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

This section has been moved here.

AFFRE Mars DRA

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

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

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


AFTERBURNING FISSION FRAGMENT ROCKET ENGINE

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

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

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

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

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


HEAT RADIATORS

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

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

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


BRAYTON POWER CONVERSION UNITS

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

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

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


PERFORMANCE

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.

AIST-NTR

This is from Affordable In-Space Transportation (1996)

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

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

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

They found this is very hard to do.

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

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


Ground Rules:

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

NUCLEAR THERMAL IN SPACE TRANSPORT

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

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

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

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

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

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

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

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

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

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

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

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

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

Antares Dawn Battlecruiser

Battlecruiser Discovery
Engine
EnginePhoton drive
(with gears)
ΔV10,500,000 m/s
(10,500 km/s)
Thrust
Power
2.36×1014 W
(236 terawatts)
Photon
Power
req.
4.71×1014
(471 terawatts)
Powerfusion
(deuterium
enriched
hydrogen)
Fusion fuel
burn rate
0.73 kg/sec
Engine High Gear
Initial
Accel
9.81 m/s2
(1 g)
Thrust1,570,000 N
Exhaust
Vel
3×108 m/s
Specific
Impulse
30,600,000 sec
Engine Low Gear
Initial
Accel
39 m/s2
(4 g)
Thrust6,240,000 N
Exhaust
Vel
75,500,000 m/s
Specific
Impulse
7,690,000 sec
Ship
Length110 m
Body Dia14.5 m
Centrifuge
major
radius
22 m
Centrifuge
minor
radius
4 m
Centrifuge
volume
6,950 m3
Centrifuge
spin
1.0 g: 6.4 RPM
0.5 g: 4.5 RPM
0.1 g: 2.0 RPM
Centrifuge
type
dependant
Ship aprox
volume
32,000 m3
Ship aprox
density
5 kg/m3
Ship aprox
wet mass
160,000 kg
Parasite
craft
x4 armed scouts
Weaponsantimatter proj
particle beam
missiles
lasers
FTL energy10% fuel/jump

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

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


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


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

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

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

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

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

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

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

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

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

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

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

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

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

ANTARES DAWN

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

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

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

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

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

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

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

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

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

“Close outer doors. Stand by to repressurize.”


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

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


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


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

From ANTARES DAWN by Michael McCollum (1986)
ANTARES PASSAGE

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

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


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

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

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

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

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

Moliere had arrived.

From ANTARES PASSAGE by Michael McCollum (1998)

Asteroid Mining Crew Transport

This section has been moved here

Asteroid Survey Vehicle

This section has been moved here.

Atomic V-2 Rocket

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

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

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

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

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

Aurora CDF

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

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

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

The spacecraft is composed of four parts:

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

PROPULSION MODULE (PM)

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

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

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


Trans-Mars Injection Stages

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

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

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


Mars Orbit Insertion Stage

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

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

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

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


Trans Earth Injection Stage

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

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


TRANSFER HABITATION MODULE (THM)

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

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

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

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

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

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

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

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

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

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

MARS EXCURSION VEHICLE (MEV)

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

The MEV has three parts: the Surface Habitation Module (SHM) where the Mars explorers live, the Descent Module (DM) which does it 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.


SURFACE HABITATION MODULE

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

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

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

MARS ASCENT VEHICLE

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

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


MISSION

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

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

Aurora CDF Project Troy

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

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

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

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

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

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

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

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


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

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

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

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

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

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

Austin Mars Mission

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

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

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

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

In the diagram at left:

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

The three habitable components are:

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

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

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

The six propulsive stages are:

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

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

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

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

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

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

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

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

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

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

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


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

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

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

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


VEHICLE ASSUMPTIONS

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

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

Basic Solid Core NTR

Overview

RocketCat sez

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

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

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

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

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

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

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

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

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

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

Rescue Ship

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

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

Reactor

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

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

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

Shadow Shield

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

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

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

Plug Nozzle

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

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

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

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

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

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

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

Thrust Vectoring

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

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

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

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

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

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

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

Cascade Vanes

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

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

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

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

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

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

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

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

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

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

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

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

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

Bimodal NTR

This section has been moved here.

