Introduction

RocketCat sez

Here is your handy-dandy cheat-sheet of rocket engines. Use this as a jumping-off point, there is no way I can keep this up-to-date. Google is your friend!

I'll point out a few of the more useful items on the sheet:

  • Aluminum-Oxygen is feeble, but is great for a lunar base (the raw materials are in the dirt).
  • VASIMR is the current favorite among ion-drive fans. Use this with orbit-to-orbit ships that never land on a planet. It can "shift gears" like an automobile.
  • Solar Moth might be a good emergency back-up engine.
  • Nuclear Thermal Solid Core is better than feeble chemical rockets, but not as much as you'd expect.
  • Nuclear Thermal Vapor Core is what you design along the way while learning how to make a gas core atomic rocket.
  • Nuclear Thermal Gas Core Open-Cycle is a full-blown honest-to-Heinlein atomic rocket, spraying glowing radioactive death in its exhaust.
  • Nuclear Thermal Gas Core Closed-Cycle is an attempt to have the advantages of both nuclear solid core and gas core, but often has the disadvantages of both. It has about half the exhaust velocity of an open-cycle atomic rocket.
  • Orion Nuclear Pulse is a rocket driven by detonating hundreds of nuclear bombs. If you can get past freaking out about the "bomb" part, it actually has many advantages. Don't miss the Medusa variant.
  • Magneto Inertial Fusion This is the best fusion-power rocket design to date.
  • Zubrin's Nuclear Salt Water This is the most over-the-top rocket. Imagine a continuously detonating Orion drive. There are many scientist who question how the rocket can possibly survive turning the drive on.

There is a nice basic overview of propulsion systems here.

You can spend lots of time researching spacecraft propulsion systems. But you are in luck, I've got some data for you. Most of this is from Philip Eklund's out of print boardgame Rocket Flight, the impressive Spaceship Handbook, and the indispensable Space Propulsion Analysis and Design. The rest is from various places I found around the internet, and no, I didn't keep track of where I got them. Use at your own risk.

Philip Eklund has a new boardgame out called High Frontier, which has the Atomic Rockets seal of approval (be sure to get the expansion pack as well). It has even more cutting-edge but scientifically accurate propulsion systems, which will eventually find there way onto this web page. (more details here, here, here, and here.)

If you don't like the values in the table, do some research to see if you can discover values you like better. Also note that the designs in the list are probably optimized for high exhaust velocities at the expense of thrust. There is a chance that some can be altered to give enough thrust for lift-off at the expense of exhaust velocity. Or you can just give up and go beg Mr. Tyco Bass for some atomic tri-tetramethylbenzacarbonethylene. Four drops should do the trick.

Some engines require electricity in order to operate. These have their megawatt requirements listed under "Power Requirements". With these engines, the Engine Mass value includes the mass of the power plant (unless the value includes "+pp", which means the mass value does NOT include the mass of the power plant). The power plant mass can be omitted if the spacecraft relies on beamed power from a remote power station. Alas, I could find no figures on the mass of the power plant. If the plant is nuclear, it probably has a mass of around 0.5 to 10 tons per megawatt. If it is beamed power the mass is of course zero. Efficiency is the percentage of the power requirements megawatts that are actually turned into thrust. The rest becomes waste heat and has to be removed with heat radiators.

T/W >1.0 = Thrust to Weight ratio greater than zero? This boils down to: can this engine be used to take off from Terra's surface? If the answer is "no" use it only for orbit to orbit maneuvers. It is calculated by figuring if the given thrust can accelerate the engine mass greater than one gee of acceleration. As a rule of thumb, a practical spacecraft capable of lifting off from the Earth's surface will require a T/W of about 50 to 75.

Most propulsion systems fall into two categories: SUV and economy. SUV propulsion is like an SUV automobile: big and muscular, but the blasted thing gets a pathetic three miles to the gallon. Economy propulsion has fantastic fuel economy, but has trouble climbing low hills. In the world of rockets, good fuel economy means a high "specific impulse" (Isp) and high exhaust velocity. And muscle means a high thrust.

The only vaguely possible propulsion system that has both high exhaust velocity and high thrust is the Nuclear Salt Water Rocket, and not a few scientist have questions about its feasibility. Well, actually there is also Project Orion, but that has other problems (see below). In science fiction, one often encounters the legendary "fusion drive" or "torchship", which is a high exhaust velocity + high thrust propulsion system that modern science isn't sure is even possible.

With ion engines, chemical engines, and nuclear torches we're facing a classic Newton's Third Law problem. Somehow the exhaust needs to have sufficient momentum for the opposite reaction to give the ship a good acceleration.

Chemical rockets solve the problem by expelling a ton of mass at a relatively low velocity. (high propellant mass flow but low exhaust velocity: SUV)

Ion drives expel a tiny amount of mass, so to get anywhere they get it moving FAST, but even at gigawatts of power they get a measly 0.0001g. (low propellant mass flow but high exhaust velocity: Economy)

Torch drives take a small-to-moderate amount of mass and use nuclear destruction to get it moving insanely fast. (medium propellant mass flow and high exhaust velocity: Torch) They're the only ones (insert disclaimer) with enough power per unit of reaction mass to get 0.3g constant acceleration conveniently. Even a perfect ion drive would need a phenomenal (read: impossible) amount of power input to match the performance of a nuclear explosion.

(A low propellant mass flow and low exhaust velocity engine would be utterly worthless)

From On Torchships comment by Eric (2010)

The Drive Table

All drives listed in the table whose names end in "MAX" require some sort of technological breakthrough to to prevent the engine from vaporizing and/or absurdly large reaction chamber sizes.

If these figures result in disappointing rocket performance, in the name of science fiction you can tweak some of them and claim it was due to a technological advance. You are allowed to tweak anything who's name does not end in "MAX". You can alter the Thrust, Engine Mass, and/or the Eff, but no other values. If there is a corresponding "MAX" entry for the engine you are tweaking, you cannot alter any of the values above the "MAX" entry (i.e., you are not allowed to tweak NTR-SOLID-DUMBO's thrust above 7,000,000, which is the value in the NTR-SOLID MAX entry).

The engines are sorted by thrust power, since that depends on both exhaust velocity and thrust. So engines that high in both of those parameters will be towards the end of the list. This is useful for designers trying to make spacecraft that can both blast-off from a planet's surface and do efficient orbital transfers.

If one was trying to design a more reasonable strictly orbit-to-orbit spacecraft one would want the engine list sorted by exhaust velocity. And surface-to-orbit designers would want the list sorted by thrust.

PropulsionCodeThrust
Power
(MW)
Exhaust
Velocity
(m/s)
Specific
Impulse
(s)
Thrust
(N)
Engine
Mass
(kg)
T/W
DAWN mission NSTARESTAT1.37e-0630,4113,1009.00e-05263.60e-07
ResistojetETHERM7.25e-042,9002961
RadioisotopeNTR SOLID5.85e-037,8007952
ArcJetETHERM2.00e-0220,0002,0392
HOPE Cargo MPDEMAG4.32e-0178,5008,00211
HOPE Tanker MPDEMAG4.32e-0178,5008,00211
HOPE Crew MPDEMAG178,5008,00228
Magneto Inertial Fusion (low)PULSE350,4205,140103
VASIMR (low gear)EMAG629,0002,95640010,0000.004
VASIMR (high gear)EMAG6294,00029,9694010,0004.08e-04
VASIMR (med gear)EMAG6147,00014,9858010,0008.15e-04
Space Shuttle RCSCHEM LIQUID63,1003163,8704106.620
Mirror SteamerBEAM139,8101,0002,60020,9770.013
Monatomic-H MITEENTR SOLID1512,7501,3002,3502001.198
HybridMITEEETHERM1517,6601,8001,70010,0000.017
PropulsionCodeThrust
Power
(MW)
Exhaust
Velocity
(m/s)
Specific
Impulse
(s)
Thrust
(N)
Engine
Mass
(kg)
T/W
AIMPULSE16598,00060,95855
Solar MothBEAM189,0009174,0001004.077
Umbrella ShipESTAT2080,4428,200490
ArcJetETHERM3119,6202,0003,20022,3690.015
Hall EffectESTAT3219,6202,0003,30085,4690.004
Pulsed Plasmoid ThrusterPULSE4378,4808,0001,10083,6110.001
Ponderomotive VASIMREMAG4439,2404,0002,25043,7960.005
Wakefield E-BeamETHERM4519,6202,0004,60041,8370.011
Ablative LaserBEAM4739,2404,0002,40022,2220.011
MPD T-WaveEMAG4778,4808,0001,20082,6750.001
Tungsten ResistojetETHERM499,8101,0009,90042,6010.024
NASA space tugCHEM LIQUID494,40044922,400199,6000.011
Mass DriverOTHER519,8101,00010,400163,0000.007
Ion DriveESTAT5778,4808,0001,444120,1490.001
MET Steamer AmplitronsETHERM599,8101,00012,000123,3020.010
NERVA (CO or N2)NTR SOLID652,64927049,00010,0000.499
Basic MITEENTR SOLID699,8101,00014,0002007.136
NERVA (CO2)NTR SOLID813,30633749,00010,0000.499
NERVA (H2O)NTR SOLID994,04241249,00010,0000.499
HOPE FFRENTR FRAG1115,170,000527,01343
Werka FFRENTR FRAG1115,170,000527,01343113,4003.90e-05
NERVA (NH3)NTR SOLID1255,10152049,00010,0000.499
D-D Fusion InertialPULSE12678,4808,0003,200243,3330.001
NERVA (CH4)NTR SOLID1556,31864449,00010,0000.499
Dusty Plasma (550AU)NTR FRAG16515,000,0001,529,052229,0002.49e-04
PropulsionCodeThrust
Power
(MW)
Exhaust
Velocity
(m/s)
Specific
Impulse
(s)
Thrust
(N)
Engine
Mass
(kg)
T/W
Colloid ThrusterESTAT17243,0004,3838,00020,0000.041
LARSNTR LIQUID19619,6202,00020,0001,0002.039
NERVA (H2)NTR SOLID1988,09382549,00010,0000.499
Laser ThermalBEAM26040,0004,07713,00020,0000.066
Mass DriverOTHER30030,0003,05820,000150,0000.014
LighterCHEM LIQUID3094,410450140,000
LANTR (high gear)NTR SOLID3099,22194067,000
Magneto Inertial Fusion (high)PULSE34850,4205,14013,800
H-B Cat InertialPULSE369156,96016,0004,70065,0890.007
Aluminum/LOX rocketCHEM LIQUID3882,649270292,60056,0000.533
NERVA (H)NTR SOLID39216,0001,63149,00010,0000.499
Kuck MosquitoCHEM LIQUID4844,400449220,000
Vortex Confined (H2)NTR GAS OP49419,6202,00050,400114,1160.045
LH2/LOX rocketCHEM LIQUID5404,905500220,00026,6670.841
VCR Light Bulb (H2)NTR GAS CL55319,6202,00056,40072,5660.079
LANTR (low gear)NTR SOLID5846,347647184,000
Dual-mode Fission (H2)NTR SOLID6129,8101,000124,70033,0000.385
Polywell FusorFUSION65819,6202,00067,10054,0000.127
Cermet NERVA (H2)NTR SOLID6599,8101,000134,40032,5460.421
Z-Pinch MicrofissionPULSE667156,96016,0008,500193,3330.004
Dual-mode PB (H2)NTR SOLID8479,8101,000172,70058,0000.304
Bimodal NTR Solid (NASA)NTR SOLID8988,980915200,0006,6723.056
IonESTAT1,050210,00021,40710,000400,0000.003
D-T FusionFUSION1,18822,0002,243108,00010,0001.101
NERVA Deriv (H2)NTR SOLID1,3508,085824334,06110,1003.372
PropulsionCodeThrust
Power
(MW)
Exhaust
Velocity
(m/s)
Specific
Impulse
(s)
Thrust
(N)
Engine
Mass
(kg)
T/W
Antimatter BottlePULSE1,36278,4808,00034,700180,0000.020
Metastable He*CHEM1,37643,0004,38364,00010,0000.652
Resuable Nuclear ShuttleNTR SOLID1,3768,000815344,000
Metastable HeliumCHEM1,56729,4303,000106,500
n-6Li MicrofissionPULSE1,570156,96016,00020,000106,6670.019
Pebble Bed (H2)NTR SOLID1,5909,530971333,6171,70020.005
Vapor Core (H2)NTR VAPOR1,6179,800999330,0006,8304.925
3He-D Mirror CellFUSION1,664313,92032,00010,600106,6670.010
Cermet (H2)NTR SOLID2,0309,120930445,2679,0005.043
Tokamak MCFUSION2,23166,8006,80966,800197,0000.035
Widmer Mars MissionNTR SOLID2,3208,000815580,000
Antimatter Solid maxAM SOLID2,37410,7911,100440,000
Dusty Plasma (0.5LY)NTR FRAG2,58015,000,0001,529,0523449,0000.004
Proton RD-253 x1CHEM LIQUID2,8363,1003161,830,0001,260148.051
Discovery IIFUSION3,123347,00035,37218,000
MPDEMAG3,140314,00032,00820,0001,540,0000.001
HOPE Z-PinchPULSE3,617189,78019,34638,12095,1380.041
HELIOS 2nd StageNTR SOLID3,8267,800795981,000
NTR Gas/Closed (H2)NTR GAS CL4,54020,4052,080445,00056,8000.799
Dumbo (CO or N2)NTR SOLID4,6362,6492703,500,0005,00071.356
Atomic V-2NTR SOLID4,7148,9809151,050,0004,20025.484
ORION FissionPULSE5,65443,0004,383263,000200,0000.134
Dumbo (CO2)NTR SOLID5,7863,3063373,500,0005,00071.356
THS Fusion Pulse high gearPULSE6,000300,00030,58140,0004,0001.019
THS Fusion Pulse low gearPULSE6,000150,00015,29180,0004,0002.039
PropulsionCodeThrust
Power
(MW)
Exhaust
Velocity
(m/s)
Specific
Impulse
(s)
Thrust
(N)
Engine
Mass
(kg)
T/W
Dumbo (H2O)NTR SOLID7,0744,0424123,500,0005,00071.356
Dumbo (NH3)NTR SOLID8,9275,1015203,500,0005,00071.356
ORION FusionPULSE10,65873,0007,441292,000200,0000.149
Dumbo (CH4)NTR SOLID11,0566,3186443,500,0005,00071.356
Saturn-V F-1 x1CHEM LIQUID11,5412,9823047,740,5009,15386.206
ACMF (ICAN-II)PULSE11,919132,43513,500180,00027,0000.680
Space Shuttle SSME x3CHEM LIQUID12,1154,4444535,452,2009,53158.313
Dumbo (H2)NTR SOLID14,1638,0938253,500,0005,00071.356
Solid rocketCHEM SOLID15,0003,00030610,000,000
Proton RD-253 x6CHEM LIQUID16,2283,10031610,470,0007,560141.174
VISTAPULSE20,400170,00017,329240,000
Project OrionPULSE21,73119,6202,0002,215,200203,6801.109
Nuclear DC-X NERVANTR SOLID27,2729,8101,0005,560,000199,6002.840
Dumbo (H)NTR SOLID28,00016,0001,6313,500,0005,00071.356
Mini-Mag Orion (DRM-3)PULSE29,85393,0009,480642,000199,6000.328
Antimatter Plasma (H2O)AM PLASMA29,890980,00099,89861,000500,0000.012
H-B FusionFUSION29,890980,00099,89861,000300,0000.021
Mini-Mag Orion (DRM-1)PULSE29,90693,1649,497642,000119,0460.550
APCP Space Shuttle SRB x2CHEM SOLID31,2002,60026524,000,0001,180,0002.073
ORION USAF 10mPULSE32,90032,9003,3542,000,000107,9001.889
NTR Solid MAXNTR SOLID42,00012,0001,2237,000,00015,00047.571
Gasdynamic MirrorFUSION46,0601,960,000199,79647,000
Nuclear DC-X LANTRNTR SOLID49,2065,90060116,680,000199,6008.519
NTR Liquid maxNTR GAS OP56,00016,0001,6317,000,00070,00010.194
LunaNTR LIQUID56,65010,3001,05011,000,0009,000124.589
PropulsionCodeThrust
Power
(MW)
Exhaust
Velocity
(m/s)
Specific
Impulse
(s)
Thrust
(N)
Engine
Mass
(kg)
T/W
Saturn-V F-1 x5CHEM LIQUID58,0543,00030638,702,50045,76586.206
NTR Gas/Open (H2)NTR GAS OP61,25035,0003,5683,500,000200,0001.784
TankerNTR GAS OP61,80335,3163,6003,500,000
AV:T high gearPULSE102,138832,92884,906245,250
NTR Gas/Open 2nd GenNTR GAS OP125,00050,0005,0975,000,000200,0002.548
Mini-Mag OrionPULSE146,795157,00016,0041,870,000199,6000.955
NTR Gas MAXNTR GAS OP147,00098,0009,9903,000,00015,00020.387
NTR Gas/Coaxial (H2)NTR GAS OP157,15617,6581,80017,800,000127,00014.287
Antimatter Plasma (H2)AM PLASMA192,0807,840,000799,18549,000500,0000.010
He3-D FusionFUSION192,0807,840,000799,18549,0001,200,0000.004
MC-Fusion MAXFUSION200,0008,000,000815,49450,0006008.495
Salt-water ZubrinNTR GAS OP341,26678,4808,0008,696,900495,4671.789
NSWR (20% UTB)NTR GAS OP425,70066,0006,72812,900,00033,00039.848
Liberty ShipNTR GAS CL560,70030,0003,05837,380,000378,00010.080
Cargo Tug SlingshotESTAT764,400280,00028,5425,460,000
IBS AgamemnonESTAT1,100,000220,00022,42610,000,000
ORION battleshipPULSE1,560,00039,0003,97680,000,0001,700,0004.797
AV:T low gearPULSE2,541,895104,11610,61348,828,125
PropulsionCodeThrust
Power
(MW)
Exhaust
Velocity
(m/s)
Specific
Impulse
(s)
Thrust
(N)
Engine
Mass
(kg)
T/W
ORION 10k ton advPULSE24,000,000120,00012,232400,000,0003,250,00012.546
NSWR (90% UTB) MAXNTR GAS OP30,550,0004,700,000479,10313,000,000
ORION MAXPULSE39,200,0009,800,000998,9818,000,0008,000101.937
Antimatter Beam MAXAM BEAM500,000,000100,000,00010,193,68010,000,00010,000101.937
IC-Fusion MAXPULSE500,000,00010,000,0001,019,368100,000,0001,000,00010.194

