solar proton storms.

The design might be improved by replacing the ion drive with an ion drive with more thrust, or with a magnetoplasmadynamic, VASIMR or other similar drive invented since 1957.

In his paper, Dr. Stuhlinger proposed that the Mars expedition be composed of a fleet of several ships. The Mars exploration equipment would be shared among all the ships. In addition, there would be some "cargo" ships. These would only carry enough propellant for a one-way trip, so they could transport a payload of 300 metric tons instead of 150. They would be manned by a skeleton crew, who would ride back to Terra on other ships.

Master artist Nick Stevens has recreated the umbrella ship in a series of images. Click to enlarge.

Blender artist Owen Egan is making his own recreation of the original Disney animation. I am quite impressed, looks just like the original.

I am not quite the artist that Nick Stevens and Owen Egan are, but I had to try my hand at it. Click to enlarge.

Martin Mars Mission System Study

Martin MMSS
EngineNTR Solid Core
5,000 MWth
Isp900 sec
Exhaust Vel8,830 m/s
10,000 kg
Rad Shield
10,000 kg
25,000 kg
25,800 kg
185,000 kg60,000 kg
Dry Mass255,800 kg130,800 kg
225,000 kg
Wet Mass480,800 kg355,800 kg
Mass Ratio1.882.72
ΔV5,574 m/s8,836 m/s
Transfer Time
220 to
300 days
100 to
170 days

This is from Manned Mars System Study (MMSS): Mars transportation and facility infrastructure study. Volume 2: Technical report (1990).

The basic design used a conventional liquid oxygen—liquid hydrogen cryogenic propulsion system, and was quickly mired in the boil-off problem. Plus the propellant mass was sizable.

They did a quick analysis of solid-core nuclear thermal, nuclear-electric with MPD thrusters, and solar-electric. The electric versions had a lower Initial Mass In LEO (IMLEO) but much longer transfer times. The nuclear-thermal on the other hand was far superior to the chemical cryogentic, with lower IMLEO and shorter transfer time.

Assumed ΔVs
Terra Departure ΔV3,800 m/s4,500 m/s5,500 m/s
Mars Arrival ΔV1,500 m/s2,500 m/s4,400 m/s
Mars Departure ΔV1,500 m/s2,500 m/s4,400 m/s
Terra Arrival ΔV3,800 m/s4,500 m/s6,200 m/s
TOTAL ΔV10,600 m/s14,000 m/s20,500 m/s
Flight Time
(each way)
200—330 days120—170 days80—120 days

For the cryogentic chemical vs nuclear thermal analysis, used a sample mission of a roundtrip voyage from LEO to a 250 km × 1 sol Martian orbit, with delta-Vs as per the above table. In addition the analysis assumed:

  • 60,000 kg roundtrip payload
  • Aerobraking at Terra and Mars
  • Aerobraking mass fractions are 15% for Conjunction mission, 20% for Medium energy mission, and 30% for High energy mission
  • NTR engine mass of 15,000 kg for Conjunction mission, 20,000 kg for Medium energy mission, and 30,000 kg for High energy mission
  • Staged engine burns
  • Cryogenic rocket stages with mass fraction of 0.9
  • Cryogenic rocket engine specific impulse of 470 seconds
  • NTR engine specific impulse of 900 seconds

The analysis used IMLEO as the metric, the lowest IMLEO wins. The results were:

Initial Mass in LEO
Cryo/no aerobrake958,000 kg2,479,000 kg23,211,000 kg
Cryo/aerobrake317,000 kg555,000 kg1,567000 kg
NTR/no aerobrake289,000 kg480,000 kg1,408,000 kg
NTR/aerobrake195,000 kg282,000 kg547,000 kg

As the missions increase in energy, so do the benefits of the NTR. However for High Energy missions, the NTR/no aerobrake is quite close to the mass of the Cryo/aerobrake. For that mission the NTR should use aerobraking to have a clear advantage over Cryo.

METTLE Mission To Europa

METTLE Mission
Length200 m
Hab Ring Radius45 m
Engine Power2.5 MW
Specific Impulse
(high gear)
29,969 s
Exhaust Velocity
(high gear)
294,000 m/s
(high gear)
17 N
Specific Impulse
(low gear)
2,956 s
Exhaust Velocity
(low gear)
29,000 m/s
(low gear)
172 N
Num Enginesx8
Total Engine
Power Req.
~20 MWe
Total Thrust
(high gear)
136 N
Total Thrust
(low gear)
1,376 N
Reactor Power5 MWe
Waste Heat
15 MWt
Num Reactors6
Total Reactor
30 MWe
Total Reactor
Waste Heat
90 MWt
Habitat Ring300 MT
Descent Vehicle300 MT
Remaining S/C1,400 MT
ΔV29,000 to
32,000 m/s
Mass Ratio1.10 to 1.12

This is from a student study Human Missions to Europa and Titan - Why Not? (2004)

The spacecraft has an overall length of 200 meters with a habitat module in the form of a ring with a 45 meter outer radius.

The thrust frame supports eight VASIMR engines mounted as four pairs (the report was skeptical about how much the VASIMR could actually vary its thrust, so they played it safe and varied it by assuming non-variable thrust and using multiple engines). Engine heat radiators are mounted betwen the engines.

The six nuclear reactors are mounted on three reactor support booms, two reactors per boom. Each reactor can produce 5 megawatts of electricity for a total of 30 megawatts (!!?!). The engines only require 20 MW so the mission can survive the loss of one boom. The reactors produce 90 MW of waste heat total, so the booms are coated with heat radiators. It is hard to tell from the diagram, but it looks to me like they have four radiator panels per boom, which drastically reduces the efficency to about 70% since the panels are shining heat into each other.

The reactors have shadow shields, you can tell because the heat radiators have been trimmed to stay in the shadow. Unfortunately if the radiators indicate the outline of the safe radiation shadow, to my eye it appears that the habitat ring is sticking into the deadly radiation zone.

The reactor support booms also houses the superconductive magnets and plasma injectors which create the artificial magnetosphere to protect the spacecraft from radiation.

From the point where the booms attach to the spacecraft's spine back to the aft end are mounted six propellant tanks containing argon reaction mass for the engines.

On the fore end of the spacecraft is the payload: the habitat ring and the Europa landing craft.


The habitat module is a ring with a 45 meter outer radius. It spins at 3 rpm to provide 0.45 g of artificial gravity. It is composed of 24 modules. Each has a floor space of 43 square meters, for a total of 1,032 square meters. The relatively large number of modules is to allow redundancy in module function, and to allow emergency isolation in case of depressurization or fire.


As always the delta V cost goes up when you lower the trip duration. See the graphs below. However the report warns that the graphs were calculated simplistically for an "impulse burn" which can only be performed by a high-thrust rocket, not a VASIMR which needs a long period of constant thrust. So the delta V cost should be given an additional margin of 20% to 30% more so as to take care of gravitational losses.

The report says for the outbound journey they selected a delta V of 14 km/s and 15 to 18 km/s for the return journey, giving a total flight time of 2.7 to 3.9 years. I tried drawing lines for these on the charts below but they do not seem to fit. Anyway according to my slide rule this implies an economical mass ratio of from 1.10 to 1.12.

Michael Nuclear Pulse Battleship

RocketCat sez

Oooooh, Yeah!!! The Orion-drive Michael Battleship is the biggest meanest son-of-a-spacer in the cosmos! Well, maybe second to the Project Orion Battleship.

Just look at that bad boy! Can't you just see that unstoppable titan blazing into orbit on a pillar of multiple nuclear explosions, ready to kick that alien bussard ramjet's buns up between its shoulder blades? The drawback to Orion-drive is that it don't scale down worth a darn. So they didn't even try. No "every gram counts" worries here, they freaking chopped the main guns off the freaking Battleship New Jersey and welded them on!

Any casaba howitzer weapons? Naw, spears of nuclear flame are too feeble. They are using full-blown freaking Excalibur bomb-pumped x-ray lasers! Not infrared, not visible light, not even ultraviolet. X-rays. Just like Teller intended.