Bimodal Hybrid NTR NEP

This section has been moved here.

Bimodal Hybrid NTR NEP 2

This section has been moved here.

Benton Spaceship Discovery

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

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

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

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

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

Four types of landers were designed:

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

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

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

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

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


Crew Module (CM)

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

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

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

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

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

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

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


Docking Module (DM)

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


Service Module (SM)

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

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

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


Engineering Module (EM)

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

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

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


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

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

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

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

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

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

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

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


Strap-On Boosters

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

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

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

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


Crew Reentry Module (RM) for Terra Return


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


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


Design Reference Mission 2 (DRM 2) – Mars Exploration


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


Design Reference Mission 5 (DRM 5) – Callisto Exploration


Design Reference Mission 6 (DRM 6) – Ganymede Exploration


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


Comparison of Design Reference Missions

Boeing IMIS

RocketCat sez

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

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

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

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

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

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

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


Overview

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

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

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


Mission

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

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

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

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

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

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

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

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

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

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

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

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


Propulsion Module

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

These look suspiciously like the NASA reusable nuclear shuttle.

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

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

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

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

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

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

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

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

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


Mission Module

CREW COMPARTMENT

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

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


DECK 1

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


DECK 2

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


DECK 3

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

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

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


DECK 4

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

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

Boeing STCAEM Mars NTR

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

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

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

Boeing STCAEM Cryo/Aerobrake

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

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

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

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

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

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

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

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

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

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

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


Spacecraft


Aerobraking Shield


Transfer Vehicle


Terra Reentry Vehicle


Mars Lander


Spin Gravity Configuration


Aerobrake Shield Booster Vehicle

Boeing STCAEM Mars NEP

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

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

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

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

In the diagram below:

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

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

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


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

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

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

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

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

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


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


Microgravity Version (μg NEP)

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

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


Spin Gravity Version (Ga NEP)

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

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

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

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



Bono Mars Glider

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

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

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

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

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


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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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


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

Borowski Inspired Designs

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

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

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

All use solid core nuclear thermal rocket engines:

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

A. C. Clark

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

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

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

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

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

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

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

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


Habitat modules

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

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

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

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

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

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


Asteroid Survey Vehicle

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

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

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

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

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

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

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


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

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

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

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

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

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


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

ASV OPTION 1

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

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

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

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

ASV OPTION 2

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

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

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

ASV OPTION 3

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

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

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


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

ASV OPTION 4

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

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

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

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

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

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

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

ASV OPTION 5

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

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

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

ASV OPTION 6

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

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

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

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


Bimodal NTR

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

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

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

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

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

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

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

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

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

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

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

Bimodal Hybrid NTR NEP

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

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

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

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


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

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

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


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

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

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

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

Bimodal Hybrid NTR NEP 2

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

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

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

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

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

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



THE MISSION

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

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

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

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

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

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

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

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

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

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

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

NCPS Mars Mission

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

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

Abbreviations in table:

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

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


THE MISSION

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

Abbreviations in Trajectory Constraints table:

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

THE SPACECRAFT

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

Abbreviations in Design Constraints table:

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

NTP system consists of 3 elements:

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

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

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


BONUS SPACECRAFT

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

NLTV

This is from Robust Exploration and Commercial Missions to the Moon Using NTR LANTR Propulsion and Lunar-Derived Propellants (2017) doc, slides.

NLTV stands for Nuclear Lunar Transport Vehicle.

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

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

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

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

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

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


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

The Crewed Cargo Transport has its own dedicated habitat module weighing ~10 t, plus a 4-sided, concave star truss that has attached to it four 1.25 t payload pallets. The LO2 tank is smaller and customized for this particular application resulting in a lower Initial Mass In Low Earth Orbit (IMLEO or wet mass in 407 km altitude orbit) and LLO2 refueling requirement (~35 t).