Antimatter

Solid Core

AM: Solid
Thrust Power2.4 GW
Exhaust velocity10,791 m/s
Thrust440,000 n
T/W >1.0yes

Basically a NERVA design where a tungsten target replaces the reactor. 13 micrograms per second of antiprotons are annihilated. The gamma rays and pions are captured in the tungsten target, heating it. The tungsten target in turn heats the hydrogen. Produces high thrust but the specific impulse is limited due to material constraints (translation: above a certain power level the blasted tungsten melts)

Gas Core

AM: Gas
Exhaust Velocity24,500 m/s
Specific Impulse2,497 s
FuelAntimatter:
antihydrogen
ReactorLiquid Core
RemassWater
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorMagnetic Nozzle

Microscopic amounts of antimatter are injected into large amounts of water or hydrogen propellant. The intense reaction flashes the propellant into plasma, which exits through the exhaust nozzle. Magnetic fields constrain the charged pions from the reaction so they heat the propellant, but uncharged pions escape and do not contribute any heating. Less efficient than AM-Solid core, but can achieve a higher specific impulse. For complicated reasons, a spacecraft optimized to use an antimatter propulsion system need never to have a mass ratio greater than 4.9, and may be as low as 2. No matter what the required delta V, the spacecraft requires a maximum of 3.9 tons of reaction mass per ton of dry mass, and a variable amount of antimatter measured in micrograms to grams.

Well, actually this is not true if the delta V required approaches the speed of light, but it works for normal interplanetary delta Vs. And the engine has to be able to handle the waste heat.

Plasma Core

AM: Plasma
Water
Exhaust Velocity980,000 m/s
Specific Impulse99,898 s
Thrust61,000 N
Thrust Power29.9 GW
Mass Flow0.06 kg/s
T/W0.01
RemassWater
Specific Power17 kg/MW
Hydrogen
Exhaust Velocity7,840,000 m/s
Specific Impulse799,185 s
Thrust49,000 N
Thrust Power0.2 TW
Mass Flow0.01 kg/s
T/W0.01
RemassLiquid Hydrogen
Specific Power3 kg/MW
Both
Total Engine Mass500,000 kg
FuelAntimatter:
antihydrogen
ReactorPlasma Core
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorMagnetic Nozzle

Similar to antimatter gas core, but more antimatter is used, raising the propellant temperature to levels that convert it into plasma. A magnetic bottle is required to contain the plasma.

Beam Core

AM: Beam
Exhaust Velocity100,000,000 m/s
Specific Impulse10,193,680 s
Thrust10,000,000 N
Thrust Power500.0 TW
Mass Flow0.10 kg/s
Total Engine Mass10,000 kg
T/W102
FuelAntimatter:
antihydrogen
ReactorAntimatter Catalyzed
RemassReaction
Products
Remass AccelAnnihilation
Thrust DirectorMagnetic Nozzle
Specific Power2.00e-05 kg/MW

Microscopic amounts of antimatter are reacted with equal amounts of matter. Remember: unless you are using only electron-positron antimatter annihilation, mixing matter and antimatter does NOT turn them into pure energy. Instead you get some energy, some charged particles, and some uncharged particles.

The charged pions from the reaction are used directly as thrust, instead of being used to heat a propellant. A magnetic nozzle channels them. Without a technological break-through, this is a very low thrust propulsion system.

All antimatter rockets produce dangerous amounts of gamma rays. The gamma rays and the pions can transmute engine components into radioactive isotopes. The higher the mass of the element transmuted, the longer lived it is as a radioisotope.

Positron Ablative

Positron Ablative
Exhaust velocity49,000 m/s

This engine produces thrust when thin layers of material in the nozzle are vaporized by positrons in tiny capsules surrounded by lead. The capsules are shot into the nozzle compartment many times per second. Once in the nozzle compartment, the positrons are allowed to interact with the capsule, releasing gamma rays. The lead absorbs the gamma rays and radiates lower-energy X-rays, which vaporize the nozzle material. This complication is necessary because X-rays are more efficiently absorbed by the nozzle material than gamma rays would be.

Drawbacks include the fact that you need 1836 positrons to equal the energy of a single anti-proton, and only half the positrons will hit the pusher plate limiting the efficiency to 50%.

This system is very similar to Antiproton-catalyzed microfission

Beamed Power

Laser Thermal

Laser Thermal
Exhaust Velocity40,000 m/s
Specific Impulse4,077 s
Thrust13,000 N
Thrust Power0.3 GW
Mass Flow0.33 kg/s
Total Engine Mass20,000 kg
T/W0.07
Thermal eff.30%
Total eff.30%
FuelExternal
Laser
ReactorCollector Mirror
RemassSeeded Hydrogen
Remass AccelThermal Accel:
Collector Mirror
Thrust DirectorNozzle
Specific Power77 kg/MW

Similar to Solar Moth, but uses a stationary ground or space-station based laser instead of the sun. Basically the propulsion system leaves the power plant at home and relies upon a laser beam instead of an incredibly long extension cord.

As a rule of thumb, the collector mirror of a laser thermal rocket can be much smaller than a comparable solar moth, since the laser beam probably has a higher energy density than natural sunlight.

With the mass of the power plant not actually on the spacecraft, more mass is available for payload. Or the reduced mass makes for a higher mass ratio to increase the spacecraft's delta V. The reduced mass also increases the acceleration. In some science fiction novels, combat "motherships" will have batteries of lasers, used to power hordes of ultra-high acceleration missiles and/or fighter spacecraft.

The drawback include the fact that there is a maximum effective range you can send a worthwhile laser beam from station to spacecraft, and the fact that the spacecraft is at the mercy of whoever is controlling the laser station.

Propellant is hydrogen seeded with alkali metal. As always the reason for seeding is that hydrogen is more or less transparent so the laser beam will mostly pass right through without heating the hydrogen. The seeding make the hydrogen more opaque so the blasted stuff will heat up. Having said that, the Mirror Steamer has an alternate solution.


The equations for delta V and mass ratio are slightly different for a Solar Moth or Laser Thermal rocket engine:

Δv = sqrt((2 * Bp * Bε) / mDot) * ln[R]

R = ev/sqrt((2 * Bp * Bε) / mDot)

where

  • Δv = ship's total deltaV capability (m/s)
  • R = ship's mass ratio
  • Bp = Beam power (watts) of either laser beam or solar energy collected
  • = efficiency with which engine converts beam power into exhaust kinetic energy (0.0 to 1.0, currently about 0.3)
  • ln[x] = natural logarithm of x, the "ln" key on your calculator
  • ex = antilog base e or inverse of natural logarithm of x, the "ex" key on your calculator
Ablative Laser
Ablative Laser
Exhaust Velocity39,240 m/s
Specific Impulse4,000 s
Thrust2,400 N
Thrust Power47.1 MW
Mass Flow0.06 kg/s
Total Engine Mass22,222 kg
T/W0.01
Frozen Flow eff.88%
Thermal eff.90%
Total eff.79%
FuelExternal
Laser
ReactorCollector Mirror
RemassGraphite
Remass AccelThermal Accel:
Collector Mirror
Thrust DirectorNozzle
Specific Power472 kg/MW

A rocket can be driven by high-energy, short-duration (<10-10 sec) laser pulses, focused on a solid propellant.

A double-pulse system is used: the first pulse ablates material and the second further heats the ablated gas. A low Z propellant, such as graphite, obtains the best specific impulse (4 ksec). Unfortunately, ice is not a suitable medium due to melting and “dribbling” losses.

Primary and secondary mirrors focus the pulses at irradiances of 3 × 1013 W/cm2. The mass-removal rate is 3 μg per laser pulse. Powered with a 60 MW beam, an ablative laser thruster has a thrust of 2.4 kN and, with a fuel tuned to the firing sequences and an efficient double-pulsed shape, the overall efficiency is 80%.