What's that you ask? What about the pumping bomb? Well, this is an Orion-drive, moron. That's whats driving the ship. Spit out a few Excaliburs, they aim their hundreds of laser rods on their targets, then the next pulse unit simultaneously thrusts the ship and energizes the graser beams. Another jumbo-sized order of crispy-fried elephant, coming right up!

Still have megatons of payload allowance left over? Well, how about carrying a small fleet of gunships with nuclear missiles? And all four space shuttles?

The look on the elephant's faces was priceless! Michael is coming. And is he pissed!

Battleship Michael
PropulsionOrion Drive
Height226 meters
Diameter113 meters
Massbetween 35,000
and 50,000
metric tons

Warning: spoilers for the book Footfall by Larry Niven and Jerry Pournelle to follow. On the other hand, the novel came out decades ago in 1985. I mean, in the novel the U.S.S.R. still exists. It takes place in the far flung future year 1995.

Footfall is arguably the best "alien invasion" novel ever written. Just like The Mote in God's Eye is arguably the best "first contact" novel ever written. But I digress.

Aliens (called "Fithp") who look like baby elephants arrive from Alpha Centauri in a Bussard ramjet starship (hybrid Sleeper ship and Generation ship). The starship is named "Message Bearer." They immediately ditch the Bussard drive module into the Sun, destroying it. If the Fithp are defeated, the humans can jolly well build their own Bussard drive from scratch to travel to Alpha C and attack the Fithp homeworld.

The Fithp evolved from herd animals, unlike humans. They have a very alien idea of conflict resolution. When two herds meet, they fight until it was obvious which one was superior. Then everybody immediately stops fighting, and the inferior herd is peacefully incorporated into the superior tribe as second-class citizens. Fithp do not comprehend the concept of "diplomacy".

They make the unwise assumption that human beings operate the same way. Big mistake!

The Fithp have somewhat superior technology compared to humans. They attack and seize the Russian space station (the ISS was not started until 13 years after the novel was written), annihilate military sites and important infrastructure with rods from God, then invade Kansas. The Fithp think "Look, humans. We are obviously superior. Now is the time to stop fighting and be peacefully incorporated into our herd." The Fithp calmly wait for the human surrender.

Humans don't work that way (and they have no idea that the Fithp have such a bizarre way of interacting). They savagely counterattack with the National Guard and three US armored divisions. The Fithp are taken aback, and beat off the counterattack with orbital lasers and more rods from God. The humans respond with a combined Russian and US nuclear strike on Kansas, obliterating the Fithp invasion force and most of the Kansas heartland.

The Fithp start panicking. What is it going to take to make these crazy humans surrender?

Finally the Fithp decide to forgo all half-measures. They drop a small "dinosaur killer" asteroid on Terra. The asteroid is called "The Foot." This causes global environmental damage, and more or less kills everybody living in India. Surely this will make the humans surrender!

The Fithp obviously don't know humans very well.

The humans have their backs to the wall, since surrender is not in their nature. The US president has a tiger team of advisers, who were drawn from the ranks of science fiction authors. After all, they are the only experts on alien invasions (in the novel, the various advisers are thinly disguised versions of actual real-world authors. Nat Reynolds is Larry Niven, Wade Curtis is Jerry Pournelle, and Bob Anson is Robert Anson Heinlein). They have got to find a way to carry the battle to the enemy: the orbiting starship and the fleet of "digit" ships. But how do you get thousands of tons of military hardware into orbit quickly enough not to be shot down while in flight using only technology they can develop in a dozen months?

There is only one answer. Project Orion. Old boom-boom. And to heck with the limited nuclear test-ban treaty that killed the project in 1963.

Orion has already been developed. Orion is mass-insensitive, it doesn't care if you are boosting tens of thousands of metric tons. This also means you can use quick and dirty engineering, since you are not stopping every five minutes trying to shave off a few grams of excess mass. You don't have to spend a decade trying to engineer featherweight kinetic energy weapons, just go tear the gun turrets off the Battleship New Jersey and weld 'em on. You can also carry a fleet of gunboats. And all four space shuttles.

The gunboats are going to be quick and dirty as well. Spaceships built around a main gun off a Navy ship, firing nuclear shells. Yes, a spinal mounted weapon

What about the Orion drive battleship's main weapon? Heh. Another cancelled project rises from the grave.

Back in the days of the Strategic Defense Initiative, Edward Teller et al came up with Project Excalibur. What was that? No less than bomb-pumped x-ray lasers. But wait, what about the bomb you need to pump the laser? Well, Orion is an nuclear-bomb-powered drive, remember? Make the propulsive bombs do double duty.

The weapons are called "spurt bombs." Dispensers on the pusher plate eject a flight of the little darlings. The spurt bombs unfurl their 100 laser rods apiece and aim them at Fithp ships. The next nuclear pulse unit is positioned, then detonated. This simultaneously gives thrust to the spacecraft, and pumps all of the spurts bombs. The Fithp ships are sliced and diced by a hail of x-ray laser beams. Spurt bombs look like fasces, "bundles of tubes around an axis made up of attitude jets and cameras and a computer."

Note that the nuclear pulse units will have to be specially designed. Standard Orion pulse units are nuclear shaped charges, designed to channel 80% of the x-rays upwards into the pusher plate (well, to create a jet of plasma directed at the plate but I digress). The battleship's pulse units need to be designed to also direct x-rays at the spurt bombs.

What is the battleship's name? Michael of course. The Biblical Archangel who cast Lucifer out of heaven.

The Michael launches through a cloud of Fithp digit ships, cutting them to pieces but suffering serious damage. The Fithp defecate in their pants and frantically rip the starship out of orbit and start running away. Their superior acceleration make escape possible, up until the point where the crew of the Space Shuttle Atlantis commits suicide and rams the starship. The main drive is damaged, and their acceleration is no longer higher than the Michael. Who catches up and starts pounding the living snot out of the starship.

There is something breathtaking about the Michael that captures the imagination of science fiction fans. On pretty much every single online forum about spacecraft combat, it isn't long until somebody brings it up. There have been many examples of fans trying to make blueprints, illustrations, or even scratch-build models of the battleship.

The original Michael diagram was made by Aldo Spadoni, president of Aerospace Imagineering. Mr. Spadoni is an MIT educated mechanical/aerospace engineer with over 30 years of experience designing and developing advanced aerospace vehicle and weapon system concepts (with most of the more advanced work being classified). He is also a personal friend of Larry Niven and Jerry Pournelle.

Mr. Spadoni did the Michael diagrams around 1997, working directly with Niven and Pournelle. They went through several iterations to arrive at the resulting diagrams.

Aldo Spadoni's Michael

However, this does bring up a good point that Scott alluded to. Footfall is a novel of course, not an engineering proposal for a space battleship. You glean details regarding the various Footfall spacecraft from the conversations of characters in the story, many of which are not experts wrt what they are describing. As Scott also pointed out, there are inconsistencies in the descriptions that are either intentional or simply mistakes on the part of the authors. Thus, the design of the Footfall spacecraft are open to interpretation.

As an engineer and concept designer, I particularly like the way Larry and Jerry write their stories. They provide enough big picture detail to determine the general design direction for their concepts, but leave the smaller details and the system integration issues to anyone willing to take a crack at envisioning their concepts. Fun stuff! So, I think my overall design captures what the authors intended, but many of the details are open to different interpretations, as some of you have done here.

As I move into discussing some of Michael’s details, I want to note that my primary design goal was to be true to the novel and the authors’ intentions as I understood them. I have my own vision of what a space battleship might look like, as I’m sure many of you have. But that’s not the subject of this design exercise.

As did Scott (Lowther), I struggled to determine Michael’s overall dimensions, given the novel’s inconsistencies. Whatever they wrote, Larry and Jerry envisioned the most compact possible vehicle that would get the job done. Note that Scott is showing an older version of my drawing that shows Michael with the shock absorber array fully compressed along with incorrect dimensions. The final dimensions I came up with are somewhat larger, on the same order as those Scott mentions in a separate post.