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

ItemCargo TransportCommuter Shuttle
MissionLEO⇒LLO⇒LEO
Duration36-hr “1-way” transit times
Habitat Module~11.2 tn/a
Passenger
Transport
Module
n/a15.2 t
Crew42
Passengersn/a18
Star Truss
w/ 5 t payload
~8.6 tn/a
In-line
LO2 tank
~86.6 t~74.5 t
LH2 NTPS~70.9 t
IMLEO
(wet mass)
~177.4 t160.6 t
Refueled LLO2~71.6 t~67.9 t
Total
Burn Time
~25.3 min

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

Sampling of LANTR Vehicle Types
Case DescriptionObjectiveTrajectory/OrbitsIn-line LO2 TankResults
1cCrewed LANTR LTV
with MPCV
and 12 m saddle truss
carrying 5 t cargo to LLO
Determine LLO2 refueling needed
to deliver 5 t cargo to LLO
72 hour 1-way transit times
LEO–LLO–LEO
ΔV ~7.984 km/s
7.6 m OD x ~5.23 m L
(~163.5 t LO2)
IMLEO ~ 152.4 t
~48.8 t LO2 supplied in LEO
~46.9 t LLO2 refueling in LLO
2cCrewed space-based LANTR LTV
with 9.9 t habmodule
and 11 m star truss
carrying 5 t cargo to LLO
Determine LLO2 refueling needed
to deliver 5 t cargo to LLO
using alternative LTV configuration
72 hour 1-way transit times
LEO–LLO–LEO
ΔV ~7.996 km/s
4.6 m OD x ~3.4 m L
(~35.9 t LO2)
IMLEO ~ 131.1 t
~35.9 t LO2 supplied in LEO
~35.1 t LLO2 refueling in LLO
3cCrewed space-based LANTR LTV
with 9.9 t hab module
and 11 m star truss
carrying 5 t cargo to LLO
Determine LLO2 refueling needed
to deliver 5 t cargo to LLO
while also cutting transit times to 48 hrs
48 hour 1-way transit times
LEO–LLO–LEO
ΔV ~8.695 km/s
4.6 m OD x ~4.1 m L
(~48.0t LO2)
IMLEO ~ 143.4 t
~48.0 t LO2 supplied in LEO
~47.0 t LLO2 refueling in LLO
4cCrewed space-based LANTR LTV
with 9.9 t hab module
and 11 m star truss
carrying 5 t cargo to LLO
Determine LLO2 refueling needed
to deliver 5 t cargo to LLO
while also cutting transit times to 36 hrs
36 hour 1-way transit times
LEO–LLO–LEO
ΔV ~9.838 km/s
4.6 m OD x ~6.1 m L
(~81.2 t LO2)
IMLEO ~ 177.4 t
~81.2 t LO2 supplied in LEO
~71.6 t LLO2 refueling in LLO
5s LANTR commuter shuttle
carrying 15 t Passenger Transport Module (PTM)
to LLO then back to LEO
Determine LLO2 refueling needed
to deliver the PTM to and from LLO
with transit times of 36 hrs
36 hour 1-way transit times
LEO–LLO–LEO
ΔV ~9.835 km/s
4.6 m OD x ~5.4 m L
(~69.3 t LO2)
IMLEO ~ 160.6 t
~69.3 t LO2 supplied in LEO
~67.9 t LLO2 refueling in LLO
6sLANTR commuter shuttle
carrying 15 t PTM to LPO
then back to LEO
Determine LLO2 refueling needed
to deliver the PTM to and from LPO
with transit times of 36 hrs
36 hour 1-way transit times
LEO–LPO–LEO
ΔV ~10.006 km/s
4.6 m OD x ~6.0 m L
(~80.0 t LO2)
IMLEO ~ 172.5 t
~80.0 t LO2 supplied in LEO
~72.1 t LLO2refueling in LLO
7sLANTR commuter shuttle
carrying 15 t PTM to LPO
then back to LEO
Determine LLO2 refueling needed
to deliver the PTM to and from LPO
NTPS tops off with excess LLH2
36 hour 1-way transit times
LEO–LPO–LEO
ΔV ~10.047 km/s
4.6 m OD x ~4.6 m L
(~56.4 t LO2)
IMLEO ~148.2 t
LTV refuels with ~55.3 t LLO2
and NTPS tops off with ~6.9 t excess LLH2
8sRapid commuter shuttle
carrying 15 t PTM to LPO
then back to LEO
Determine feasibility of 24 hour transits
using twin LANTR engines
NTPS tops off with excess LLH2
24 hour 1-way transit times
LEO–LPO–LEO
ΔV ~13.225 km/s
4.6 m OD x ~8.3 m L
(~116.6 t LO2)
IMLEO ~204.3 t
LTV refuels with ~105.6 t LLO2
and NTPS tops off with ~13.2 t excess LLH2