“Specific impulse and other characteristics of elementary propellants for ablative laser propulsion”, Dr. Andrew V. Pakhomov, Associate Professor at the Department of Physics, UAH, 2002.

From High Frontier by Philip Eklund

Laser Sail

A Laser Sail is a photon sail beam-powered by a remote laser installation.

As an important point, the practical minimum acceleration for a spacecraft is about 5 milligees. Otherwise it will take years to change orbits. Photo sails can only do up to 3 milligees, but a laser sail can do 5 milligees easily.

Solar Moth

Solar Moth
Exhaust Velocity9,000 m/s
Specific Impulse917 s
Thrust4,000 N
Thrust Power18.0 MW
Mass Flow0.44 kg/s
Total Engine Mass100 kg
T/W4
Thermal eff.65%
Total eff.65%
FuelSolar Photons
ReactorCollector Mirror
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Collector Mirror
Thrust DirectorNozzle
Specific Power6 kg/MW

Solar thermal rocket. 175 meter diameter aluminum coated reflector concentrates solar radiation onto a window chamber hoop boiler, heating and expanding the propellant through a regeneratively-cooled hoop nozzle. The concentrating mirror is one half of a giant inflatable balloon, the other half is transparent (so it has an attractive low mass).

The advantage is that you have power as long as the sun shines and your power plant has zero mass (as far as the spacecraft mass is concerned). The disadvantage is it doesn't work well past the orbit of Mars. The figures in the table are for Earth orbit.

The solar moth might be carried on a spacecraft as an emergency propulsion system, since the engine mass is so miniscule.


The equations for delta V and mass ratio are slightly different for a Solar Moth or Laser Thermal rocket engine:

Δv = sqrt((2 * Bp * Bε) / mDot) * ln[R]

R = ev/sqrt((2 * Bp * Bε) / mDot)

where

  • Δv = ship's total deltaV capability (m/s)
  • R = ship's mass ratio
  • Bp = Beam power (watts) of either laser beam or solar energy collected
  • = efficiency with which engine converts beam power into exhaust kinetic energy (0.0 to 1.0)
  • ln[x] = natural logarithm of x, the "ln" key on your calculator
  • ex = antilog base e or inverse of natural logarithm of x, the "ex" key on your calculator

For the Solar Moth in the data block Bε = 0.63, for the Mirror Steamer Bε = 0.87

Bp = Marea * (☉constant * (1 / (☉dist2)))

where

  • Bp = Beam power (watts) of solar energy collected
  • Marea = total area of collecting mirrors (m2)
  • dist = distance between Sun and spacecraft (Astronomical Units, Earth = 1.0)
  • constant = Solar Constant (w/m2)

1.0 astronomical units is defined as 149,597,870,700 meters.

1 / (☉dist2) is the sunlight energy density. In Earth's orbit, the density is 1.0, at Mars orbit it is 0.44 (44%), at Jupiter orbit it is 0.037, at Neptune orbit it is 0.001, at Mercury orbit it is 6.68

The Solar Constant varies from 1,361 w/m2 at solar minimum and 1,362 w/m2 at solar maximum.

Mirror Steamer Robonaut
Mirror Steamer
Exhaust Velocity9,810 m/s
Specific Impulse1,000 s
Thrust2,600 N
Thrust Power12.8 MW
Mass Flow0.27 kg/s
Total Engine Mass20,977 kg
T/W0.01
Frozen Flow eff.97%
Thermal eff.90%
Total eff.87%
FuelSolar Photons
ReactorCollector Mirror
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Collector Mirror
Thrust DirectorNozzle
Specific Power1,645 kg/MW

Water is an attractive volumetric absorber for infrared laser propulsion. Diatomic species formed from the disassociation of water such as OH are present at temperatures as high as 5000 K, and can be rotationally excited by a free electron laser operating in the far infrared. The OH molecules then transfer their energy to a stream of hydrogen propellant in a thermodynamic rocket nozzle by relaxation collisions.

Beamed heat can also be added by a blackbody cavity absorber. This heat exchanger is a series of concentric cylinders, made of hafnium carbide (HfC). Focused sunlight or lasers passes through the outermost porous disk, and is absorbed in the cavity. Heat is transferred to the propellant by the hot HfC without the need for propellant seeding. The specific impulse is materials-limited to 1 ks.

“Solar Rocket System Concept Analysis”, F.G. Etheridge, Rockwell Space Systems Group. (I resized the Rockwell “Solar Moth” design for 3 kN thrust).

From High Frontier by Philip Eklund.

Chemical

A barely contained chemical explosive. Noted for very high thrust and very low exhaust velocity. One of the few propulsion systems where the fuel and the propellant are the same thing. There is a list of chemical propellants here

Solid Rocket

Space Shuttle SRB x2
Exhaust Velocity2,600 m/s
Specific Impulse265 s
Thrust/Engine12,000,000 N
Number Thrustersx2
Thrust24,000,000 N
Thrust Power31.2 GW
Mass Flow9,231 kg/s
Total Engine Mass1,180,000 kg
T/W2
FuelChemical Solid:
APCP
ReactorCombustion
Chamber
RemassReaction
Products
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power38 kg/MW

Liquid Rocket

Methane-Oxygen

Chemical: Methane-Oxygen
Exhaust Velocity3,700 m/s
Specific Impulse377 s

Methane and oxygen are burned resulting in an unremarkable specific impulse of about 377 seconds. However, this is the highest performance of any chemical rocket using fuels that can be stored indefinitely in space. Chemical rockets with superior specific impulse generally use liquid hydrogen, which will eventually leak away by escaping between the the molecules composing your fuel tanks. Liquid methane and liquid oxygen will stay put.

ISRU Sabatier

The Sabatier reactor uses In-Situ Resource Utilization (ISRU) to create a closed hydrogen and oxygen cycle for life support on planets with CO2 atmospheres such as Mars or Venus.

It contains two chambers, one for mixing and the other for storing a nickel catalyst. When charged with hydrogen and atmospheric carbon dioxide, it produces water and methane. (The similar Bosch reactor uses an iron catalyst to produce elemental carbon and water.)

A condenser separates the water vapor from the reaction products. This condenser is a simple pipe with outlets on the bottom to collect water; natural convection on the surface of the pipe is enough to carry out the necessary heat exchange.

Electrolysis of the water recovers the hydrogen for reuse.

NASA 2007.

From High Frontier by Philip Eklund

Hydrogen-Fluorine

Chemical: LH2/Fluorine
Exhaust Velocity4,700 m/s
Specific Impulse479 s

Hydrogen-Oxygen

Chemical: LH2/LOX
Exhaust Velocity4,400 m/s
Specific Impulse449 s
Space Shuttle SSME x3
Propulsion SystemChemical: LH2/LOX
Exhaust Velocity4,444 m/s
Specific Impulse453 s
Thrust/Engine1,817,400 N
Number Thrustersx3
Thrust5,452,200 N
Thrust Power12.1 GW
Mass Flow1,227 kg/s
Total Engine Mass9,531 kg
T/W58
Specific Power1 kg/MW
NASA space tug
Propulsion SystemChemical: LH2/LOX
Thrust22,400 N
Thrust Power49.3 MW
Mass Flow5 kg/s
Total Engine Mass199,600 kg
T/W0.01
Wet Mass32,000 kg
Dry Mass14,000 kg
Mass Ratio2.29 m/s
ΔV3,637 m/s
Specific Power4,050 kg/MW
Lighter
Propulsion SystemChemical: LH2/LOX
Exhaust Velocity4,410 m/s
Specific Impulse450 s
Thrust140,000 N
Thrust Power0.3 GW
Mass Flow32 kg/s
Wet Mass56,300 kg
Dry Mass25,898 kg
Mass Ratio2.17 m/s
ΔV3,424 m/s
Kuck Mosquito
Propulsion SystemChemical: LH2/LOX
Exhaust Velocity4,400 m/s
Specific Impulse449 s
Thrust220,000 N
Thrust Power0.5 GW
Mass Flow50 kg/s
Wet Mass350,000 kg
Dry Mass100,000 kg
Mass Ratio3.50 m/s
ΔV5,512 m/s

Hydrogen and oxygen are burned resulting in close to the theoretical maximum specific impulse of about 450 seconds. However, liquid hydrogen cannot be stored permanently in any tank composed of matter. The blasted stuff will escape atom by atom between the molecules composing the fuel tanks.

LH2/LOX Rocket
LH2/LOX Rocket
Exhaust Velocity4,905 m/s
Specific Impulse500 s
Thrust220,000 N
Thrust Power0.5 GW
Mass Flow45 kg/s
Total Engine Mass26,667 kg
T/W0.84
Frozen Flow eff.55%
Thermal eff.98%
Total eff.54%
FuelChemical:
Single-H/LOX
ReactorCombustion
Chamber
RemassReaction
Products
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power49 kg/MW

The combustion of the cryogenic fuels hydrogen and oxygen produces an ideal specific impulse of 528 seconds. The product is water, which is exhausted through a converging-diverging tube called a De Laval nozzle.

The engine illustrated is similar to the Space Shuttle main engine, with a specific impulse of 460 seconds. The De Laval nozzle has a 180:1 area ratio, and is regeneratively-cooled with liquid hydrogen. The chamber temperature is 3500K, and the chamber pressure is 2.8 MPa. The engine has a thermal efficiency of 98%, a mixture ratio of 5.4, and a frozen-flow efficiency of 55%. A 2000 MWth chamber generates 440 kN of thrust and a thrust to weight ratio of one gravity.

Space Transportation Systems, American Institute of Aeronautics and Astronautics, 1978.

From High Frontier by Philip Eklund

RP-1 - Oxygen

Chemical: RP-1/LOX
Exhaust Velocity3,500 m/s
Specific Impulse357 s
Saturn-V F-1 x1
Exhaust Velocity2,982 m/s
Specific Impulse304 s
Thrust7,740,500 N
Thrust Power11.5 GW
Mass Flow2,596 kg/s
Total Engine Mass9,153 kg
T/W86
FuelChemical:
RP-1/LOX
ReactorCombustion
Chamber
RemassReaction
Products
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power1 kg/MW
Saturn-V F-1 x5
Exhaust Velocity3,000 m/s
Specific Impulse306 s
Thrust/Engine7,740,500 N
Number Thrustersx5
Thrust38,702,500 N
Thrust Power58.1 GW
Mass Flow12,901 kg/s
Total Engine Mass45,765 kg
T/W86
FuelChemical Liquid:
RP-1/LOX
Specific Power1 kg/MW

RP-1 is Rocket Propellant-1 or Refined Petroleum-1) is a highly refined form of kerosene outwardly similar to jet fuel, used as rocket fuel. It is not as powerful as liquid hydrogen but it is a whole lot less trouble. Compared to LH2 it is cheaper, stabler at room temperature, non-cryogenic less of an explosive hazard, and denser.

NASA uses it a lot.

Hypergolic Fuels

Chemical: UDMH/N204
Exhaust Velocity3,267 m/s
Specific Impulse333 s
Chemical: MMH/N204
Exhaust Velocity3,296 m/s
Specific Impulse336 s
Space Shuttle RCS
Thrust3,870 N
Thrust Power6.0 MW
Mass Flow1 kg/s
Total Engine Mass4 kg
T/W107
FuelChemical:
MMH/N204
Specific Power1 kg/MW
Proton RD-253 x1
Thrust1,830,000 N
Thrust Power2.8 GW
Mass Flow590 kg/s
Total Engine Mass1,260 kg
T/W148
FuelChemical:
UDMH/N204
Specific Power0.44 kg/MW
Proton RD-253 x6
Thrust/Engine1,745,000 N
Number Thrustersx6
Thrust10,470,000 N
Thrust Power16.2 GW
Mass Flow3,377 kg/s
Total Engine Mass7,560 kg
T/W141
FuelChemical:
UDMH/N204
Specific Power0.47 kg/MW

Unsymmetrical dimethylhydrazine (UDMH) + nitrogen tetroxide (N204 or "NTO") and Monomethylhydrazine (MMH) + NTO are very important chemical rocket fuels.

Both are hypergolic, meaning the stuff explodes on contact with each other instead of needing a pilot light or other ignition system as do other chemical fuels. This means one less point of failure and one less maintenance nightmare on your spacecraft. Being hypergolic also prevents large amounts of fuel and oxidizer accumulating in the nozzle, which can cause a hard start or engine catastrophic failure (fancy term for "engine goes ka-blam!"). It is also non-cryogenic, liquid at room temperature and pressure. This means it is a storable liquid propellant, suitable for space missions that last years.

"Ah, what's the catch?" you ask.

The catch is that the mix is hideously corrosive, toxic, and carcinogenic. It is also easily absorbed through the skin. If UDMH escapes into the air it reacts to form dimethylnitrosamine, which is a persistent carcinogen and groundwater pollutant. MMH is only fractionally less bad.

This is the reason for all those technicians wearing hazmat suits at Space Shuttle landings. The Shuttle used MMH/NTO in its reaction control thrusters. Upon landing the techs had to drain the hellish stuff before it leaked and dissoved some innocent bystander.