Regarding the comment that this is a slick ILM Hollywood design, I think this is reading quite a lot into a hemisphere, a rectangular prism and a shallow cone! Perhaps the commenter is confusing vehicle configuration design with render quality. These drawings were never intended to portray Michael’s actual exterior finish, surface markings, etc. These drawings were created way back when using an ancient vector-based illustration software application called MacDraw Pro. They look pretty awful and it’s certainly not the way I would render Michael today. In hindsight, I should have left them as line drawings and avoided the use of MacDraw Pro’s primitive shading tools.

Regarding the battleship-derived gun turrets, I agree with Scott’s assertion that the text of the book is vague in this area. But based on my discussions with Larry and Jerry, the authors definitely intended for Michael to include two of the full-up 16-inch Iowa class turrets, as well as some smaller gun turrets, not the guns alone.

Regarding the Shock absorbers array configuration, I disagree with you guys. Thinking that Michael is a straight extrapolation of the conventional Orion design configurations is incorrect. The primary purpose of the shock absorber array is, of course, to smooth out the “ride” for the payload/passenger portion of the vehicle. Most of the Orion designs were configured for non-military applications, whereas Michael is a maneuvering warship with massive nuclear pumped steam attitude thruster arrays. In addition to primary Orion thrusting, Michael will be subjected to multi-axial mechanical loads that are NOT along the longitudinal axis of the ship. Also consider that Michael’s design incorporates a pusher “shell” that is far more massive as a fraction of total vehicle mass than the typical Orion pusher plate design. When Michael is thrusting under primary propulsion while engaging in combat maneuvers, an angled shock absorber array design is a good choice for handling the inevitable side loads and for stabilizing the shell wrt the passenger/payload “brick”. Consider a high performance off road vehicle, which must provide chassis stability while the wheels and suspension are being subjected to loads from many directions. You don’t see any parallel straight up and down shock absorbers in the suspension system, do you?

If you look carefully at me design, you can see that that central shock absorber is longitudinal and more massive than the rest. This one is primarily responsible for handling the Orion propulsive loads. Perhaps it should be a bit beefier than I’ve depicted it in the original drawing. The remaining angled shock absorbers handle some of that propulsive load while also providing multi-axial stability. Admittedly, these 2D drawings don’t convey the Shock absorber array configuration that I have envisioned very well.

Since the time these drawing were created, I’ve discussed Michael with Larry and Jerry on a number of occasions. I’ve reconsidered and refined many of Michael’s technical aspects and I’ve designed a more detailed and representative configuration, including an updated shock absorber array. I’m also involved in creating my own high fidelity 3D model of Michael with a few fellow conspirators. I’m looking forward to sharing that with everyone at some point.

(ed note: one of those "few fellow conspirators" was me. Another was Andrew Presby, who is featured on one or two pages of this website.)

From a comment on the Unwanted Blog by Aldo Spadoni (2012)

Around 2010 Andrew Presby and I were commissioned by Aldo Spadoni to turn his Michael blueprints into 3D renders. Click for larger images.

Scott Lowther, author of Aerospace Projects Review is working on a book about nuclear space propulsion. Of course he wouldn't dream about leaving out the coolest Orion Drive spacecraft of all.

Now, strictly by the novel, the Michael is a mile high, which is ludicrous. The protagonists would have to have built a mile-high dome to cover it, which the aliens might have found a bit suspicious. In the diagrams below, Mr. Lowther shows the "large" Michael (one mile) and the more reasonable "small" Michael (1/8th mile).

Master artist William Black also had to turn his formidable talents on the Michael.

William Black's Michael

Nuclear pulse propulsion battleship Michael from the novel Footfall by Larry Niven & Jerry Pournelle.

“Michaels nose was a thick shield … armored in layers: steel armor, fiberglass matting, more steel armor, layer after layer of hard and nonresiliant soft.” —from Footfall, pg. 446 and 472

“Two great towers stood on the curve of the hemispherical shell, with cannon showing beneath the lip, aimed inward. Four smaller towers flanked them. A brick-shaped structure rose above them. The Brick was much less massive than the Shell, but its sides were covered with spacecraft: tiny gunships, and four Shuttles with tanks but no boosters. The bricks massive roof ran beyond the flanks to shield the Shuttles and gunships.”  —from Footfall, pg. 432

Michael is one of the Orion based concepts I knew I would have to take a run at sooner or later. I referenced the novel, extensively, and Scott Lowther condensed all the design bits he gleaned from Footfall into an Excel spreadsheet, available here, for a project he set aside. The spreadsheet is an excellent guide to all the passages describing Message Bearer, the digit ships, Michael, the stovepipes and Shuttles, and it proved invaluable in my effort.

Most people are probably familiar with Aldo Spadoni's visualization of the iconic warship from Niven and Pournelle’s novel, but for those who are not, Aldo’s drawings are available here.

What I’ve done is meet the Aldo Spadoni design half-way with my own interpretations. My intent was to complement Aldo’s design-thought without entirely rewriting it, keeping in mind what Aldo had to say about the process. One point Aldo raised in conversation on Scott Lowther’s blog is in regards to who is providing description in various scenes from the novel.

Aldo Spadoni: “Footfall is a novel of course, not an engineering proposal for a space battleship. You glean details regarding the various Footfall spacecraft from the conversations of characters in the story, many of which are not experts [with regard to] what they are describing. As Scott also pointed out, there are inconsistencies in the descriptions that are either intentional or simply mistakes on the part of the authors. Thus, the design of the Footfall spacecraft are open to interpretation.”

Aldo makes a good case for the distinctive angled shock absorbers of his design, and I’ll provide his commentary below, the sticking point for me, however, is the parabolic pusher plate Niven and Pournelle describe—early design work on Orion solidly ruled out a parabolic pusher. With shaped-charge nuclear pulse units the parabolic plate will only heat up while offering almost no thrust advantage. Heating and impact stress on the pusher would be of no small concern, the bombs necessary to loft something the scale and mass of Michael would not be the tame little devices used to propel a dinky NASA/USAF 10-meter Orion. Heating is the cost of even partially containing the ionized plasma resulting from nuclear detonation.

Orion works because the plasma is dynamically shaped (as the explosion happens) by the specially designed shaped charge nuclear explosive, X-rays are channeled by the radiation case in the instant before the weapon is vaporized, these exit a single aperture, striking and heating up a beryllium oxide channel filler and propellant disk (tungsten), resulting in a narrow conical jet of ionized tungsten plasma, traveling at high velocity (in excess of 1.5 × 10⁵ meters per second). This crashes into the pusher plate, accelerating the spacecraft. The jet is not physically contained by the pusher, and contact with the pusher is infinitesimally brief, so the pusher is not subject to extreme heating during thrust maneuvers. So, while offering very little performance difference compared to a flat pusher design, the parabolic plate would need regenerative cooling in the bargain, adding weight and complexity to the system. Engineering such a pusher plate would be fraught with difficulties, and conditions under which Michael is built, in my opinion, rule out any eccentric messing with the baseline system. A legion of Ted Taylors would already be kept busy night and day with the mere task of readying a conventional Orion designed under such circumstances—for delivery under a one year drop-rocks-from-orbit-dead deadline.

As Aldo points out, the text of Footfall leaves room for different interpretations and here is where I took some of Aldo’s design-thought and creatively merged it with my own toward the end of addressing the design as presented in the novel. (No, not the army of Ted Taylor clones inhabiting a maze of cubicles in some deep bunker somewhere—that’s just me.)

It occurred to me that what Aldo had done (following Niven and Pournelle’s description), was move the functions of the Orion standard propulsion module down, mounting them directly on the top of the plate, so really it’s a built up intermediate platform/propulsion module. What I’ve done is run with that thought: I chose to treat the entire pusher plate as an early large Orion: a dome sitting on flat pusher plate, concentric rows of toroidal shock absorbers surrounding a core array of gas-piston shock absorbers. There is no central hole-and-bomb-placement-gun-protection-tube in my design (but there is an anti-ablation oil spray system). Instead, pulse units are shot by bomb placement guns mounted to fire around the edge; exactly as in Aldo’s design (the early large Orion had rocket assisted bombs riding tracks on the exterior of the spacecraft—imagine the show that would make). The body of the “dome” in my design is stowage for tanked pressurization gas (for the shock absorbers), anti-ablation oil, and perhaps a reserve number of pulse units.