Cases 1-4 are Crewed Cargo Transport Missions. Cases 5-8 are Commuter Shuttle Missions. Cases use a “Common NTPS” (carries ~39.7 t LH2). Propellant depots assumed in LEO, LLO and LPO. LANTR engines use optimized MRs. LEO=407 km Low Earth Orbit, LLO=300 km equatorial Low Lunar Orbit, LPO=300 km polar Lunar Polar Orbit. Total round trip mission ΔV values shown include g-losses

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

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

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

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

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

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

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

NTPS

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

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

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

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


PREVIOUS DESIGNS


GROUND RULES AND ASSUMPTIONS FOR NTPS MISSION AND PAYLOADS

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

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

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

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

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

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

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

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

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


GROUND RULES AND ASSUMPTIONS OF NUCLEAR THERMAL PROPULSION STAGE

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

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

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

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

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

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

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

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


LUNAR CARGO AND CREWED LANDING MISSION

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

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

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

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

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

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

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

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


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

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

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

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

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

To accommodate the MPCV, a saddle truss is used. The truss provides a nook for the MPCV to inhabit, a docking port and transfer tunnel connecting the MPCV with the LDAV, and photovoltaic arrays to energize the MPCV's life support system. It also has additional RCS jets. The MPCV does have its own photovoltaic arrays but they are difficult to deploy when inside the nook.

Unlike the uncrewed mission, the crewed mission carries more mass back to Terra (saddle truss, MPCV and LDAV). It needs more propellant, so the longer 23.7 m/49,000 kg LH2 in-line tank is used.

The resuable Lunar crew transfer mission departs from LEO (C3 or bare minimum escape velocity of 1,516 m2/s2) using a 2-perigee burn into Trans-Lunar Insertion requiring a delta-V (ΔVTLI) of 3,214 m/s (including a g-loss of 117 m/s).

About 72 hours later (3 days) arrives at Luna with an arrival Vinf (V) of 1,217 m2/s2. It captures into a 300 km circular Low Lunar Orbit (LLO) requiring a delta-V (ΔVLOC) of 913 m/s (including g-loss).

The key phases of the uncrewed Lunar crew transfer mission are shown below:

LUNAR DESCENT ASCENT VEHICLE
LDAV
Mass Schedule
Inert Mass6,100 kg
Payload: Crew Cabin2,200 kg
Payload: to Surface5,000 kg
Payload: x4 Crew
with EVA Suits
800 kg
DRY MASS14,100
LOX/LH2 Fuel20,900 kg
WET MASS35,000 kg
ΔV
Mass Ratio2.48
Isp450 sec
Exhaust Vel4,420 m/s
ΔV4,014 m/s
(actually 4,100
surface payload is unloaded)
ΔV Budget
Descent Start Mass35,000 kg
Descent Fuel Burnt13,400 kg
Descent End Mass21,600
Descent Mass Ratio1.62
Descent ΔVdes2,115 m/s
Ascent Start Mass15,100 kg
(-5,000 surface payload
-1,500 consumables)
Ascent Fuel Burnt5,500 kg
Ascent End Mass9,600
Ascent Mass Ratio1.57
Ascent ΔVasc1,985
TOTAL ΔV4,100 m/s

In LLO, the crew enters the LDAV and undocks. The surface payload containers rotate 180° into their landing position. They are stored snug to the crew cabin so that the LDAV has a small enough diameter to fit into the booster payload faring. They are rotated so on the lunar surface the payload is next to the ground for easy access, instead of inconveniently one story up in the air out of reach.