Hybrid Rocket

Aluminum-Oxygen

Chemical: Aluminum-Oxygen
Exhaust Velocity2,800 m/s
Specific Impulse285 s
FuelChemical:
Aluminum/LOX
ReactorCombustion
Chamber
RemassReaction
Products
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle

Aluminum and oxygen are burned resulting in an unremarkable specific impulse of about 285 seconds. However, this is of great interest to any future lunar colonies. Both aluminum and oxygen are readily available in the lunar regolith, and such a rocket could easily perform lunar liftoff, lunar landing, or departure from a hypothetical L5 colony for Terra (using a lunar swingby trajectory). The low specific impulse is more than made up for by the fact that the fuel does not have to be imported from Terra. It can be used in a hybrid rocket (with solid aluminum burning in liquid oxygen), or using ALICE (which is a slurry of nanoaluminium powder mixed in water then frozen).

Of course the aluminum oxide in lunar regolith has to be split into aluminum and oxygen before you can use it as fuel. But Luna has plenty of solar power. As a rule of thumb, in space, energy is cheap but matter is expensive.

Aluminum/LOX rocket
Aluminum/LOX rocket
Exhaust Velocity2,649 m/s
Specific Impulse270 s
Thrust292,600 N
Thrust Power0.4 GW
Mass Flow110 kg/s
Total Engine Mass56,000 kg
T/W0.53
Frozen Flow eff.79%
Thermal eff.98%
Total eff.77%
FuelChemical:
APCP
ReactorCombustion
Chamber
RemassReaction
Products
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power145 kg/MW

Solar Carbothermal Refinery

Although aluminum is common in space, it stubbornly resists refining from its oxide Al3O2. It can be reduced by a solar carbothermal process, using carbon as the reducing agent and solar energy. Compared to carbo-chlorination, this process needs no chlorine, which is hard to obtain in space. Furthermore, the use of solar heat instead of electrolysis allows higher efficiency and less power conditioning. The solar energy required is 0.121 GJ/kg Al.

The aluminum and oxygen produced can be used to fuel Al-O2 chemical boosters, which burn fine sintered aluminum dust in the presence of liquid oxygen (LO2). Unlike pure solid rockets, hybrid rockets (using a solid fuel and liquid oxidizer) can be throttled and restarted. The combustion of aluminum obtains 3.6 million joules per kilogram. At 77% propulsion efficiency, the thrust is 290 kN with a specific impulse of 285 seconds. The mass ratio for boosting off or onto Luna using an Al-O2 rocket is 2.3. In other words, over twice as much as much fuel as payload is needed.

Gustafson, White, and Fidler of ORBITECTM, 2010.


Carbochlorination Refinery

Metal sulfates may be refined by exposing a mixture of the crushed ore and carbon dust to streams of chlorine gas. Under moderate resistojet heating (1123 K) in titanium chambers (Ti resists attack by Cl), the material is converted to chloride salts such as found in seawater, which can be extracted by electrolysis.

The example shown is the carbochlorination of Al2Cl3 to form aluminum. Al is valuable in space for making wires and cables (copper is rare in space). The electrolysis of Al2Cl3 does not consume the electrodes nor does it require cryolite. However, due to the low boiling point of Al2Cl3, the reaction must proceed under pressure and low temperatures.

Other elements produced by carbochlorination include titanium, potassium, manganese, chromium, sodium, magnesium, silicon and also (with the use of plastic filters) the nuclear fuels 235U and 232Th. Both C and Cl2 must be carefully recycled (the recycling equipment dominates the system mass) and replenished by regolith scavenging.

Dave Dietzler

From High Frontier by Philip Eklund

ISRU Metal-Oxygen

Propulsion Fuels From Indigenous Lunar And Asteroidal Metals
Table 1: Metal/Oxygen Combustion Properties
MetalSpecific
Enthalpy
(joules/kg)
Isp
(seconds)
hydrogen1.39×107457
aluminum1.63×107270
calcium1.41×107213
iron4.7×106184
magnesium1.83×107260
silicon1.58×107272
titanium1.17×107255

Lunar and asteroidal surface materials are ubiquitous and abundant sources of metals like silicon, aluminum, magnesium, iron, calcium, and titanium. Many schemes have been proposed for extracting these metals and oxygen for structural, electrical, and materials processing space operations.

However, all the metals burn energetically in oxygen and could serve as in-situ rocket fuels for space transportation applications.

Table 1 lists the specific heats of combustion (enthalpy) at 1800 K and corresponding specific impluses at selected mixture ratios with oxygen of the above pure metals assuming rocket combustion at 1000 psia and an expansion ratio of 50. Hydrogen is included for comparison.

All the metals appear to offer adequate propulsion performance from low or moderate gravity bodies and are far more abundant than hydrogen on many terrestrial planets and asteroids.

It is noteworthy that silicon, the most abundant nonterrestrial metal, is potentially one of the best performers. In addition, iron with the lowest specific impulse is sufficiently energetic for cislunar and asteroidal transportation. Further, silicon and iron are the most readily obtained nonterrestrial metals. They can be separated by distillation of basalts and other nonterrestrial silicates in vacuum solar furnaces.

Efficient rocket combustion of metal fuels could be realized by injecting them as a fine powder into the combustion chamber. This could be done by mixing the fuel with an inert carrier gas or in liquid oxygen (LOX) to form a slurry. Preliminary studies indicate that a mixture of metal/LOX can be stored and handled safely without danger of autoignition. Lean fuel mixtures would be used to achieve the maximum specific impluse by reducing the exhaust molecular weight without excessivly lowering the combustion temperature. Two phase flow losses are estimated to be acceptable for anticipated throat sizes based on measured thrust loss data from solid rocket motors ustng aluminized propellants.

The metals could be atomized by condensing droplets in vacuum from a liquid metal stream forced through a fine ceramic nozzle. Brittle metals like silicon and calcium might be pulverized to sub 20 micrometer size in vacuum in autogenous grinders that operate by centrifugal impact and are independent of the gravity level.

From Propulsion Fuels From Indigenous Lunar And Asteroidal Metals by William N. Agosto and John H. Wickman

Metastable

Atomic Hydrogen

100% Atomic Hydrogen
Exhaust velocity20,600 m/s
15% Atomic Hydrogen in solid H2
Exhaust velocity7,300 m/s
Single-H/LOX
Exhaust Velocity4,600 m/s
Specific Impulse469 s

Atomic hydrogen is also called free-radical hydrogen or "single-H". The problem is that it instantly wants to recombine. The least unreasonable way of preventing this is to make a solid mass of frozen hydrogen (H2) at liquid helium temperatures which contains 15% single-H by weight.

Free Radical Hydrogen
Free Radical Hydrogen
Exhaust velocity39,240 m/s
Thrust73,900 N
Specific Power55 kg/MW
Engine Power2,000 MW
Frozen Flow eff.77%
Thermal eff.94%
Thrust Power1448 MW

Free radicals are single atoms of elements that normally form molecules. Free radical hydrogen (H) has half the molecular weight of H2.

If used as propellant, it doubles the specific impulse of thermodynamic rockets.

If used as fuel, its specific energy (218 MJ/kg) produces a theoretical specific impulse of 2.13 ksec.

Free radicals extracted by particle bombardment are cooled by VUV laser chirping, and trapped in a hybrid laser-magnet as a Bose-Einstein gas at ultracold temperatures. A Pritchard-Ioffe trap keeps their mobile spins aligned, using the interaction of the atomic magnetic moment with the inhomogeous magnetic field. The trapping density of >1014 atoms/cc is much higher than in Penning traps.

Free radical deuterium that has been spin-vector polarized is stable against ionization and atomic collisions. Because of its large fusion reactivity cross-sectional area, it makes a useful fusion fuel.

Robert L. Forward, 1983

From High Frontier by Philip Eklund

Metallic Hydrogen

Metallic Hydrogen
Exhaust velocity17,000 m/s

Hydrogen subjected to enough pressure to turn it into metal, then contained under such pressure. Release the pressure and out comes all the stored energy that was required to compress it in the first place. It will require storage that can handle millions of atmospheres worth of pressure. The mass of the storage unit might be enough to negate the advantage of the high exhaust velocity.

The hope is that somebody might figure out how to compress the stuff into metal, then somehow release the pressure and have it stay metallic.

Metastable He*

Metastable He*
Exhaust Velocity43,000 m/s
Specific Impulse4,383 s
Thrust64,000 N
Thrust Power1.4 GW
Mass Flow1 kg/s
Total Engine Mass10,000 kg
T/W0.65
FuelMetastable He*
ReactorCombustion
Chamber
RemassReaction
Products
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Specific Power7 kg/MW

Spin-polarized triplet helium. Two electrons in a helium atom are aligned in a metastable state (one electron each in the 1s and 2s atomic orbitals with both electrons having parallel spins, the so-called "triplet spin state", if you want the details). When it reverts to normal state it releases 0.48 gigjoules per kilogram. Making the stuff is easy. The trouble is that it tends to decay spontaneously, with a lifetime of a mere 2.3 hours. And it will decay even quicker if something bangs on the fuel tank. Or if the ship is jostled by hostile weapons fire. To say the fuel is touchy is putting it mildly. The fuel is stored in a resonant waveguide to magnetically lock the atoms in their metastable state but that doesn't help much. There were some experiments to stablize it with circularly polarized light, but I have not found any results about that.

Metastable He IV-A

Metastable He IV-A
Exhaust Velocity21,600 m/s
Specific Impulse2,202 s
Total Engine Mass10,000 kg
FuelMetastable He IV-A
ReactorCombustion
Chamber
RemassReaction
Products
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle
Meta from Saturn Rukh
Exhaust Velocity30,900 m/s
Specific Impulse3,150 s

Meta-helium would be such a worthwhile propulsion system that scientists have been trying real hard to get the stuff to stop decaying after a miserable 2.3 hours. One approach is to see if metastable helium can be formed into a room-temperature solid if bonded with diatomic helium molecules, made from one ground state atom and one excited state atom. This is called diatomic metastable helium. The solid should be stable, and it can be ignited by heating it. The exhaust velocity is about half that of pure He* which is disappointing, but not as disappointing as a dust-mote sized meteorite blowing your ship into atoms.

Theoretically He IV-A would be stable for 8 years, have a density of 0.3 g/cm3, and be a solid with a melting point of 600 K (27° C). The density is a plus, liquid hydrogen's annoying low density causes all sorts of problems.


Dr. Robert Forward in his novel Saturn Rukh suggested bonding 64 metastable helium atoms to a single excited nitrogen atom, forming a stable super-molecule called Meta. Whether or not this is actually possible is anybody's guess. In theory it would have a specific impulse of 3150 seconds.

Metastable Helium
Metastable Helium
Exhaust Velocity29,430 m/s
Specific Impulse3,000 s
Thrust106,500 N
Thrust Power1.6 GW
Mass Flow4 kg/s
Frozen Flow eff.90%
Thermal eff.87%
Total eff.78%
FuelMetastable He IV-A
ReactorCombustion
Chamber
RemassReaction
Products
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorNozzle

Metastable helium is the electronically excited state of the helium atom, easily formed by a 24 keV electron beam in liquid helium.

If the spin-orbit decay is suppressed by a coherent laser pump, its theoretical lifetime would be eight years (as ferromagnetic solid He*2 with a melting temperature of 600 K). Spin-aligned solid metastable helium could be a useful, if touchy, high thrust chemical fuel with a theoretical specific impulse of 3.2 ksec.

J.S. Zmuidzinas, "Stabilization of He2(a 3Sigmau+) in Liquid Helium by Optical Pumping," unpublished 1976, courtesy Dr. Robert Forward.

From High Frontier by Philip Eklund

Electromagnetic (Plasma)

Electrodeless plasma

Helicon Double Layer (HDLT)

Magnetoplasmadynamic (MPD)

Magnetoplasmadynamic
Exhaust Velocity314,000 m/s
Specific Impulse32,008 s
Thrust20,000 N
Thrust Power3.1 GW
Mass Flow0.06 kg/s
Total Engine Mass1,540,000 kg
T/W1.00e-03
Thermal eff.79%
Total eff.79%
Fuel4GWe input
RemassHelium
Remass AccelElectromagnetic
Acceleration
Specific Power490 kg/MW
HOPE Cargo MPD
Propulsion SystemMPD
Exhaust Velocity78,500 m/s
Specific Impulse8,002 s
Thrust11 N
Number Thrusters2
Thrust Power0.4 MW
Mass Flow1.40e-04 kg/s
Fuel60MWe input
RemassHelium
Wet Mass242,000 kg
Dry Mass182,000 kg
Mass Ratio1.33 m/s
ΔV22,367 m/s
HOPE Tanker MPD
Propulsion SystemMPD
Exhaust Velocity78,500 m/s
Specific Impulse8,002 s
Thrust11 N
Number Thrusters2
Thrust Power0.4 MW
Mass Flow1.40e-04 kg/s
Fuel60MWe input
RemassHelium
Wet Mass244,000 kg
Dry Mass184,000 kg
Mass Ratio1.33 m/s
ΔV22,155 m/s
HOPE Crew MPD
Propulsion SystemMPD
Exhaust Velocity78,500 m/s
Specific Impulse8,002 s
Thrust28 N
Number Thrusters4
Thrust Power1.1 MW
Mass Flow3.57e-04 kg/s
Fuel60MWe input
RemassHelium
Wet Mass262,000 kg
Dry Mass188,000 kg
Mass Ratio1.39 m/s
ΔV26,054 m/s

Magnetoplasmadynamic thruster, a travelling wave plasma accelerator. Propellant is potassium seeded helium.