I’ve retained the scheme of duel pulse unit magazines. Niven & Pournelle called them “thrust bomb” towers. Four “spurt bomb” towers are also mounted to the base—the “spurt bomb” Niven and Pournelle describe is a type of bomb-pumped laser using gamma-radiation rather than X-rays. All of my towers are a good deal beefier than those on Aldo’s design. Narrative in the novel describes the “thrust bomb” towers as doing double duty, providing an extra layer of armor and shielding for the CIC/control room, the nerve center of the spacecraft, which is located in the lower portion of the Brick, wedged between two large water tanks (and two nuclear reactor containment vessels). The water tanks are frozen at lift-off, providing Michael with an ample heat-sink.  

As I mentioned above, Aldo makes an excellent case for the angled shock absorbers on his design, his description below:

Aldo Spadoni: “Most of the Orion designs were configured for non-military applications, whereas Michael is a maneuvering warship with massive nuclear pumped steam attitude thruster arrays. In addition to primary Orion thrusting, Michael will be subjected to multi-axial mechanical loads that are NOT along the longitudinal axis of the ship. … When Michael is thrusting under primary propulsion while engaging in combat maneuvers, an angled shock absorber array design is a good choice for handling the inevitable side loads and for stabilizing the shell [with regard to] the passenger/payload “brick.” Consider a high performance off road vehicle, which must provide chassis stability while the wheels and suspension are being subjected to loads from many directions. You don’t see any parallel straight up and down shock absorbers in the suspension system, do you?

If you look carefully at my design, you can see that that central shock absorber is longitudinal and more massive than the rest. This one is primarily responsible for handling the Orion propulsive loads. … The remaining angled shock absorbers handle some of that propulsive load while also providing multi-axial stability.”

Scott Lowther (of Aerospace Projects Review) offers this insight in regards to angled shock absorbers:

Scott Lowther: “I remain unconvinced at the off-axis "angled" shock absorbers, but they seem to be the popular approach. However, if you do go that route, you have to deal with the central piston in the same way... ball joints fore and aft. *All* the pistons must be free to swing from side to side. If one, even the central one, is locked, then either the pusher assembly cannot move sideways *thus negating the value of the angled shocks), or it'll simply get ripped off its mounts the first time there's an off-axis blast.

Given that the ship is clearly described as having nuclear steam rockets for attitude control, I don't see the value in off-axis blasts for steering. But... shrug.”

I spent a good deal of time reproducing Aldo’s shock absorber array because frankly I think it is brilliant, going back and forth between Aldo’s drawings and my file … in the end the detail would be invisible, so I created a cutaway render with two of the “spurt bomb” towers removed to reveal the system.

True to the novel Michael’s main guns are the 16"/50 caliber Mark 7 gun and turret taken directly off the New Jersey. There is a good deal of discussion (on Scott’s blog and elsewhere) on the suitability of the guns and turrets—the mounting is rotated ninety degrees to vertical relative to the orientation turret, guns, and loading mechanisms were designed for—however, Aldo is quite clear that mounting the full turrets “as is” reflects the author’s intention, and so I’ve kept to their vision in this regard.

In the novel the guns are described firing a nuclear artillery round, this would be a modern version of the W23 15-20 kiloton nuclear round. The Mark 23 was a further development of the Army's Mk-9 & Mk-19 280mm artillery shell. This was a 15-20 kiloton nuclear warhead adapted to a 16 in naval shell used on the 4 Iowa Class Battleships1. 50 of these weapons were produced starting in 1956 but shortly after their introduction the four Iowa's were mothballed. The weapon stayed in the nuclear inventory until October 1962. Presumably under war conditions a new production run would produce the numbers necessary for Michael’s assault on Message Bearer.

Secondary batteries: a generic turret roughly based on the secondary turrets of the Iowa class.

Missile launchers based on the MK-41 Vertical Launching System (VLS).

The “Battle Management Array” is a set of phased-array radars and tight-beam communications antenna for passing targeting information to Michaels secondary spacecraft, all mounted to a pair of shock-isolated cab, each riding its own set of shock absorbers, one mounted atop each “thrust bomb” tower.  A fall-back set of communications antenna and radar are mounted beneath the overhang of the forward shield atop the Brick.

I’ve gone with the dimensions Scott arrived at, which Aldo confirmed in his comments on Scott’s blog: Length:742’ Diameter: 371’.

Different opinions have been offered in regards to Michael’s mass, between 35,000 and 50,000 tons have been opinioned on Scott Lowther’s blog. Pournelle was quoted as saying 2 million tons on one occasion, and 7 million tons on another.

Michaels launch, in the novel, is shortened for reasons of narrative brevity; one character wonders if there were perhaps 30 or more nuclear detonations. Putting Michael in orbit would require 8 minutes of powered flight and about 480 bombs lit off at one bomb per second.

The novel is clear that Michael carries four Space Shuttles mounted to their external tanks sans their SRBs. The number of Gunships is less clear. Nine Gunships are described as destroyed in combat, an unspecified number survive to confront Message Bearer in the final scene. Designing the most compact spacecraft necessary to fill the role, my Gunship measures 100 feet in length, 25 feet in diameter. At these dimensions, 14 Gunships total can be comfortably mounted to Michaels flanks.

For detail on my Gunship design see my following post, Gunship.

1 W23

From Michael by William Black (2015)
William Black's Gunship

Gunship from Larry Niven and Jerry Pournelle's novel Footfall. See my related post Michael for additional detail.

“They take one of the main guns off a Navy ship. Wrap a spaceship around it. Not a lot of ship, just enough to steer it. Add an automatic loader and nuclear weapons for shells. Steer it with TV.” —from Footfall, pg. 354

In the novel these Gunships are referred to as “Stovepipe’s.” I was far less concerned with designing to match that narrative description than I was with designing the most compact spacecraft possible capable of the mission described. Michaels construction (including all its auxiliary spacecraft and subsystems) takes place in secret under wartime conditions, perhaps the moniker is derived from a code name picked randomly (that’s how the 1958 Project Orion was named), or perhaps dockworkers handling the vehicle sections, packed in featureless cylindrical shipping containers strapped to pallets, named the craft, and it stuck. See Aldo Spadoni’s commentary on character-delivered descriptions on my Michael post.    

I built my Gunship around the 5"/54 caliber Mark 45 gun.

Nuclear Round

The nuclear round fired by the Gunship would be something akin to the UCLR1 Swift, a 622 mm long, 127 mm diameter nuclear shell, weighing in at 43.5 kg.

In 1958 a fusion warhead was developed and tested. At its test it yielded only 190 tons; it failed to achieve fusion and only the initial fission explosion worked correctly. There are unconfirmed reports that work on similar concepts continued into the 1970s and resulted in a one-kiloton warhead design for 5-inch (127 mm) naval gun rounds, these, however, were never deployed as operational weapons. See paragraph 9 (not counting the bulleted list) under United States Nuclear Artillery.

Gunship Crew & Crew Module  

The text of the novel is unclear on the number of crew manning the Gunships, but my opinion is no more than 2 would be required, and dialogue in the novel tends to back this up. The loading mechanism is automated, so only targeting and piloting skills are involved. Considering urgency involved in readying Michael, I doubt an entirely new capsule, man-rated for spaceflight, would be considered. Michaels designers would fall back on tried and tested designs and modify them as required. In this case a stripped down Gemini spacecraft and its Equipment Module fits the bill nicely. The life support system matches the mission requirements. Leave off the heat shield (these are one-way missions), and reaction control system—the capsule never operates separate from the Gunship rig. Mount targeting and firing controls for the gun. Probably a single hatch rather than Gemini’s double hatch, and internal flat-screen displays rather than viewports—looking on this battle with naked eyes would leave the astronaut seared, radiation burned, and blinded.