On the lunar surface, the crew can operate for three to 14 days using life support consumables carried in the surface payload containers. Alternatively they can operate for up to 180 days using one of the habitat landers. Upon departure they can carry back to the orbiting spacecraft up to 100 kg of lunar samples. The surface payload containers are left behind.

The LDAV climbs into orbit and does a rendezvous with the ship. It docks with the transfer tunnel on the saddle truss and the crew transfers to the MPCV. Then the ship departs from LLO (C3 of 949 m2/s2) into Trans-Earth Injection burn requiring a delta-V (ΔVTEI) of 856 m/s (including a g-loss).

72 hours later it arrives at Terra and captures into a 24-hour Eccentric Earth Orbit (EEO). The MPCV separates from the ship and the crew returns to Terra in the command module via aerobraking.


ASTEROID EXPLORATION MISSION

This is a crewed mission to explore Near Earth Asteroids like 2000 SG344. This spacecraft could support a month-long mission to an asteroid at Earth-Moon Lagrange 2, a reusable 327-day mission to 2000 SG344, or a non-resuable 178-day mission to the same.

The E-ML2 mission is 9.8 days outbound, 5.8 day stay exploring the asteroid, and a 17.4 day inbound transit. Total mission delta-V is 5,150 m/s including gravity losses, plus lunar flyby impulsive burns on both outbound and inbound legs.

The spacecraft uses an inflatable TransHab as a habitat module, so the poor astronauts don't have to live for months to years inside a cramped MPCV with less living space than three phone booths. Kids, ask your grandparents what a phone booth was, and why Superman got arrested for indecent exposure.

The spacecraft carries a MMSEV space pod, which is a sort of space-going Alvin submarine that was born to explore asteroids.

The far asteroid mission carries a MPCV, while the near asteroid mission carries a HL-20 Lifting Body instead. Either is to deliver the astronauts from Terra to the spacecraft, and to aerobrake them back to Terra Firma at the end of the mission.

The total spacecraft has an IMLEO wet mass of 170,800 kilograms, and a length of 79.3 meters. With 222,400 Newtons total from the engines at an Isp of 900 s, the 5,150 m/s delta V can be performed with a total burn time of 48.3 minutes.


SPACE TOURISM MISSION


COMPARATIVE SIZE GALLERY

NTR FIRST LUNAR OUTPOST

NTR FIRST LUNAR OUTPOST
Payload96,000 kg
EngineSolid core NTR
Pewee-class
Specific Impulse900 s
Exhaust Velocity8,800 m/s
Thrust110,000 N
Engine Mass3,437 kg
Shadow Shield Mass1,500 kg
Num Enginesx3
Total Thrust330,000 N
Total Engine Mass14,810 kg
Inert Mass33,330 kg
Dry Mass129,330 kg
Propellant Mass67,570 kg
Wet Mass196,900 kg
Mass Ratio1.5
ΔV3,658 m/s
MASS SCHEDULE
Structural Mass13,360 kg
Avionics & Power1,000 kg
Reaction Control460 kg
x3 Engines10,310 kg
x3 Shadow Shields4,500 kg
Contingency3,700 kg
INERT MASS33,330 kg
Payload 1
(FLO Lander)
93,000 kg
Payload 2
(FLO Adapter)
3,000 kg
Total Payload96,000 kg
DRY MASS129,330 kg
LH2 Propellant66,540 kg
RCS Propellant1,030 kg
Total Propellant67,570 kg
WET MASS196,900 kg

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

First Lunar Outpost (FLO) was one of NASA's "reference missions" studies. It was created in 1992. As with the other reference missions the mission parameters were nailed down, and researchers could design spacecraft capable of carrying out said missions. It got the ax shortly after 1992 for a variety of reasons.