MPD T-Wave
MPD T-Wave
Exhaust Velocity78,480 m/s
Specific Impulse8,000 s
Thrust1,200 N
Thrust Power47.1 MW
Mass Flow0.02 kg/s
Total Engine Mass82,675 kg
T/W1.00e-03
Frozen Flow eff.90%
Thermal eff.81%
Total eff.73%
Fuel60MWe input
RemassRegolith
Remass AccelElectromagnetic
Acceleration
Specific Power1,756 kg/MW

Impulsive electric rockets can accelerate propellant using magnetoplasmadynamic traveling waves (MPD T-waves).

In the design shown, superfluid magnetic helium-3 is accelerated using a megahertz pulsed system, in which a few hundred kiloamps of currents briefly develop extremely high electromagnetic forces. The accelerator sequentially trips a column of distributed superconducting L-C circuits that shoves out the fluid with a magnetic piston.

The propellant is micrograms of regolith dust entrained by the superfluid helium. The dust and helium are kept from the walls by the inward radial Lorentz force, with an efficiency of 81%.

Each 125 J pulse requires a millifarad of total capacitance at a few hundred volts. Compared to ion drives, MPDs have good thrust densities and have no need for charge neutralization. However, they run hot and have electrodes that will erode over time. Moreover, small amounts of an expensive superfluid medium are continually required.

From High Frontier by Philip Eklund

Pulsed Inductive (PIT)

Pulsed inductive thruster

Pulsed Plasma (PPT)

Pulsed plasma thruster

Pulsed Plasmoid Thruster
Pulsed Plasmoid Thruster
Exhaust Velocity78,480 m/s
Specific Impulse8,000 s
Thrust1,100 N
Thrust Power43.2 MW
Mass Flow0.01 kg/s
Total Engine Mass83,611 kg
T/W1.00e-03
Frozen Flow eff.90%
Thermal eff.80%
Total eff.72%
Fuel60MWe input
RemassRegolith
Remass AccelThermal Accel:
Reaction Heat
Thrust DirectorMagnetic Nozzle
Specific Power1,937 kg/MW

A plasmoid is a coherent torus-shaped structure of plasma and magnetic fields.

An example from nature is “Kugelblitz” (ball lightning). (One of my mentors, Dr. Roger C. Jones of the University of Arizona, has worked out the physics of this.)

A plasmoid rocket creates a torus of ball lightning by directing a mega-amp of current onto the propellant. Almost any sort of propellant will work. The plasmoid is expanded down a diverging electrically conducting nozzle. Magnetic and thermal energies are converted to directed kinetic energy by the interaction of the plasmoid with the image currents it generates in the nozzle. Ionization losses are a small fraction of the total energy; the frozen flow efficiency is 90%.

Unlike other electric rockets, a plasmoid thruster requires no electrodes (which are susceptible to erosion) and its power can be scaled up simply by increasing the pulse rate.

The design illustrated has a 50-meter diameter structure that does quadruple duty as a nozzle, laser focuser, high gain antenna, and radiator. Laser power (60 MW) is directed onto gap photovoltaics to charge the ultracapacitor bank used to generate the drive pulses.

R. Bourque, General Atomics, 1990.

From High Frontier by Philip Eklund

VASIMR

VASIMR
VASIMR (high gear)
Exhaust Velocity294,000 m/s
Specific Impulse29,969 s
Thrust40 N
Thrust Power5.9 MW
Mass Flow1.36e-04 kg/s
Total Engine Mass10,000 kg
T/W4.08e-04
Specific Power1,701 kg/MW
VASIMR (med gear)
Exhaust Velocity147,000 m/s
Specific Impulse14,985 s
Thrust80 N
Thrust Power5.9 MW
Mass Flow5.44e-04 kg/s
Total Engine Mass10,000 kg
T/W8.15e-04
Specific Power1,701 kg/MW
VASIMR (low gear)
Exhaust Velocity29,000 m/s
Specific Impulse2,956 s
Thrust400 N
Thrust Power5.8 MW
Mass Flow0.01 kg/s
Total Engine Mass10,000 kg
T/W4.08e-03
Specific Power1,724 kg/MW
All
Thermal eff.60%
Total eff.60%
Fuel19.6MWe input
RemassLiquid Hydrogen
Remass AccelElectromagnetic
Acceleration
Thrust DirectorMagnetic Nozzle

Some classify this as an electromagnetic plasma, some as an electrodeless electrothermal

The variable specific impulse magnetoplasma rocket is a plasma drive with the amusing ability to "shift gears." This means it can trade exhaust velocity for thrust and vice versa. Three "gears" are shown on the table. There are more details here and here.

VASIMR has been suggested for use in a space tug aka Orbital Transfer Vehicle. A VASIMR powered tug could move 34 metric tons from Low Earth Orbit (LEO) to Low Lunar Orbit (LLO) by expending only 8 metric tons of argon propellant. A chemical rocket tug would require 60 metric tons of liquid oxygen - liquid hydrogen propellant. Granted the VASIMR tug would take six month transit time as opposed to the three days for the chemical, but there are always trade offs.

Ponderomotive VASIMR
Ponderomotive VASIMR
Exhaust Velocity39,240 m/s
Specific Impulse4,000 s
Thrust2,250 N
Number Thrusters15
Thrust/Engine150 N
Thrust Power44.1 MW
Mass Flow0.06 kg/s
Total Engine Mass43,796 kg
T/W5.00e-03
Frozen Flow eff.90%
Thermal eff.80%
Total eff.72%
Fuel60MWe input
RemassLiquid Hydrogen
Remass AccelElectromagnetic
Acceleration
Thrust DirectorMagnetic Nozzle
Specific Power992 kg/MW

The variable-specific-impulse magnetoplasma rocket (VASIMR) has two unique features: the removal of the anode and cathode electrodes (which greatly increases its lifetime compared to other electric rockets) and the ability to throttle the engine, exchanging thrust for specific impulse. A VASIMR uses low gear to climb out of planetary orbit, and high gear for interplanetary cruise.

Other advantages include efficient resonance heating (80%), and a low current, high voltage power conditioner, which saves mass.

Propellant (typically hydrogen, although many other volatiles can be used) is first ionized by helicon waves and then transferred to a second magnetic chamber where it is accelerated to ten million degrees K by an oscillating electric and magnetic fields, also known as the ponderomotive force.

A hybrid two-stage magnetic nozzle converts the spiraling motion into axial thrust at 97% efficiency.

Franklin Chang-Diaz, et al., “The Physics and Engineering of the VASIMR Engine,” AIAA conference paper 2000-3756, 2000.

From High Frontier by Philip Eklund

Electrostatic

Colloid

ESTAT: Colloid
Exhaust Velocity43,000 m/s
Specific Impulse4,383 s
Thrust8,000 N
Thrust Power0.2 GW
Mass Flow0.19 kg/s
Total Engine Mass20,000 kg
T/W0.04
Thermal eff.85%
Total eff.85%
Fuel200MWe input
RemassColloid
Remass AccelElectrostatic
Acceleration
Specific Power116 kg/MW

Similar to Ion, but utilizing tiny droplets instead of ions.

Field-Emission Electric (FEEP)

Field-emission electric propulsion, a type of Colloid thruster.

Hall Effect (HET)

Hall Effect Thruster

Hall Effect
Hall Effect
Exhaust Velocity19,620 m/s
Specific Impulse2,000 s
Thrust3,300 N
Number Thrusters300
Thrust/Engine11 N
Thrust Power32.4 MW
Mass Flow0.17 kg/s
Total Engine Mass85,469 kg
T/W4.00e-03
Frozen Flow eff.73%
Thermal eff.73%
Total eff.53%
Fuel60MWe input
RemassMagnesium
Remass AccelElectrostatic
Acceleration
Specific Power2,640 kg/MW

This ion rocket accelerates ions using the electric potential maintained between a cylindrical anode and negatively charged plasma which forms the cathode.

To start the engine, the anode on the upstream end is charged to a positive potential by a power supply. Simultaneously, a hollow cathode at the downstream end generates electrons. As the electrons move upstream toward the anode, an electromagnetic field traps them into a circling ring at the downstream end.

This gyrating flow of electrons, called the Hall current, gives the Hall thruster its name.

The Hall current collides with a stream of magnesium propellant, creating ions. As magnesium ions are generated, they experience the electric field between the anode (positive) and the ring of electrons (negative) and exit as an accelerated ion beam.

A significant portion of the energy required to run the Hall Effect thruster is used to ionize the propellant, creating frozen flow losses.

This design also suffers from erosion of the discharge chamber.

On the plus side, the electrons in the Hall current keep the plasma substantially neutral, allowing far greater thrust densities than other ion drives.

Novosti Kosmonavtiki, 1999.

From High Frontier by Philip Eklund

Ion

Ion
Exhaust Velocity210,000 m/s
Specific Impulse21,407 s
Thrust10,000 N
Thrust Power1.1 GW
Mass Flow0.05 kg/s
Total Engine Mass400,000 kg
T/W3.00e-03
Thermal eff.96%
Total eff.96%
Fuel800MWe input
RemassArgon
Remass AccelElectrostatic
Acceleration
Specific Power381 kg/MW
DAWN mission NSTAR
Propulsion SystemIon
Exhaust Velocity30,411 m/s
Specific Impulse3,100 s
Thrust9.00e-05 N
Thrust Power1.4 W
Mass Flow2.96e-09 kg/s
Total Engine Mass26 kg
T/W3.60e-07
FuelSolar Photons
ReactorPhotovoltaic array
RemassXenon
Remass AccelElectrostatic
Acceleration
Wet Mass1,210 kg
Dry Mass785 kg
Mass Ratio1.54 m/s
ΔV13,159 m/s
Specific Power1.86e+07 kg/MW
Umbrella Ship
Propulsion SystemIon
Exhaust Velocity80,442 m/s
Specific Impulse8,200 s
Thrust490 N
Thrust Power19.7 MW
Mass Flow0.01 kg/s
FuelFission:
Uranium 235
ReactorNuclear Power
Reactor (electric)
RemassCesium
Remass AccelElectrostatic
Acceleration
Wet Mass660,000 kg
Dry Mass328,000 kg
Mass Ratio2.01 m/s
ΔV56,247 m/s

Gridded Electrostatic Ion Thruster. Potassium seeded argon is ionized and the ions are accelerated electrostatically by electrodes. Other propellants can be used, such as cesium and buckyballs. Though it has admirably high exhaust velocity, there are theoretical limits that ensure all Ion drives are low thrust.

It also shares the same problem as the other electrically powered low-thrust drives. In the words of a NASA engineer the problem is "we can't make an extension cord long enough." That is, electrical power plants are weighty enough to make the low thrust an even larger liability. A high powered ion drive will generally be powered by a nuclear reactor, Nuclear Electric Propulsion (NEP). Low powered ion drives can get by with solar power arrays, all ion drive space probes that exist in the real world use that system. Researchers are looking into beamed power systems, where the ion drive on the spaceship is energized by a laser beam from a remote space station.

If you are interested in the technical details about why ion drives are low thrust, read on.

And it suffers from the same critical thrust-limiting problem as any other ion engine: since you are accelerating ions, the acceleration region is chock full of ions. Which means that it has a net space charge which repels any additional ions trying to get in until the ones already under acceleration manage to get out, thus choking the propellant flow through the thruster.

The upper limit on thrust is proportional to the cross-sectional area of the acceleration region and the square of the voltage gradient across the acceleration region, and even the most optimistic plausible values (i.e. voltage gradients just shy of causing vacuum arcs across the grids) do not allow for anything remotely resembling high thrust.

You can only increase particle energy so much; you then start to get vacuum arcing across the acceleration chamber due to the enormous potential difference involved. So you can't keep pumping up the voltage indefinitely.

To get higher thrust, you need to throw more particles into the mix. The more you do this, the more it will reduce the energy delivered to each particle.

It is a physical limit. Ion drives cannot have high thrusts.

Ion Drive
Ion Drive
Exhaust Velocity78,480 m/s
Specific Impulse8,000 s
Thrust1,444 N
Number Thrustersx361
Thrust/Engine4 N
Thrust Power56.7 MW
Mass Flow0.02 kg/s
Total Engine Mass120,149 kg
T/W1.00e-03
Frozen Flow eff.96%
Thermal eff.99%
Total eff.95%
Fuel60MWe input
RemassMagnesium
Remass AccelElectrostatic
Acceleration
Specific Power2,120 kg/MW

In space, an electrostatic particle accelerator is effectively an electric rocket.

The illustrated design uses a combination of microwaves and spinning magnets to ionize the propellant, eliminating the need for electrodes, which are susceptible to erosion in the ion stream.

The propellant is any metal that can be easily ionized and charge-separated. A suitable choice is magnesium, which is common in asteroids that were once part of the mantles of shattered parent bodies, and which volatilizes out of regolith at the relatively low temperature of 1800 K.

The ion drive accelerates magnesium ions using a negatively charged grid, and neutralizes them as they exit. The grids are made of C-C, to reduce erosion.

Since the stream is composed of ions that are mutually repelling, the propellant flow is limited to low values proportional to the cross-sectional area of the acceleration region and the square root of the voltage gradient.