“The exhausts of the gunboats were bright and yellow: solid fuel rockets.” —from Footfall, pg. 454

Eight SRBs akin to the GEM-40 allow options: they could be fired in pairs, allowing four separate burns, or two burns of 4, or a single burn of all eight – needs depending. The SRBs are strapped around a ten foot diameter 40 foot long core containing ample tank stowage for hypergolic reaction control propellants, pressurization gas, and nitrogen for clearing the breech and gun barrel. The reaction control system is used to aim the gun; propellant expenditure would be prodigious.

1UCRL - University of California Radiation Laboratory

From Gunship by William Black (2015)

Mini-Mag Orion

Mini-Mag Orion
PropulsionMini-Mag Orion
Thrust1,870,000 n
Exhaust Velocity157,000 m/s
Thrust Power147 GW
Pulse Unit Energy340 GJ
Nozzle Efficiency87.1%
Nozzle Mass199.6 metric tons

Data is from Mini-Mag Orion Program Document: Final Report from Ralph Ewig's website.

The nuclear pulse Orion drive propulsion system had both reasonably high exhaust velocity coupled with incredible amounts of thrust, a rare and valuable combination. A pity it was driven by sequential detonation of hundreds of nuclear bombs, and required two stages of huge shock absorbers to prevent the spacecraft from being kicked to pieces.

Andrews Space & Technology tried to design a variant on the nuclear Orion that would reduce the drawbacks but keep the advantages. The result was the Mini-Mag Orion.

First off, they crafted the explosive pulses so each was more 50 to 500 gigajoules each, instead of the 20,000 gigajoules typically found in the nuclear Orion. Secondly they made the explosions triggered by the explosive charge being squeezed into critical mass using an external power source instead of each charge being a self-contained easily-weaponized device. Thirdly they made the blast thrust against the magnetic field of a series of superconducting rings (Magnetic Nozzle) instead of the nuclear Orion's flat metal pusher plate.

In the standard nuclear pulse Orion, the pulse units are totally self-contained, that is, they are bombs. Since this makes it too easy to use the pulse units as impromptu weapons (which alarms the people in charge of funding such a spacecraft) a non-weaponizable pulse unit was designed. The Mini-Mag Orion pulse unit has the fissionable curium-245 nuclear explosive, an inexpensive Z-pinch coil to detonate it, but no power supply for the coil. The Z-pinch power comes from huge capacitor pulse power banks mounted on the spacecraft, i.e., the pulse unit ain't anywhere near being "self contained". The banks have a mass of a little over seven metric tons, far too large to use in a weapon (especially one that explodes with a pathetic 0.03 kilotons of yield). The Z-pinch coil should be inexpensive since it will be destroyed in the blast.

For a 50 gigajoule yield (with a burn fraction of 10%), the nuclear explosive is 42.9 grams of curium-245 in the form of a hollow sphere 1.27 centimeters radius (yes, I know the diagram above says the compression target is 0.47 centimeters radius. I think they mean the compressed size). This is coated with 15.2 grams of beryllium to act as a neutron reflector. According to the table below, a 120.7 gigajoule yield uses 21 grams of curium, which does not make sense to me. Usually you need more nuclear explosive to make a bigger burst. I guess the pulse units in the table have a larger burn fraction. The Z-pinch will squeeze the curium sphere from a radius of 1.27 centimeter down to 0.468 centimeters, leading to a chain reaction and nuclear explosion. Since curium-245 has a low spontaneous fission rate, the pulse unit will need a deuterium/tritium diode to provide the triggering neutrons. The pulse units will be detonated about one per second (1 Hz).

The Z-pinch needs 70 megaAmps of electricity. This is 70 million amps, which is a freaking lot of amps. The trouble is that you cannot lay big thick cables to the Z-pinch coil in the pulse unit. The cable will be vaporized by the nuclear explosion, which is OK. But a vaporized massive cable composed of heavy elements will drastically lower the exhaust velocity. This is very not OK. Remember that one of the selling points of the Mini-Mag Orion is the high exhaust velocity. Reduce the exhaust velocity and Mini-Mag Orion becomes much less attractive.

So instead of heavy cables the pulse unit uses gossamer thin sheets of Mylar (20 μ thick). I know that Mylar is usually considered an insulator, but 70 megaAmps does not care if it is an insulator or not. The report calls these Mylar cables Low Mass Transmission Lines (LMTL). They have a total mass of only 2 kilograms, which is good news for the exhaust velocity.

The 70 megaAmps go from the pulse power banks to permanent electrodes mounted on the magnetic nozzle. These take the form of five meter diameter metal rings. Two rings, positive and negative, just like the two slots in an electrical wall socket. The pulse unit proper is a minimum of 0.0244 meters diameter (double the 1.27 centimeter radius). So the LMTL has to stretch from the permanent electrodes to the pulse unit. This makes a five meter diameter disk of Mylar with with the grape sized pulse unit in the center. Actually two stacked Mylar disks (positive and negative) separated by about 2 centimeters of space (g0 in diagram above) so they won't short circuit. Ordinarily you'd use an insulator to prevent a short, something like, for instance, Mylar. Unfortunately here you are using Mylar as the conductor so instead you need a gap. The edge of each Mylar disk has an aluminum rim, each making contact with one of the magnetic nozzle's two permanent electrodes.

To place the pulse unit in the proper detonation point inside the magnetic nozzle, the pulse unit has to be five meters lower than the permanent electrodes in the nozzle. This forces the Mylar LMTL to be an upside down cone instead of a flat disk.

The pulse unit, Mylar LMTL and the aluminum rims are all vaporized during detonation. The magnetic nozzle with its permanent electrodes remain.

There are two power supplies: the steady-state reactor and the pulse power banks.

The reactor is the "charger." It charges up the superconducting magnetic nozzle, and gives the pulse power banks their initial charge. Finally it supplies power to the payload (including the habitat module). In the reference designs below, it outputs 103 kilowatts, has a mass of 9 metric tons, and is expected to supply 50 kilowatts to the payload. It takes 1 hour to give the pulse power banks (main and backup) their initial charge, and takes 39 hours to charge up the superconducting magnetic nozzle. Since the nozzle uses superconductors, its charge will last a long time before it leaks out.

The reactor has to supply 192 megajoules over one hour to charge up the main and backup pulse power banks. The reactor has to supply 7,446 megajoules over 39 hours to charge up the superconducting nozzle.

The tiny bombs need 70 megaAmps in 1.2 microseconds in order to detonate, but the reactor can only produce that many amps in one hour. The standard solution is to use capacitors, which can be gradually filled up but can dump all their stored energy almost instantly. This is the pulse power banks, a Marx bank of capacitors.

The reactor takes half an hour to charge up one pulse power bank, one hour to charge up the bank and the backup bank. The bank discharges all that energy into the pulse unit to detonate it. A separate system in the magnetic nozzle converts about 1 percent of the explosion into electricity and totally recharges the pulse power bank. For subsequent detonations, the reactor is not needed, the detonating bombs supply the power.

In the reference design, the pulse power banks hold 96 megaJoules per bank, there is a main bank and a backup bank for a total of 192 megaJoules, each bank has a mass of 3.5 metric tons, main and backup bank have a combined mass of 7.1 metric tons. The banks have to sustain a pulse unit detonation rate of 1 per second (1 Hz).

The backup bank is in case of a misfire, resulting in a lack of a recharge for the main bank. The still-full backup bank takes over energizing the pulse detonations while the reactor starts slowly re-charging the main bank.

Since the electrical system will be operating at megawatt levels, it will need a sizable set of heat radiators (Thermal Management System). By "sizable" we mean "up to 30% of the spacecraft's dry mass." In the first reference mission, the radiators have to handle 2,576 kW of waste heat, with the radiators having a mass of 15,456 kg and a surface area of 7,728 square meters.

The heat radiators are tapered in order to keep them inside the shadow cast by the radiation shadow shield. This keeps the radiators relatively free of neutron activation and neutron embrittlement. It also prevents the radiators from backscattering deadly nuclear radiation into the crew compartment.

The engine core and feed mechanism will have to inject the pulse units into the detonation point at rates of up to 1 Hz. It too will need redundancy and a minimum of moving parts.