The payload is 96 metric tons of lunar lander. This is 60 metric tons of lander stage which carries 36 tons of either: [a] cargo, [b] surface habitat, or [c] manned crew module with ascent/Terra-return stage.

The standard designs assumed that the lander would be transported to Lunar orbit by a conventional chemical propulsion module based around J-2S engines. But Stanley Borowski et al figured mission could be performed much more economically by using solid core nuclear rocket engines. The single J-2S chemical engine had a great thrust of 265 kilopounds force (klbf) (1,180,000 newtons) but a crummy specific impulse of 436 seconds (4,300 m/s exhaust velocity). A trio of NERVA derivative rocket (NDR) engines would only have a combined thrust of 75 klbf (25 klbf each) (330,000 newtons) but a much better Isp of 900 s (8,800 m/s Ve).

The nuclear stage carries 66.5 metric tons of liquid hydrogen propellant.

Bottom line is that the chemical stage had wet mass of 155 metric tons but an equivalent nuclear stage was only 101 metric tons. A savings of 54 metric tons is nothing to sneeze at. The nuclear stage is four meters longer than the chemical stage, but who cares?

After the lander detaches from the nuclear stage, the latter uses the RCS system to do a trailing edge lunar swingby. This provides enough of a gravity assist to put the spent stage into a disposal heliocentric orbit that will keep it away from Terra for at least one hundred thousand years.

"One Burn" Lunar Scenario
Trans-Lunar Injection (TLI) Payload
96 MT (pilot vehicle and TLI stage adaptor)
TLI Maneuver
ΔV3,200 m/s + gravity losses
Initial Orbit185 km circular LEO
NTR System
PropellantCryogenic hydrogen
Isp870 sec (graphite)
900 sec (composite)
960 sec (ternary carbide)
External shield mass60 kg/klbf thrust
(0.014 kg/Newton)
Burn Duration≤ 30 minutes
Flight Performance
Reserve
1% usable propellant
Cooldown (effective)3% usable propellant
Residual1.5% total tank capacity
RCS System
PropellantHydrazine
Isp237 seconds
TLI Burnout ΔV60 m/s
(30 m/s for trailing edge lunar flyby)
RCS System
Material2219-T87 Al
Geometry10 m diameter cylindrical tank
with √2/2 domes
Insulation2 inch MLI +
micrometeoroid shield (3.97 kg/m2)
Boiloff12.40 kg/day
Contingency
Engine &
external shields
15%
All other dry masses10%

Blue Max Studio Liberty Bell

Liberty Bell
ParamLaunchLunar Run
Structure
Mass
15075
Cargo
Capacity
250300
Dry
Mass
400375
Wet
Mass
820585
Thrust
(engine)
6.18 MN2.2 MN
Mass
Ratio
10.5 (4.2)*1.56
Accel10 m/s24.58 m/s2
Mass
Flow
1,132.4 kg/s1,000 kg/s
Exhaust
Velocity
5,457 m/s2,200 m/s
Specific
Impulse
556.2 s452.3 s
Burn
Duration
770 s105 s
Flight
Time
≈2 hrs5.56 days
Δv7,842 m/s4,947 m/s

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

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

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

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

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


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

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


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

However, our Liberty Bell is heading to Luna.

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

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

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

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


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

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

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


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

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

Closed-Cycle MHD Nuclear-Electric

This is from Prospects for Nuclear Electric Propulsion Using Closed-Cycle Magnetohydrodynamic Energy Conversion (2001).

NASA’s Dr. Ernst Stuhlinger, a leading authority on electric (ion) propulsion, has often said that such a rocket system would be ideal for a manned journey to Mars.

“Yeah,” a wag once cracked, “if you can just find an extension cord long enough."

From A FUNNY THING HAPPENED ON THE WAY TO THE MOON by Bob Ward (1969)

What the joke is saying is that Electrostatic (ion drives) and Electromagnetic (VASIMR) rockets are power hogs. While they have outstanding exhaust velocity/specific impulse, they need huge solar photovoltaic arrays or nuclear reactors whose mass is measured in metric tons. Which really cuts into your payload mass.