Decoupling the acceleration from the extraction process into a two-stage system allows the voltage gradients to reach 30 kV without vacuum-arcing, corresponding to exit velocities of 80-210 km/sec.

A 60 MWe system with a thrust of 1.5 kN utilizes a hexagonal array, 25 meters across, containing 361 accelerators. Frozen flow efficiencies are high (96%).

To boost the acceleration (corresponding to the “open-cycle cooling” game rule), colloids are accelerated instead of ions. Colloids (charged sub-micron droplets of a conducting non-metallic fluid) are more massive than ions, allowing increased thrust at the expense of fuel economy.

J. Beatty of Hughes, 1990.

From High Frontier by Philip Eklund

{ IBS Agamemnon }

IBS Agamemnon
Propulsion SystemIon
Exhaust Velocity220,000 m/s
Specific Impulse22,426 s
Thrust10,000,000 N
Thrust Power1.1 TW
Mass Flow45 kg/s
FuelDeuterium-Deuterium
Fusion
ReactorFusion Power
Reactor(electric)
RemassCadmium
Remass AccelElectrostatic
Acceleration
Wet Mass100,000,000 kg
Dry Mass28,000,000 kg
Mass Ratio3.57 m/s
ΔV280,052 m/s
Ship Mass8,000,000 kg
Cargo Mass20,000,000 kg
Length400 m
Length spin arm100 m
Cargo Tug Slingshot
Propulsion SystemIon
Exhaust Velocity280,000 m/s
Specific Impulse28,542 s
Thrust5,460,000 N
Thrust Power764.4 GW
Mass Flow20 kg/s
FuelDeuterium-Deuterium
Fusion
ReactorFusion Power
Reactor(electric)
RemassCadmium
Remass AccelElectrostatic
Acceleration
Wet Mass512,600,000 kg
Dry Mass501,600,000 kg
Mass Ratio1.02 m/s
ΔV6,074 m/s

Fictional Interplanetary BoostShip Agamemnon from Jerry Pournelle's short story "Tinker". This fictional ship is a species of Ion drive utilizing cadmium and powered by deuterium fusion. Looking at its performance I suspect that in reality no Ion drive could have such a high thrust. The back of my envelope says that you'd need one thousand ultimate Ion drives to get this much thrust.

Electrothermal

ArcJet

ArcJet
Exhaust Velocity20,000 m/s
Specific Impulse2,039 s
Thrust2 N
Thrust Power20.0 kW
Mass Flow1.00e-04 kg/s
Fuel100kWe input
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Arc Heater
Thrust DirectorNozzle

Hydrogen propellant is heated by an electrical arc.

Arcjet
Arcjet
Exhaust Velocity19,620 m/s
Specific Impulse2,000 s
Thrust3,200 N
Number Thrusters32
Thrust/Engine100 N
Thrust Power31.4 MW
Mass Flow0.16 kg/s
Total Engine Mass22,369 kg
T/W0.01
Frozen Flow eff.60%
Thermal eff.87%
Total eff.52%
Fuel60MWe input
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Arc Heater
Thrust DirectorNozzle
Specific Power713 kg/MW

A working fluid such as hydrogen can be heated to 12,000 K by an electric arc. Since the temperatures imparted are not limited by the melting point of tungsten, as they are in a sold core electrothermal engine such as a resistojet, the arcjet can burn four times as hot. However, the thoriated tungsten electrodes must be periodically replaced.

When used as an electrothermal thruster, the arcjet attains a specific impulse of 2 ksec with frozen-flow efficiencies of 60%. When used for mining beneficiation, regolith or ore is initially processed with a 1 Tesla magnetic separator and impact grinder (3.5 tonnes), before being vaporized in the arcjet. The arcjet can also be used for arc welding.

From High Frontier by Philip Eklund

Microwave Electrothermal

Microwave Electrothermal Thruster

MET Steamer Amplitrons
MET Steamer Amplitrons
Exhaust Velocity9,810 m/s
Specific Impulse1,000 s
Thrust/Engine30 N
Number Thrustersx400
Thrust12,000 N
Thrust Power58.9 MW
Mass Flow1 kg/s
Total Engine Mass123,302 kg
T/W0.01
Frozen Flow eff.95%
Thermal eff.85%
Total eff.81%
Fuel60MWe input
RemassWater
Remass AccelThermal Accel:
Microwave Heater
Thrust DirectorMagnetic Nozzle
Specific Power2,095 kg/MW

This device works by generating microwaves in a cylindrical resonant, propellant-filled cavity, thereby inducing a plasma discharge through electromagnetic coupling. The discharge performs either mining or thrusting functions.

In its mining capacity, the head brings to bear focused energy, tuned at close quarters by the local microwave guides, to a variety of frequencies designed to resonate and shatter particular minerals or ice.

In its electrothermal thruster (MET) capacity, the microwave-sustained plasma superheats water, which is then thermodynamically expanded through a magnetic nozzle to create thrust. The MET needs no electrodes to produce the microwaves, which allows the use of water propellant (the oxygen atoms in a steam discharge would quickly dissolve electrodes).

MET steamers can reach 900 seconds of specific impulse due to the high (8000 K) discharge source temperatures, augmented by rapid hydrogen-oxygen recombination in the nozzle. Vortex stabilization produces a well-defined axisymmetric flow. However, the specific impulse is ultimately limited by the maximum temperature (~ 2000 K) that can be sustained by the thruster walls.

The illustration shows a microwave plasma discharge created by tuning the TM(011) mode for impedance-matched operation. This concentrates the most intense electric fields along the cavity axis, placing 95% of the energy into the propellant, with less than 5% lost into the discharge tube walls. Regenerative water cooling is used throughout.

For pressures of 45 atm, each unit can produce 30 N of thrust. The thrust array contains 400 such units, at 50 kg each.

“Development of a High Power Microwave Thruster, with a Magnetic Nozzle, for Space Applications.” John L. Power and Randall A. Chapman, Lewis Research Center, 1989.

From High Frontier by Philip Eklund

Resistojet

Resistojet
Exhaust Velocity2,900 m/s
Specific Impulse296 s
Thrust1 N
Thrust Power0.7 kW
Mass Flow2.00e-04 kg/s
Thermal eff.80%
Total eff.80%
Fuel100kWe input
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Resistance Heater
Thrust DirectorNozzle

In a resistojet, ropellant flows over a resistance-wire heating element (much like a space heater or toaster) then the heated propellant escapes out the exhaust nozzle. They are mostly used as attitude jets on satellites, and in situations where energy is more plentiful than mass.

Tungsten Resistojet
Tungsten Resistojet
Exhaust Velocity9,810 m/s
Specific Impulse1,000 s
Thrust9,900 N
Thrust Power48.6 MW
Mass Flow1 kg/s
Total Engine Mass42,601 kg
T/W0.02
Thermal eff.80%
Total eff.80%
Fuel60MWe input
RemassLiquid Hydrogen
Remass AccelThermal Accel:
Resistance Heater
Thrust DirectorNozzle
Specific Power877 kg/MW

Tungsten, the metal with the highest melting point (3694 K), may be used to electric-resistance heat ore for smelting or propellant for thrusting. In the latter mode, the resistojet is an electro-thermal rocket that has a specific impulse of 1 ksec using hydrogen heated to 3500K. The frozen flow efficiency (without hydrogen recombination) is 85%. Internal pressures are 0.1 MPa (1 atm). To reduce ohmic losses, the heat exchanger uses a high voltage (10 kV) low current (12.5 kiloamp) design. The specific power of the thruster is 260 kg/MWj and the thrust to weight ratio is 8 milli-g.

Once arrived at a mining site, the tungsten elements, together with wall of ceramic lego-blocks (produced in-situ from regolith by magma electrolysis) are used to build an electric furnace. Tungsten resistance-heated furnaces are essential in steel-making. They are used to sand cast slabs of iron from fines (magnetically separated from regolith), refine iron into steel (using carbon imported from Type C asteroids), and remove silicon and sulfur impurities (using CaAl2O4 flux roasted from lunar highland regolith).

From High Frontier by Philip Eklund

Wakefield E-Beam

Wakefield E-Beam
Wakefield E-Beam
Exhaust Velocity19,620 m/s
Specific Impulse2,000 s
Thrust4,600 N
Thrust Power45.1 MW
Mass Flow0.23 kg/s
Total Engine Mass41,837 kg
T/W0.01
Frozen Flow eff.85%
Thermal eff.89%
Total eff.76%
Fuel60MWe input
RemassRegolith
Remass AccelThermal Accel:
Arc Heater
Thrust DirectorNozzle
Specific Power927 kg/MW

An e-beam (beam of electrons) is a versatile tool. It can bore holes in solid rock (mining), impart velocity to reaction mass (rocketry), remove material in a computer numerical control cutter (finished part fabrication), or act as a laser initiator (free electron laser).

A wakefield electron accelerator uses a brief (femtosecond) laser pulse to strip electrons from gas atoms and to shove them ahead. Other electrons entering the electron-depleted zone create a repulsive electrostatic force. The initial tight grouping of electrons effectively surf on the electrostatic wave.

Wakefield accelerators a few meters long exhibit the same acceleration as a conventional rf accelerator kilometers in length. In a million-volt-plus electron beam the electrons are approaching lightspeed, so the term relativistic electron beam is appropriate.

The wakefield can be used as an electrothermal rocket similar in principle to the arcjet, but far less discriminating in its choice of propellant.

Tajima 1979.

From High Frontier by Philip Eklund

Fusion

Fusion Containment

There are five general methods for confining plasmas long enough and hot enough for achieving a positive Q (more energy out of a reaction than you need to ignite it, "break even"):


  • Closed-field magnetic confinement (see D-T Fusion Tokamak)
  • Open-field magnetic confinement (see 3He-D mirror cell)
  • Inertial confinement (see D-D inertial fusion)
  • Electrostatic inertial confinement (see 6Li-H fusor)
  • Cold fusion (see H-B cat fusion)
  • Of these reactions, the fusion of deuterium and tritium (D-T), has the lowest ignition temperature (40 million degrees K, or 5.2 keV). However, 80% of its energy output is in highly energetic neutral particles (neutrons) that cannot be contained by magnetic fields or directed for thrust.

    In contrast, the 3He-D fusion reaction (ignition temperature = 30 keV) generates 77% of its energy in charged particles, resulting in substantial reduction of shielding and radiator mass. However, troublesome neutrons comprise a small part of its energy (4% at ion temperatures = 50 keV, due to a D-D side reaction), and moreover the energy density is 10 times less then D-T. Another disadvantage is that 3He is so rare that 240,000 tonnes of regolith scavenging would be needed to obtain a kilogram of it. (Alternatively, helium 3 can be scooped from the atmospheres of Jupiter or Saturn.)

    Deuterium, in contrast, is abundant and cheap. The fusion of deuterium to itself (D-D) occurs at too high a temperature (45 keV) and has too many neutrons (60%) to be of interest. However, the neutron energy output can be reduced to 40% by catalyzing this reaction to affect a 100% burn-up of its tritium and 3He by-products with D.

    The fusion of 10% hydrogen to 90% boron (using 11B, the most common isotope of boron, obtained by processing seawater or borax) has an even higher ignition temperature (200 keV) than 3He-D, and the energy density is smaller. Its advantage is that is suffers no side reactions and emits no neutrons, and hence the reactor components do not become radioactive.

    The 6Li-H reaction is similarly clean. However, both the H-B and 6Li-H reactions run hot, and thus ion-electron collisions in the plasma cause high bremsstrahllung x-ray losses to the reactor first wall.

    From High Frontier by Philip Eklund

    The samples below are from Nuclear Propulsion—A Vital Technology for the Exploration of Mars and the Planets Beyond (1987).

    There are two types of mission. One way missions go from planet A to planet B (AB or A→B) or from planet B to planet A (BA or B→A). Round trip (RT or A→A) missions go from A to B and back to A.

    The bottom line is that inertial confinement fusion is far superior to magnetic confinement fusion.