In the second diagram above:

  1. Cycle begins. A pulse unit is at the detonation point with its LMTL contact rings touching the magnetic nozzle's permanent electrodes. Both blast doors are closed. The nozzle is fully extended.
  2. 70 megaAmps detonates the pulse unit. The explosion transmits force into the magnetic nozzle, producing thrust. 1% of the blast energy is converted into electricity which re-charges the pulse power bank. The nozzle moves upward along the feed system as part of the compression cycle. Meanwhile, the upper blast door opens to allow the next pulse unit to enter the feed system.
  3. The explosion plasma dissipates. The nozzle continues to move upward. As the next pulse unit enters the feed system, the upper blast door closes.
  4. The lower blast door opens. The nozzle reaches its highest position. The fresh pulse unit is injected into nozzle at the detonation point with a velocity matching the nozzle, LMTL contact rings of pulse unit touching nozzle's permanent electrodes. The lower blast door closes as the nozzle starts to travel downward along the feed system. When the nozzle reaches it lowest point, a new cycle begins.

The report had three sample "Design Reference Missions", and created optimal spacecraft using MiniMag Orion propulsion. As it turns out, the spacecraft for mission 1 and mission 2 were practically identical, so they only showed the two ship designs.

Design Reference Missions

DRM-1: Crewed Mars Mission:
50 kWe, 100 km/s Δv, 100 ton payload, 90 to 100 days one way trip time.
DRM-2: Crewed Jupiter Mission:
50 kWe, 100 km/s Δv, 100 ton payload, 2 years one way trip time.
DRM-3: Robot Pluto Sample Return:
50 kWe, 150 km/s Δv, 5 ton payload, 8 years one way trip time.

DRM-1/DRM-2 Spacecraft

DRM-1 Mass Budget
Mission delta-v100 km/s
Specific Power347 kW/kg
(347,400 W/kg)
Thrust Power87 gigawatts
Payload Mass100,000 kg
Specific Impulse9,500 sec
Exhaust Velocity93,000 m/s
Power System Mass (Charge)9,038 kg
Power System Mass (Pulse Banks)7,115 kg
Heat Radiators15,456 kg
Magnetic Nozzle Mass102,893 kg
Propellant Mass481,625 kg
Dry Mass (no remass, no payload)150,300 kg
Burnout Mass (no remass)250,300 kg
Ignition (Wet) Mass731,924 kg
Payload Fraction0.137
Propellant Fraction0.66
Dry Mass Fraction0.21
Power System - Pulse Banks
Peak Compression Current89 MA
Capacitor Voltage170 kV
Energy per Bank96 MJ
Capacitor Energy Density54 kJ/kg
Capactior Mass (one bank)1,779 kg
Pulse Bank Mass (total for 2 banks)7115 kg
Power System - Charge Power
Pulse Bank Charge Time60 minutes
Nozzle Charge Time39 hours
Pulse Banks Energy Content192 MJ
Nozzle Energy Content7,446 MJ
Payload Power Requirement50 kW
Power Output Electric103 kW
System Power Density11.4 W/kg
Thermal to Electric Efficiency0.04 fraction
Total Mass9,038 kg
Heat Radiators
Waste Heat Load2,576 kW
Area per Watt3 m2/kW
Mass per Area2 kg/m2
Radiator Area7,728 m2
Radiator Mass15,456 kg
Engine Performance
Specific Impulse9,500 sec
Exhaust Velocity93,164 m/s
Nozzle Efficiency0.45 fraction
Coupling Efficiency0.55 fraction
Pulse Yield120.7 GJ
Pulse Unit Mass6.9 kg
Standoff Distance5.9 m
Fission Assembly Mass21 g
Firing Rate1 Hz
Mass Flow6.9 kg/s
Thrust642 kN
Power29,894 MW
Gain563,631 ratio
Alpha (specific power)222,251 W/kg
Maximum Acceleration0.26 g's
Minimum Acceleration0.09 g's

DRM-3 Spacecraft

DRM-3 Mass Budget
Mission delta-v150 km/s
Specific Power551 kW/kg
(551,300 W/kg)
Thrust Power87 gigawatts
Payload Mass5,000 kg
Specific Impulse9,500 sec
Exhaust Velocity93,000 m/s
Power System Mass (Charge)9,067 kg
Power System Mass (Pulse Banks)7,115 kg
Heat Radiators15,505 kg
Magnetic Nozzle Mass102,895 kg
Propellant Mass630,963 kg
Dry Mass (no remass, no payload)152,723 kg
Burnout Mass (no remass)157,723 kg
Ignition (Wet) Mass788,686 kg
Payload Fraction0.006
Propellant Fraction0.8
Dry Mass Fraction0.19
Power System - Pulse Banks
Peak Compression Current89 MA
Capacitor Voltage170 kV
Energy per Bank96 MJ
Capacitor Energy Density54 kJ/kg
Capactior Mass (one bank)1,779 kg
Pulse Bank Mass (total for 2 banks)7115 kg
Power System - Charge Power
Pulse Bank Charge Time60 minutes
Nozzle Charge Time39 hours
Pulse Banks Energy Content192 MJ
Nozzle Energy Content7,447 MJ
Payload Power Requirement50 kW
Power Output Electric103 kW
System Power Density11.4 W/kg
Thermal to Electric Efficiency0.04 fraction
Total Mass9,038 kg
Heat Radiators
Waste Heat Load2,576 kW
Area per Watt3 m2/kW
Mass per Area2 kg/m2
Radiator Area7,752 m2
Radiator Mass15,505 kg
Engine Performance
Specific Impulse9,500 sec
Exhaust Velocity93,164 m/s
Nozzle Efficiency0.45 fraction
Coupling Efficiency0.55 fraction
Pulse Yield120.7 GJ
Pulse Unit Mass6.89 kg
Standoff Distance5.9 m
Fission Assembly Mass21 g
Firing Rate1 Hz
Mass Flow6.89 kg/s
Thrust642 kN
Power29,894 MW
Gain560,167 ratio
Alpha (specific power)222,122 W/kg
Maximum Acceleration0.41 g's
Minimum Acceleration0.08 g's

MOVERS Orbital Transfer Vehicle

Specific Impulse880 s
Exhaust Vel8,600 m/s
Thrust134,000 N
Endurance7 days
(21 person-days)
PowerFuel cells
Hab Module1,361 kg
Command Module363 kg
Power Systems
1,814 kg
RCS472 kg
and Rendezvous
471 kg
3,583 kg
NTR Engine1,814 kg
Shadow Shield3,856 kg
Propellant Tanks2,994 kg
w/ rad shielding
9,024 kg
DRY MASS25,753 kg
Payload0 kg
DRY MASS25,753 kg
Propellant42,317 kg
WET MASS68,070 kg
Mass Ratio2.64
ΔV8,380 m/s
Payload6,804 kg
DRY MASS32,557 kg
Propellant54,968 kg
WET MASS87,525 kg
Mass Ratio2.69
ΔV8,540 m/s

This is from Conceptual Design of a Manned Orbital Transfer Vehicle (1988). The function of the spacecraft was to deploy, recover, and repair satellites. Those things are expensive, it would be a vast saving to repair and refurbish satellites in place instead of sending up an entire new satellite. The report was prepared by the Modular Orbital Vehicle Engineering Research Society (MOVERS) of the University of Virginia.

The design criteria specified an ability to deliver and retrieve a payload of 6,800 kg from geosynchroneous orbit. A crew of three, life support for seven days, support for extra-vehicular activites. In addition the basic spacecraft should be adaptable to Terra-Luna missions with payloads up to 36,290 kg. This will be done by attaching more modules and propellant tanks.

The basic spacecraft has a delta V of about 8,400 m/s. Varying amounts of propellant are carried depending upon the payload mass, if any.

The elongated tanks are the main propellant tanks. There are four. Dimensions are 11.9 meters long by 4.5 meters in diameter. They carry a total of 42,317 kg of propellant (10,579 kg each), enough for flying with zero payload and 8,380 m/s of delta V.

The spherical tanks are the secondary propellant tanks. There are four. Dimensions are 4.6 meters long by 4.5 meters in diameter. They carry a total of 12,654 kg of propellant (3,162 kg each). With both the main and secondary tanks filled there is a total of 54,968 kg of propellant, enough for flying with 6,800 kg of payload and 8,540 m/s of delta V.