Photovoltaics are an attempt to use Sol as the power plant and sunlight as the extension cord. Trouble is that sunpower is relatively dilute, and the inverse square law shortens the length of the extension cord to about the orbit of Mars. A "robust" mission at any rate.

This design is an attempt to reduce the mass of a nuclear power plant so it can be used in an ion-drive ship without reducing the payload mass to a couple of kilograms.


MHD Nuclear-Electric
General
Payload100,000 kg
MissionTerra to Mars
Mission Durations
  • 120 days
  • 150 days
  • 180 days
Power Plant
TypeClosed-loop nuclear MHD Brayton cycle
Reactor Power100 MWth
Enthalpy Extraction Ratio40%
Isentropic Efficiency70%
Power Output40 MWe
Propulsion System
Propulsion Alpha1.1 kg/kWe
Isp5,000 to 8,000 sec
Engines
TypeIsp
(sec)
ηtα
(kg/kWe)
Ion (Kr)≥5,0000.81.0
MPD (Li)4,000—8,0000.50.5
MPD (H2)≥8,0000.50.5
VASIMR (H2)3,000—30,0000.50.2—1.0

NASA and other space agencies tend to focus on technologies that can be realized in the near term. I mean, antimatter power would be nice but that ain't gonna be available anytime short of a century (i.e., 25 presidential election cycles, which is the average time between radical NASA policy shifts). However, the paper makes a case that expending some effort on a technology that is just a couple of steps over near term can give huge returns.

The paper makes the case that the penalty-mass problem with spacecraft nuclear power plants is due to the fact they are based on closed-cycle gas turbine technology. These are limited to low-heat rejection temperatures, which result in large and massive heat radiators. And just guess which nuclear power component is the major culprit affecting power plant mass? Yep, the heat radiators. The mass of the radiators is huge compared to the mass of the reactor and energy conversion equipment.

How do you reduce the mass of the heat radiators? You run them at a higher temperature, that's how. Why don't they do that? Well, gas turbines contain turbine wheels. If you run the system at a higher temperature the blasted turbine blades melt and the turbine is destroyed. Even if the turbine wheels have an active cooling system. Sort of like spitting on a blast furnace in order to cool it down. You don't want to run the turbine much hotter than 1,200 K or so.

This is why the authors of the paper say we should abandon closed-cycle gas turbine technology and make the jump to closed-cycle magnetohydrodynamic (MHD) technology. On the minus side this technology is not quite as mature. On the plus side you can run that sucker at 2,500 K with little or no problem (report says "minimum development risk") with a corresponding drastic reduction in heat radiator mass. If you did some work high-temperature fissile fuels for the reactor, you could push that to 3,000 K. And in the future if you developed gas core nuclear reactors, it is estimated that the theoretical limit of MHD generators is as high as 8,000 to 10,000 K. That will really shrink the heat radiators down to size.

You see, with a gas turbine, the turbine blades are bathing in the ultra-hot blast of gas. With an MHD on the other hand, none of it actually touches the gas, it just surrounds it. Which drastically reduces the "melting generator" problem. The MHD can be cooled with good old regenerative cooling, just like the nozzle of a chemical rocket. MHDs also have no moving mechanical parts, which improves reliability and reduces maintenance.

The main problem is that the gas has to conduct electricity, which generally means you have to seed it with cesium dust or something like that. Then it becomes electric charges moving through a magnetic field, which is the basis for all electrical generators. It is just that the electric charges are moving at hypervelocity so it generates lottsa current.

The fact that the gas is accelerated by a fission reactor opens up another seeding possibility. If you seed the gas with an isotope with a large neutron interaction cross section, as it passes through the intense neutron flux inside the reactor the isotope dust will create nuclear ionization events. Not just one or two as with chemical ionization, each nuclear interaction can produce hundreds to thousands of ionization events.


As a benchmark the report authors set up a sample space mission to demonstrate the performance of this propulsion system.

The mission was to deliver a 100 tonne payload from a 1,000 km circular Terran orbit (i.e., high enough so that the reactor radiation would not reach Terra) to a 500 km Mars orbit. Several 2018 mission opportunities were examined for trip times of 120, 150, and 180 days.