    Sample Closed-field
    Magnetic Confinement
    (Tokamak)
    Fusion Rocket
    FuelD-3He (spin polarized)
    Specific Impulse20,000 s
    Mass Flow0.308 kg/s
    Engine Alpha5.75 kW/kg
    Engine Mass1,033,000 kg
    Payload Mass200,000 kg
    Sample
    Inertial
    Confinement
    Fusion Rocket
    FuelCat-DD
    Specific Impulse270,000 s
    Mass Flow0.015 kg/s
    Engine Alpha110 kW/kg
    Engine Mass486,000 kg
    Payload Mass200,000 kg
    Sample Tokamak Fusion Rocket
    One-way continuous-burn constant-Isp trajectory
    Mission


    Distance

    DAB (A.U.)
    Mass
    Ratio
    RM
    Initial
    Mass
    Mi (mT)
    Propellant
    Mass
    Mp (mT)
    Payload
    Mass
    ML/Mi (%)
    Travel
    Time
    τAB (days)
    Initial
    Acceleration
    ai (10-3 g0)
    Mars0.5241.7322,1359029.433.0~2.9
    Ceres1.7672.4973,0791,8466.569.42.0
    Jupiter4.2033.5904,4273,1944.5120.0~1.4
    Sample Tokamak Fusion Rocket
    Round-trip trajectory
    Mission



    Mass
    Ratio

    RM
    Propellant
    Mass
    MpA→B
    (mT)
    Propellant
    Mass
    MpB→A
    (mT)
    Propellant
    Mass
    MpA→A
    (mT)
    Initial
    Mass
    Mi
    (mT)
    Travel
    Time
    τAB
    (days)
    Travel
    Time
    τBA
    (days)
    Travel
    Time
    τRT
    (days)
    Mars2.6641,1499022,0513,28443.233.977.1
    Ceres4.6672,6751,8464,5215,754100.569.4169.9
    Jupiter7.7835,1693,1948,3639,596194.3120.0314.3
    Sample Inertial Confinement Fusion Rocket
    Round-trip continuous-burn constant-Isp trajectory
    Mission



    Distance

    DAB (A.U.)
    Mass
    Ratio

    RM
    Initial
    Mass
    Mi
    (mT)
    Propellant
    Mass
    MpA→A
    (mT)
    Payload
    Mass
    ML/Mi (%)
    Travel
    Time
    τAB
    (days)
    Travel
    Time
    τRT
    (days)
    Mars0.5241.104757.371.326.427.755.0
    Ceres1.7671.196820.5134.524.453.1103.7
    Jupiter4.2031.309898212.022.384.6163.6
    Saturn8.5391.453997311.020.1125.5239.8
    Uranus18.1821.6891,159473.017.3194.1364.7
    Neptune29.0581.9011,304618.015.3257.3476.9
    Pluto38.5182.0631,415729.014.1306.6562.7

    The above tables were calculated with the following equations:

    Wf = Mf * g0

    MB = Mf + MpB→A

    1 / α = Mi / MB

    1 / β = MB / Mf

    Pf = Mp / Mi

    RM = 1 / (α * β) (two way)

    RM = 1 / β (one way)

    τAB = (Isp / (F / Wf)) * (1 / β) * ((1 / α) -1) (equation 10)

    τBA = (Isp / (F / Wf)) * (1 / β - 1) (equation 11)

    τRT = τAB + τBA (equation 12a)

    τRT = (Isp / (F / Wf)) * (1 / (α * β) - 1) (equation 12b)

    DAB(m) = ((g0 * Isp2) / (F / Wf))) * (1 / β) * ((1 / sqrt(α)) - 1)2 (equation 13a)

    DBA(m) = ((g0 * Isp2) / (F / Wf))) * ((1 / sqrt(β)) - 1)2 (equation 14)

    DAB(m) = DBA(m) (equation 13b)

    where:

    αp = engine alpha (W/kg)
    DAB = distance between A and B (meters)
    DBA = distance between B and A (meters)
    Isp = engine specific impulse (seconds)
    IMEO = initial mass in Earth orbit (kg)
    MB = dry mass plus just propellant to travel from B to A (kg)
    ML = mass of payload (kg)
    MW = mass of engine (kg)
    Mf = dry mass (kg)
    Mi = initial mass in Earth orbit (kg)
    MpA→A = mass of propellant used traveling round-trip from A to B to A (kg)
    MpA→B = mass of propellant used traveling one-way from A to B (kg)
    MpB→A = mass of propellant used traveling one-way from B to A (kg)
    p = propellant mass flow (kg/s)
    Pf = propellant mass fraction
    RM = spacecraft mass ratio
    τAB = time to travel one way from A to B (seconds)
    τBA = time to travel one way from B to A (seconds)
    τRT = time to travel round trop from A to B to A (seconds)
    Wf = dry weight (Newtons)

    Fusion Fuels

    For more details about fusion fuels, go here.

    Torchship Fusion

    (ed note: Luke Campbell is giving advice to somebody trying to design a torchship. So when he says that magnetic confinement fusion won't work, he means won't work in a torchship. It will work just fine in a weak low-powered fusion drive.)

    For one thing, forget muon catalyzed fusion. The temperature of the exhaust will not be high enough for torch ship like performance.

    You might use a heavy ion beam driven inertial confinement fusion pulse drive, or a Z-pinch fusion pulse drive.

    I don't think magnetic confinement fusion will work — you are dealing with a such high power levels I don't think you want to try confining this inside your spacecraft because it would melt.


    D-T (deuterium-tritium) fusion is not very good for this purpose. You lose 80% of your energy to neutrons, which heat your spacecraft and don't provide propulsion. 80% of a terrawatt is an intensity of 800 gigawatts/(4 π r2) on your drive components at a distance of r from the fusion reaction zone.

    If we assume we need to keep the temperature of the drive machinery below 3000 K (to keep iron from melting, or diamond components from turning into graphite), you would need all non-expendable drive components to be located at least 120 meters away from the point where the fusion pulses go off.

    (ed note: 120 meters = attunation 180,000. 800 gigawatts / 180,000 = 4.2 megawatts)


    D-D (deuterium-deuterium) fusion gives you only 66% of the energy in neutrons. However, at the optimum temperature, you get radiation of bremsstrahlung x-rays equal to at least 30% of the fusion output power.

    For a terawatt torch, this means you need to deal with 960 gigawatts of radiation. You need a 130 meter radius bell for your drive system to keep the temperature down.

    (ed note: 130 meters = attunation 210,000. 960 gigawatts / 210,000 = 4.5 megawatts)


    D-3He (deuterium-helium-3) fusion gives off maybe 5% of its energy as neutrons. A bigger worry is bremsstrahlung x-rays are also radiated accounting for at least 20% of the fusion output power. This lets you get away with a 66 meter radius bell for a terawatt torch.

    (ed note: 66 meters = attunation 55,000. 250 gigawatts / 55,000 = 4.5 megawatts. I guess 4.5 megawatts is the level that will keep the drive machinery below 3000 k)

    To minimize the amount of x-rays emitted, you need to run the reaction at 100 keV per particle, or 1.16 × 109 K. If it is hotter or colder, you get more x-rays radiated and more heat to deal with.

    This puts your maximum exhaust velocity at 7,600,000 m/s, giving you a mass flow of propellant of 34.6 grams per second at 1 terawatt output, and a thrust of 263,000 Newtons per terawatt.

    This could provide 1 G of acceleration to a spacecraft with a mass of at most 26,300 kg, or 26.3 metric tons. If we say we have a payload of 20 metric tons and the rest is propellant, you have 50 hours of acceleration at maximum thrust. Note that this is insufficient to run a 1 G brachistochrone. Burn at the beginning for a transfer orbit, then burn at the end to brake at your destination.


    Note that thrust and rate of propellant flow scales linearly with drive power, while the required bell radius scales as the square root of the drive power. If you use active cooling, with fluid filled heat pipes pumping the heat away to radiators, you could reduce the size of the drive bell somewhat, maybe by a factor of two or three. Also note that the propellant mass flow is quite insufficient for open cycle cooling as you proposed in an earlier post in this thread.

    Due to the nature of fusion torch drives, your small ships may be sitting on the end of a large volume drive assembly. The drive does not have to be solid — it could be a filigree of magnetic coils and beam directing machinery for the heavy ion beams, plus a fuel pellet gun. The ion beams zap the pellet from far away when it has drifted to the center of the drive assembly, and the magnetic fields direct the hot fusion plasma out the back for thrust.

    Deuterium-Tritium

    Exhaust Velocity22,000 m/s
    Specific Impulse2,243 s
    Thrust108,000 N
    Thrust Power1.2 GW
    Mass Flow5 kg/s
    Total Engine Mass10,000 kg
    T/W1
    FuelDeuterium-Tritium
    Fusion
    Specific Power8 kg/MW

    Fuel: deuterium and tritium. Propellant: lithium. 1 atom of Deuterium fuses with 1 atom of Tritium to produce 17.6 MeV of energy. One gigawatt of power requires burning a mere 0.00297 grams of D-T fuel per second.

    Note that Tritium has an exceedingly short half-life of 12.32 years. Use it or lose it. Most designs using Tritium included a blanket of Lithium to breed more fresh Tritium fuel.

    Hydrogen-Boron

    H-B Fusion
    Exhaust Velocity980,000 m/s
    Specific Impulse99,898 s
    Thrust61,000 N
    Thrust Power29.9 GW
    Mass Flow0.06 kg/s
    Total Engine Mass300,000 kg
    T/W0.02
    FuelHydrogen-Boron
    Fusion
    Specific Power10 kg/MW

    Fuel is Hydrogen and Boron-11. Propellant is hydrogen. Bombard Boron-11 atoms with Protons (i.e., ionized Hydrogen) and you get a whopping 16 Mev of energy, three Alpha particles, and no deadly neutron radiation.

    Well, sort of. Current research indicatates that there may be some neutrons. Paul Dietz says there are two nasty side reactions. One makes a Carbon-12 atom and a gamma ray, the other makes a Nitrogen-14 atom and a neutron. The first side reaction is quite a bit less likely than the desired reaction, but gamma rays are harmful and quite penetrating. The second side reaction occurs with secondary alpha particles before they are thermalized.

    The Hydrogen - Boron reaction is sometimes termed "thermonuclear fission" as opposed to the more common "thermonuclear fusion".

    A pity about the low thrust. The fusion drives in Larry Niven's "Known Space" novels probably have performance similar to H-B Fusion, but with millions of newtons of thrust.

    It sounded too good to be true, so I asked "What's the catch?"

    The catch is, you have to arrange for the protons to impact with 300 keV of energy, and even then the reaction cross section is fairly small. Shoot a 300 keV proton beam through a cloud of boron plasma, and most of the protons will just shoot right through. 300 keV proton beam against solid boron, and most will be stopped by successive collisions without reacting. Either way, you won't likely get enough energy from the few which fuse to pay for accelerating all the ones which didn't.

    Now, a dense p-B plasma at a temperature of 300 keV is another matter. With everything bouncing around at about the right energy, sooner or later everything will fuse. But containing such a dense, hot plasma for any reasonable length of time, is well beyond the current state of the art. We're still working on 25 keV plasmas for D-T fusion.

    If you could make it work with reasonable efficiency, you'd get on the order of ten gigawatt-hours of usable power per kilogram of fuel.

    Professor N. Rostoker, et. al think they have the solution, utilizing colliding beams. Graduate Student Alex H.Y. Cheung is looking into turning this concept into a propulsion system.

    Helium3-Deuterium

    He3-D Fusion
    Exhaust Velocity7,840,000 m/s
    Specific Impulse799,185 s
    Thrust49,000 N
    Thrust Power0.2 TW
    Mass Flow0.01 kg/s
    Total Engine Mass1,200,000 kg
    T/W4.00e-03
    FuelHelium3-Deuterium
    Fusion
    Specific Power6 kg/MW

    Fuel is helium3 and deuterium. Propellant is hydrogen. 1 atom of Deuterium fuses with 1 atom of Helium-3 to produce 18.35 MeV of energy. One gigawatt of power requires burning a mere 0.00285 grams of 3He-D fuel per second.

    Inertial Confinement

    Inertial Confinement Fusion is in the Pulse section.

    Magnetic Confinement

    MC-Fusion
    Thrust Power200 GW
    Exhaust velocity8,000,000 m/s
    Thrust50,000 n
    Engine mass0.6 tonne
    T/W >1.0yes
    Gasdynamic Mirror
    Exhaust Velocity1,960,000 m/s
    Specific Impulse199,796 s
    Thrust47,000 N
    Thrust Power46.1 GW
    Mass Flow0.02 kg/s
    FuelDeuterium-Tritium
    Fusion
    ReactorMagnetic Confinement
    Linear
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle

    A magnetic bottle contains the fusion reaction. Very difficult to do. Researchers in this field say that containing fusion plasma in a magnetic bottle is like trying to support a large slab of gelatin with a web of rubber bands. Making a magnetic bottle which has a magnetic rocket exhaust nozzle is roughly 100 times more difficult.

    Since the engine is using a powerful but tightly controlled magnetic field, it might be almost impossible to have a cluster of several magnetic confinement fusion engines. The magnetic fields will interfere with each other.

    Discovery II
    Discovery II
    PropulsionHelium3-Deuterium
    MC Fusion
    Exhaust Velocity347,000 m/s
    Specific Impulse35,372 s
    Thrust18,000 N
    Thrust Power3.1 GW
    Mass Flow0.05 kg/s
    FuelHelium3-Deuterium
    Fusion
    ReactorMagnetic Confinement
    Toroidal
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle
    Wet Mass1,690,000 kg
    Dry Mass883,000 kg
    Mass Ratio1.91 m/s
    ΔV225,258 m/s
    Specific Power3.5 kW/kg (3,540 W/kg)
    Initial Acceleration1.68 milli-g
    Payload172,000 kg
    Length240 m
    Diameter60 m wide
    D-T Fusion Tokamak
    D-T Fusion Tokamak
    Exhaust Velocity66,800 m/s
    Specific Impulse6,809 s
    Thrust66,800 N
    Thrust Power2.2 GW
    Mass Flow1 kg/s
    Total Engine Mass197,000 kg
    T/W0.04
    Frozen Flow eff.77%
    Thermal eff.85%
    Total eff.65%
    FuelDeuterium-Tritium
    Fusion
    ReactorMagnetic Confinement
    Toroidal
    RemassReaction
    Products
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle
    Specific Power88 kg/MW

    Of all the fusion reactions, the easiest to attain is a mixture of the isotopes of hydrogen called deuterium and tritium (D-T). This reaction is “dirty”, only 20% of the reaction power is charged particles (alphas) that can be magnetically extracted with a diverter for power or thrust. The remaining energy (neutron, bremsstrahlung, and cyclotron radiation) must be captured in a surrounding jacket of cold dense Li plasma. The heated lithium is either exhausted as open-cycle coolant, or recirculated through a heat engine (to generate the power needed for the microwave plasma heater).