A sample mission servicing a Telstar satellite in GEO requires

Enter waiting ellipse775 m/s2 hours
11 minutes
Enter Hohmann transfer
Transit to Telestar
1,658 m/s5 hours
16 minutes
Match velocity with Telstar1,834 m/s< 14 minutes
Satellite servicing04.5 days
Enter Hohmann transfer
Transit to Terra LEO
1,834 m/s5 hours
16 minutes
Enter waiting ellipse
Dock with space station
2,169 m/s1 hour
43 minutes
TOTAL8,270 m/s5 days
2 hours
29 minutes


This is the heart of the spacecraft, how it earns its keep. The system can resupply fluid consumables to orbiting spacecraft (RCS fuel and coolant) as well as replace malfunctioned or obsolete components. It utilizes a waldo arm but has a backup of a Manned Manuvering Unit to allow an astronaut to go EVA and fix things manually. The entire servicing system is modular and be be detached from the core orbital transfer vehicle.


This supports the Remote Manipulator System (RMS) or waldo arm, which is basically the same as the one on the Space Shuttle. The RMS is approximately 15 meters long, it can safely manipulate a satellite up to 9 meters away. The notch in the ESM supports the waldo arm during periods of acceleration, so the blasted thing does not snap like a twig.

The ESM also has a cubby for the Manned Maneuvering Unit (MMU) or astronaut rocket backpack. The cubby has tanks of nitrogen propellant to recharge the MMU. There is also a cubby for the as-yet not developed Flight Telerobotic Servicer (FTS, a remote-controlled repair drone), for now the cubby is empty.


This contains up to six 1.1 meter diameter spherical tanks. These will be filled with whatever is needed to fill the empty tanks of the satellite being serivced, be it hydrazine, water, liquid helium, or whatever. Presumably future satellites will be equipped with fluid transfer connections that are established when the OTV docks. Older satellites will need the poor astronaut to go EVA, grab an umbilical from the FRS and manually fill up the satellite's empty tanks.


This is a satellite workstation, i.e., a place to tie down the blasted satellite so you can do repairs without it floating all over the place. The FSS is located along the ship's long axis so the center of gravity stays on the center. The FSS will have aids like astronaut foot restraints, propellant resupply umbilicals, power cables, and jacks for component diagnosis, testing, and checkout. The base of the FSS is a rack to store replacement satellite modules, with power feeds to keep the modules alive.


Module is 2.4 meters long by 4.5 meters in diameter. It houses all the command and control modules as well as spacesuits and other necessary equipment for EVA operations. Including the airlock. It is designed so that there is enough room in the main compartment for two astronauts to don spacesuits and allow both to enter the airlock.

The avionics and reaction control system (RCS) are more precise than most spacecraft, since they have to locate, track, and dock with relatively tiny satellites. They are described in excruciating detail in the report, but I won't bother to repeat it here. The equipment is 1988 vintage, which is laughably obsolete by now. Even if it is space-rated.


Module is 9.1 meters long by 4.5 meters in diameter. The interior consists of 21 service modules each one meter wide. These are installed on the four walls, leaving a 2.1 meter square opening in the center for the crew. The crew quarters are clustered at one end, they are wider than the service modules. Therefore the 2.1 m square center contracts to a 1.5 meter square hallway leading to the command module.

Each crew quarter displaces one and a half service bays and encloses an area of 4.3 cubic meters. The sleep restraint and personal use console are oriented parallel to exploit the free-fall environment. There is also a small window for recreational viewing. There are four rooms: three are crew quarters, one is designated as a "safe haven."

The wardroom provides space for a multi-use table and allows a large viewing window in the sidewall. The entire crew can occupy the wardroom simultaneously for eating or conferencing. It can also be used by off-duty crew for conversation or recreation.

The Environmental Control And Life Support System (ECLSS) supplies 54 kg of water per day, 18 kg per crew. Of that 18 kg per crew 6.8 kg is for drinking and food preparation while 11.1 kg for personal hygiene and wash water.

A treadmill and bicycle ergometer are provided as exercise machines to combat calcium bone loss and muscle tone depletion.

The personal hygiene facility provides privacy, contains accidental spills and controls odor. It has facilities for shaving, oral hygiene, hand/partial body washing, and a backup urinal for use in the event that the waste management compartment is occupied. It does NOT have a shower. That takes up too much room and has problems containing all the free-floating globules of water.

The waste management compartment is the toilet. It is much like the system that was used on the Space Shuttle.

The galley contains equipment for frozen, refrigerated and ambient food storage. Meal preparation subsystems include microwave/convection ovens, hot and cold water dispensers, utensil storage and pull-out counters. Clean-up and housekeeping is supported by inclusion of a trash compactor and stowage, and a convenient hand washer.

The hull of the module has a 5g/cm2 aluminum radiation shield. Dosage from the Van Allen radiation belts is estimated at 0.35 Sieverts. In case of a solar proton storm the spacecraft will aim its radiation shadow shield (atop the nuclear engine) to face the sun in lieu of equipping the ship with a full blown storm cellar. This will not provide as much radiation protection as a storm cellar, but the design cannot afford the savage reduction in payload mass.

On board power is supplied by a pair of H2-O2 fuel cells. They provide 12 kW at 2.78 VDC normally, and 16 kW at 26.5 VDC under emergency conditions. The fuel supply is 354 kg of liquid oxygen and 42 kg of liquid hydrogen. The pair of cells also create 44 kg of water per day. Solar power was too large and required constant panel adjustemnt. Nuclear power was too dangerous. Primary batteries had too low an alpha, batteries with enough watts would weigh too much. So fuel cells were chosen.

Avionics1.9 to 2.361 kW
Navigation0.8 kW
Crew Systems2.7 to 2.75 kW
Docking Equip.2.2 kW
Waldo Arm3.75 kW
TOTAL15.35 to 15.861 kW


Solid-core nuclear thermal rocket with a specific impulse of 880 seconds (exhaust velocity of 8,600 m/s) and 134,000 Newtons of thrust. Engine is 4.3 meters long with a diameter of 1.2 meters. Engine mass is 1,814 kg, the shadow shield mass is 3,856 kg.

MSFC NTR Mars Mission

This is from Nuclear Thermal Propulsion Mars Mission Systems Analysis and Requirements Definition (2007), a study by the Marshall Space Flight Center (MSFC) Advanced Concepts Office. The topic of the study was given [1] A Mars Mission and [2] Solid-core nuclear thermal rockets, what range of options where there and how did they compare?

The options depended upon a few key design decisions:

MARS STAY: Short-stay or Long-stay. Short-stay means a 30 to 70 days stay at Mars, and the total mission takes 600 days. Long-stay means a 550 day stay at Mars and the total mission takes 900 days.

ENGINE: All-propulsive or Aerocapture. This boils down to whether you save on propellant by using aerocapture or not. Both cases use solid-core NTR for Trans-Mars Injection (TMI). But for Mars Orbit Injection (MOI) the All Propulsive uses the NTR while the Aerocapture uses an aeroshell heat-shield and a close pass through the Martian atmosphere. In addition the All Propulsive uses NTR for Trans-Earth Injection (TEI) but the Aerocapture has to use chemical rockets due to packaging restrictions within the aeroshell. I think what the report is trying to say is that the NTR engine is to big to fit in the aeroshell, but an auxiliary chemical thruster will.

SPACECRAFT: All-up-mission or Split-mission. All-up means the entire mission is performed by one huge spacecraft. Split-mission uses a fleet of piloted and cargo spacecraft. Generally the uncrewed cargo ships are sent ahead. Only if they all arrive in Mars orbit is the piloted ship sent to join them. All-up and piloted split-missions are round-trip. Cargo split-missions are one-way.

So the design cases the study investigated were:

CASE 1: Short-stay, All-up, All-propulsive

CASE 2: Short-stay, Split-mission, All-propulsive

CASE 3: Short-stay, All-up, Aerobraking

CASE 4: Short-stay, Split-mission, Aerobraking

CASE 5: Long-stay, Split-mission, Piloted: All-propulsive, Cargo: Aerobraking


Obviously missions with crew are round-trip, while uncrewed cargo missions are one-way. So all-up missions and piloted split-mission are round-trip.