A 100 MWth nuclear reactor was assumed, driving an MHD generator with an enthalpy extraction ratio of 40% and an isentropic efficiency of 70%. This means it will generate 40 MWe for the ion drive. Using near-term technology assumptions for the subsystems, this implies an overal propulsion system specific mass of 1.1 kg/kWe.

Figure 14 shows the Initial Mass in Low Earth Orbit (IMLEO) for a nuclear MHD powered ship with the above specifications, over a range of engine Specific Impulse (Isp). The ranges of Isp for the four engines covered in the report are shown in color. The three black lines show the values for trip times of 120, 150, and 180 days. Example: if you had an engine with an Isp of 2,000 seconds, the 180 day transit would require an initial mass in LEO of about 270 metric tons.

The sweet spot seems to be with Isp between 5,000 and 8,000 seconds. Note that in that range the payload can account for as much as half the IMLEO.

The CHEBY-TOP software the writers used to figure the mission trajectories also had a function to determine the optimal power for a given configuration. So they gave it a try. Figure 16 shows the results for the 120 day mission, delivering a 100 metric ton payload. At a specific impulse fo 2,000 seconds the optimal power was 13.3 MWe. At 10,000 sec the optimal power was 30.6 MWe. Since the system was sized with a power level of 40 MWe, it turns out that the design is actually oversized for the mission. But that's OK, the extra power can be used.

The extra power can be used in two alternate ways: faster trip time and/or larger payloads.

Figure 17 shows the payload increase option. Here the trip time is still 120 days, the triangle line shows payload mass, the circle line shows IMLEO.

For instance, for an engine with an Isp of 2,000 sec, it could deliver 293 metric tons to Mars in 120 days. The drawback is that the IMLEO mushrooms to an unattractively monstrous 880 tonnes! An Isp of 10,000 sec can only deliver 132 metric tons of payload in 120 days, but the IMLEO is a much more reasonable 255 metric tons. Any Isp higher than 6,000 sec will have a payload mass fraction (payloadMass / IMLEO) greater than 0.56, which is pretty darn attractive for a 120 day mission.

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).

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

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

Aeronutronic EMPIRE

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

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


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

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

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


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

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

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


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

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

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

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


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

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

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



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


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

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


The ship is now reconfigured into orbit mode.

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

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

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

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


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


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

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

Electria

Electria
EngineIon
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

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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

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GCNR Spacecraft

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GCNR Liberty Ship

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Hariven-class Free Trader

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Hedrick Fusion Spacecraft

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HELIOS WATERSKI

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HELIOS BOOM-BOOM

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Hermes from The Martian

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HOPE

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HOPE (FFRE)

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HOPE (MPD)

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HOPE Cargo vehicle

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HOPE Tanker

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HOPE Crew vehicle

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Stuhlinger Ion Rocket

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HOPE (MTF)

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

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HOPE (VASIMR)

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HOPE (Z-Pinch Fusion)

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Hyde Fusion Rocket

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Hyperion

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ICAN-II

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Kuck Mosquito

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LANTR LTV

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LCOTV

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Lewis Research Center GCNR

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Lewis Research Center Ion Rocket

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Lewis Research Center Mars Landing

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Lewis Research Center Mars Ref

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Luna from Destination Moon

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Lunar Transportation System

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Lighter and Tanker

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Manned Mars Explorer

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Mars Base Camp

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Mars Expedition Spacecraft

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Mars NEP with Artificial Gravity

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Mars Umbrella Ship

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Martin Mars Mission System Study

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METTLE Mission To Europa

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Michael Nuclear Pulse Battleship

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Mini-Mag Orion

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MOVERS Orbital Transfer Vehicle

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MSFC NTR Mars Mission

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Atomic Rockets notices

This week's featured addition is Reusable Nuclear Shuttle Class 3

This week's featured addition is Pulsed Plasmoid Mars Mission

This week's featured addition is Lockheed Nuclear Space Tug

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