Uncrewed cargo split-missions are one-way, sent in advance of the piloted ships. Naturally the piloted ships are not sent until the cargo ships successfully arrive, otherwise what's the point?

All missions depart Terra from a 407 km circular parking orbit and are inserted into a 250 km by 33,793 km elliptical Mars orbit with a period of one Martian day.

When using aerocapture for Mars orbit insertion, the Martian altitude is assumed to be 125 km and maximum allowed arrival speed is 7.350 km/s (hyperbolic excess speed 5.450 km/s).

All missions (including all-propulsive) use aerocapture for Terra return, with a maximum allowed hyperbolic excess speed of 6.813 km/s.

As previously mentioned, the All-propulsive spacecraft use NTR for both TMI and TEI. But the aerobraking spacecraft use NTR for TMI and auxiliary chemical thrusters for TEI.

The report did an analysis and concluded that a 2033 mission start had the lowest Initial Mass In LEO (IMLEO) within a few decades, so that was chosen for the all-up and piloted split-missions. The cargo split-mission is started in 2030 so any failure will give enough advanced warning to abort the 2033 piloted mission.

2033 Piloted Mission Trajectory Data
Terra Departure
Mars Arrival
Earth Arrival
Aerocapture at Mars

2030 Cargo Mission Supporting Piloted Mission Trajectory Data
Terra Departure
Mars Arrival
Aerocapture at Mars

V is hyperbolic excess velocity (km/s). Ventry is atmospheric entry velocity (km/s). ΔV is delta-V, change in velocity require to perform specified maneuver.


Short-stay, All-up, All-propulsive

IMLEO602,000 kg
# Enginesx1
Main Engine
TypeSolid NTR
Thrust Nominal1,100,000 N
Thrust Range1,100,000 to
1,600,000 N
Isp Nom875 s
Isp Rng875 to 900 s
T/W Nom8.35
Engine Life Nom60 min
Engine Life Rng60 to 120 min
# Burns Nom10
# Burns Rng10 to 15
Engine Dia7 m
Engine Length15 m

Case one is for a short-stay at Mars. There is only a single spacecraft carrying everything. The payload is a habitat module with crew, a Crew Exploration Vehicle (CEV) and a Mars lander with its own atmospheric entry aeroshell. The spacecraft uses tumbling pigeon artificial gravity to create at least 0.3 g's. A nuclear thermal rocket engine is used for all maneuvers, including Trans-Mars injection, Mars orbit insertion, and Trans-Earth injection. When the spacecraft approaches Terra at the end of the mission, the crew abandons ship in the CEV and aerobrakes to a splash-down.

The main design drivers of this case was the propellant tanks and the overall vehicle length required to generate the minimum artificial gravity.

To conservatively avoid spin nausea you'll want to spin at 4 rpm and have the vehicle length be around 33.6 meters. If that is too long, you can force the astronaut to train, spin at 6 rpm, and get the vehicle length down to 15 meters. I'm just spitballing but looking at the image, if it is an isometric image, and the transhab is 10 meters tall, the spacecraft is about 106 meters long. Assuming the center of gravity in the center, the spin radius is a luxurious 53 meters. No spin nausea problems there. Even better, the center of gravity is probably closer to the engine because of the payload and the heavy nuclear engine. This means the spin radius for the hab module is even longer.


Short-stay, Split-mission, All-propulsive

IMLEO Piloted376,000 kg
IMLEO Cargo268,000 kg
# Enginesx1
Main Engine
TypeSolid NTR
Thrust Nominal670,000 N
Thrust Range645,000 to
1,600,000 N
Isp Nom875 s
Isp Rng875 to 900 s
T/W Nom7.52
Engine Life Nom60 min
Engine Life Rng60 to 120 min
# Burns Nom10
# Burns Rng10 to 15
Engine Dia7 m
Engine Length15 m

Case two is for a short-stay at Mars. An uncrewed cargo vehicle transports the lander to Mars parking orbit about 2.5 years before the piloted vehicle transports the astronauts. It uses its NTR for MOI.

The piloted vehicle carries only the habitat module, the CEV, and a transfer node for docking. It too uses its NTR for MOI. The propellant tanks are arranged to optimize the center of gravity for tumbling pigeon operations. Upon arrival in Mars orbit it docks with the cargo vessel using the transfer node.


Short-stay, All-up, Aerobraking

IMLEO439,000 kg
# Enginesx1
Main Engine
TypeSolid NTR
Thrust Nominal890,000 N
Thrust Range823,000 to
1,60,000 N
Isp Nom875 s
Isp Rng875 to 900 s
T/W Nom7.98
Engine Life Nom60 min
Engine Life Rng60 to 120 min
# Burns Nom5
# Burns Rng5 to 10
Engine Dia7 m
Engine Length15 m

Case three is for a short-stay at Mars. There is only a single spacecraft carrying everything. The NTR performs the TMI maneuver, the MOI is done by aerocapture, and the TEI maneuver is performed with a chemical stage.

The payload is carried in two separate aeroshells. Shell 1 holds the habitat module, the CEV, the transfer node, and the TEI chemical stage. Shell 2 is integrated with the Mars lander. When the vehicle approaches Mars, both aeroshells abandon the nuclear stage (letting it sail off into the wild black yonder) and both shells separately aerocapture into MOI. They then temporarily dock in Mars orbit using the transfer node, before the lander departs for the Martian surface. The lander uses its integral aeroshell a second time to get to the surface.

While tumbling pigeon gravity is provided on the Mars-bound leg of the mission, it cannot be used on the Terra-bound leg. The tumbling only works with a long spacecraft length. Unfortunately the ship's length is drastically shortened when it jettisons the nuclear stage. The crew will just have to suffer through free fall for the trip home.


Short-stay, Split-mission, Aerobraking

IMLEO Piloted290,000 kg
IMLEO Cargo198,000 kg
# Enginesx1
Main Engine
TypeSolid NTR
Thrust Nominal450,000 N
Thrust Range330,000 to
1,600,000 N
Isp Nom875 s
Isp Rng875 to 900 s
T/W Nom6.59
Engine Life Nom60 min
Engine Life Rng60 to 120 min
# Burns Nom5
# Burns Rng5 to 10
Engine Dia7 m
Engine Length15 m

Case four is for a short-stay at Mars. An uncrewed cargo vehicle transports the lander to Mars parking orbit about 2.5 years before the piloted vehicle transports the astronauts. It performs TMI with its NTR engine. The lander abandons the nuclear stage when approaching Mars and uses its aeroshell for MOI. It then waits patiently for the astronauts to arrive.

The piloted vehicle carries only the habitat module, the CEV, a transfer node for docking, and TEI chemical stage. All are housed in an aeroshell. Exactly like the cargo vehicle it performs TMI with its NTR engine, jettisons the nuclear stage when approaching Mars, and uses the aeroshell for MOI. It docks with the lander using the transfer node, then the explorers travel to the Martian surface.

Like case three, artificial gravity is only available in the Mars-bound leg of the mission.


Long-stay, Split-mission, Piloted: All-propulsive, Cargo: Aerobraking

IMLEO Piloted293,000 kg
IMLEO Cargo154,000 kg
# Enginesx1
Main Engine
TypeSolid NTR
Thrust Nominal330,000 N
Thrust Range200,000 to
450,000 N
Isp Nom875 s
Isp Rng875 to 900 s
T/W Nom6.59
Engine Life Nom60 min
Engine Life Rng60 to 120 min
# Burns Nom5
# Burns Rng5 to 10
Engine Dia7 m
Engine Length15 m

Case five is for a long-stay at Mars. It uses conjunction class trajectories instead of opposition class. Two uncrewed cargo vehicles are used. One delivers a Mars habitat to the surface, the second delivers a Mars lander into orbit.

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This week's featured addition is IEC Fusion Ship II

This week's featured addition is Martin Mars Mission System Study

This week's featured addition is Boeing STCAEM Cryo/Aerobrake

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