    The 2 GWth magnetically-confined reactor shown uses eight poloidal superconducting 30 Tesla coils, twisted into a Tokamak configuration. These weigh 22 tonnes with stiffeners and neutron shielding.

    The pulsed D-T plasma, containing tens of megamps, is super-heated by 50 MW of microwaves or colliding beams to 20 keV. The Q (gain factor) is 40. Closed field line devices such as this can ignite and burn, in which case the Q goes to infinity and microwave heating is no longer needed. However, since ignition is inherently unstable (once ignited, the plasma rapidly heats away from the ignition point), the reactor is kept at slightly below ignition.

    Fuel is replenished at 24 mg/sec by gas puffing to maintain a plasma ion density of 5 × 1020/m3 at 26 atm. At a power density of 250 MWth /m3, the lithium-cooled first wall has a neutron loading of 1 MW/m2 and a radiation loading of 5 MW/m2.

    More advanced vortex designs, which do away with the first wall, separate the hot fusion fuel from the cool lithium plasma by swirling the mixture. The thermal efficiency is 50% in open-cycle mode.

    Williams, Borowski, Dudzinski, and Juhasz, “A Spherical Torus Nuclear Fusion Reactor Space Propulsion Vehicle Concept for Fast Interplanetary Travel,” Lewis Research Center, 1998.

    (The Tokamak used in High Frontier is a smaller lower tech version of the Lewis design, which uses aneutronic 3He-D fuel.)

    From High Frontier by Philip Eklund
    3He-D Mirror Cell
    3He-D Mirror Cell
    Exhaust Velocity313,920 m/s
    Specific Impulse32,000 s
    Thrust10,600 N
    Thrust Power1.7 GW
    Mass Flow0.03 kg/s
    Total Engine Mass106,667 kg
    T/W0.01
    Frozen Flow eff.92%
    Thermal eff.90%
    Total eff.83%
    FuelHelium3-Deuterium
    Fusion
    ReactorMagnetic Confinement
    Linear
    RemassReaction
    Products
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle
    Specific Power64 kg/MW

    Helium 3 is an isotope of helium, and deuterium (abbreviated D) is an isotope of hydrogen. The 3He-D fusion cycle is superior to the D-T cycle since almost all the fusion energy, rather than just 20%, is deposited in the plasma as fast charged particles.

    Magnetic containers with a linear rather than toroidal geometry, such as steady-state mirrors, have superior ratios of plasma pressure to magnet pressure (β >30%) and higher power densities necessary for reaching the high (50 keV) 3He-D operating temperatures.

    The mirror design shown is a tube of 11 Tesla superconducting magnetic coils, with choke coils for reflection at the ends. The magnets weigh 12 tonnes, plus another 24 tonnes for 60 cm of magnet radiation shielding and refrigeration. A mirror has low radiation losses (20% bremsstrahlung, 3% neutrons) compared to its end losses (77% fast charged particles). These losses limit the Q to about unity and prevent ignition. (This is not a problem for propulsion, since reaching break-even is not required to achieve thrust. The plasma is held in stable energy equilibrium by the constant injection of auxiliary microwave heating.)

    The Q can be improved by a tandem arrangement: stacking identical mirror cells end to end so that one’s loss is another’s gain. The exhaust exiting one end can be converted to power by direct conversion (MHD), and the other end’s exhaust can be expanded in a magnetic flux tube for thrust.

    Mirrors improved by vortex technology, called field-reversed mirrors, introduce an azimuthal electron current which creates a poloidal magnetic field component strong enough to reverse the polarity of the magnetic induction along the cylindrical axis. This creates a hot compact toroid that both plugs end losses and raises the temperature of the scrape-off plasma layer fourfold (to 2.5 keV), corresponding to a specific impulse of 32 ksec.

    Mirrors, like all magnetic fusion devices, can readily augment their thrust by open-cycle cooling.

    “Considerations for Steady-State FRC-Based Fusion Space Propulsion,” M.J. Schaffer, General Atomics Project 4437, Dec 2000.

    From High Frontier by Philip Eklund

    Polywell Fusor

    H-Li6 Fusor Reactor
    H-Li6 Fusor
    Exhaust Velocity19,620 m/s
    Specific Impulse2,000 s
    Thrust67,100 N
    Thrust Power0.7 GW
    Mass Flow3 kg/s
    Total Engine Mass54,000 kg
    T/W0.13
    Frozen Flow eff.92%
    Thermal eff.90%
    Total eff.83%
    FuelHydrogen-Lithium6
    Fusion
    ReactorElectrostatic
    Confinement
    RemassReaction
    Products
    Remass AccelElectrostatic
    Acceleration
    Thrust DirectorMagnetic Nozzle
    Specific Power82 kg/MW

    A Farnsworth-Bussard fusor is little more than two charged concentric spheres dangling in a vacuum chamber, producing fusion through inertial electrostatic confinement. Electrons are emitted from an outer shell (the cathode), and directed towards a central anode called the grid. The grid is a hollow sphere of wire mesh, with the elements magnetically-shielded so that the electrons do not strike them. Instead, they zip right on through, oscillating back and forth about the center, creating a deep electrostatic well to trap the ions of lithium 6 and hydrogen that form the fusion fuel. With a one meter diameter grid and a fuel consumption rate of 7 mg/sec, the fusion power produced is 360 MWth.

    Half of this energy is bremsstrahlung X-rays, which must be captured in a lithium heat engine. The other half are isotopes of helium (3He and 4He), each at about 8 MeV. (Overall efficiency is 36%). Since both products are doubly charged, a 4 MeV electric field will decelerate them and produce two electrons from each, producing an 18 amp current at extremely high voltage.

    An electron gun using this 4 million volt energy would emit electrons at relativistic speeds. This beam dissipates quickly in space, unless neutralized by positrons or converted into a free electron laser beam.

    “Inertia-Electrostatic-Fusion Propulsion Spectrum: Air-Breathing to Interstellar Flight,” R W. Bussard and L. W. Jameson, Journal of Propulsion and Power, v. 11, no. 2, pp. 365-372.

    (Philo Farnsworth, the farm boy who invented the television, spent his last years in a lonely quest to attain break-even fusion in his ultra-cheap fusor devices. His ideas are enjoying a renaissance, thanks to Dr. Bussard, and working fusion reactors are making an appearance in high school science fairs. On the theory that the fusor is power-limited, I have scaled down Bussard’s 10 GW design to 360 MW.)

    From High Frontier by Philip Eklund

    { AV:T Fusion }

    AV:T Fusion
    Cruise mode
    Exhaust Velocity832,928 m/s
    Specific Impulse84,906 s
    Thrust245,250 N
    Thrust Power0.1 TW
    Mass Flow0.29 kg/s
    FuelHelium3-Deuterium
    Fusion
    Combat mode
    Exhaust Velocity104,116 m/s
    Specific Impulse10,613 s
    Thrust48,828,125 N
    Thrust Power2.5 TW
    Mass Flow469 kg/s
    FuelHelium3-Deuterium
    Fusion

    Fictional magnetic bottle fusion drive from the Attack Vector: Tactical wargame. It uses an as yet undiscovered principle to direct the heat from the fusion reaction out the exhaust instead of vaporizing the reaction chamber. Like the VASIMR it has "gears", a combat mode and a cruise mode. The latter increases specific impulse (exhaust velocity) at the expense of thrust.

    In the illustration, the spikes are solid-state graphite heat radiators, the cage the spikes emerge from is the magnetic bottle, the sphere is the crew quarters and the yellow rectangles are the retractable power reactor heat radiators. The ship in the lower left corner is signaling its surrender by deploying its radiators.

    { THS Fusion Pulse }

    Fusion Pulse low gear
    Exhaust Velocity150,000 m/s
    Specific Impulse15,291 s
    Thrust80,000 N
    Mass Flow0.53 kg/s
    T/W2
    Fusion Pulse high gear
    Exhaust Velocity300,000 m/s
    Specific Impulse30,581 s
    Thrust40,000 N
    Mass Flow0.13 kg/s
    T/W1
    Both
    Thrust Power6.0 GW
    Total Engine Mass4,000 kg
    FuelHelium3-Deuterium
    Fusion
    Specific Power1 kg/MW

    Fictional inertial-confinement fusion drive from the game GURPS: Transhuman Space. Like the VASIMR it has "gears", one increases specific impulse (exhaust velocity) at the expense of thrust.

    Nuclear Thermal

    These use the heat generated from a nuclear reaction to heat up propellant. The nuclear reaction is controlled by adjusting the amount of free neutrons inside the mass of fissioning material.

    As a side effect, if you have a cluster of several such engines it is vitally important to have distance and neutron shields between them. Otherwise the nuclear reaction in each engine will flare out of control due to the neutron flux from its neighbor engines.

    Solid Core

    Solid Core NTR
    3200° K
    Exhaust velocity (H1)16,000? m/s
    Exhaust velocity (H2)8,093 m/s
    Exhaust velocity (CH4)6,318 m/s
    Exhaust velocity (NH3)5,101 m/s
    Exhaust velocity (H2O)4,042 m/s
    Exhaust velocity (CO2)3,306 m/s
    Exhaust velocity (CO or N2)2,649 m/s

    Nuclear thermal rocket / solid core fission. It's a real simple concept. Put a nuclear reactor on top of an exhaust nozzle. Instead of running water through the reactor and into a generator, run hydrogen through it and into the nozzle. By diverting the hydrogen to a turbine generator 60 megawatts can be generated. The reactor elements have to be durable, since erosion will contaminate the exhaust with fissionable materials. The exhaust velocity limit is fixed by the melting point of the reactor.

    Solid core nuclear thermal rockets have a nominal core temperature of 2,750 K (4,490° F).

    Hydrogen gives the best exhaust velocity, but the other propellants are given in the table since a spacecraft may be forced to re-fuel on whatever working fluids are available locally (what Jerry Pournelle calls "Wilderness re-fuelling", Robert Zubrin calls "In-situ Resource Utilization", and I call "the enlisted men get to go out and shovel whatever they can find into the propellant tanks"). For thermal drives in general, and NTR-SOLID in particular, the exhaust velocity imparted to a particular propellant by a given temperature is proportional to 1 / sqrt( molar mass of propellant chemical ).

    The value for "hydrogen" in the table is for molecular hydrogen, i.e., H2. Atomic hydrogen would be even better, but unfortunately it tends to explode at the clank of a falling dust speck (Heinlein calls atomic hydrogen "Single-H"). Another reason to avoid hydrogen is the difficulty of storing the blasted stuff, and its annoyingly low density (Ammonia is about eight times as dense!). Exhaust velocities are listed for a realistically attainable core temperature of 3200 degrees K for the propellants Hydrogen (H2), Methane (CH4), Ammonia (NH3), Water (H2O), Carbon Dioxide (CO2), Carbon Monoxide (CO), and Nitrogen (N2).

    The exhaust velocities are larger than what one would expect given the molecular weight of the propellants because in the intense heat they break down into their components. Ammonia is nice because it breaks down into gases (Hydrogen and Nitrogen). Methane is nasty because it breaks down into Hydrogen and Carbon, the latter tends to clog the reactor with soot deposits. Water is most unhelpful since it doesn't break down much at all.

    Dr. John Schilling figures that as an order of magnitude guess, about one day of full power operation would result in enough fuel burnup to require reprocessing of the fissionable fuel elements. (meaning that while there is still plenty of fissionables in the fuel rod, enough by-products have accumulated that the clogged rod produces less and less energy) A reprocessing plant could recover 55-95% of the fuel. With reprocessing, in the long term each totally consumed kilogram of plutonium or highly enriched uranium (HEU) will yield ~1E10 newton-seconds of impulse at a specific impulse of ~1000 seconds.

    Dr. Schilling also warns that there is a minimum amount of fissionable material for a viable reactor. Figure a minimum of 50 kilograms of HEU.

    One problem with solid-core NTRs is that if the propellant is corrosive, that is, if it is oxidizing or reducing, heating it up to three thousand degrees is just going to make it more reactive. Without a protective coating, the propellant will start corroding away the interior of the reactor, which will make for some real excitement when it starts dissolving the radioactive fuel rods. What's worse, a protective coating against an oxidizing chemical is worthless against a reducing chemical, which will put a crimp in your wilderness refueling. And trying to protect against both is an engineering nightmare. Oxidizing propellants include oxygen, water, and carbon dioxide, while reducing propellants include hydrogen, ammonia, and methane. Carbon Monoxide is neither, as the carbon atom has a death-grip on the oxygen atom.

    Keep in mind that the oxidizing/reducing effect is only a problem with solid-core NTRs, not the other kinds. This is because only the solid-core NTRs have solid reactor elements exposed to the propellant (for heating).

    NERVA

    NERVA
    Thrust Power0.198-0.065 GW
    Exhaust velocitySee Table
    Thrust49,000 n
    Engine mass10 tonne
    T/W >1.0no
    NERVA (H2)
    Exhaust Velocity8,093 m/s
    Specific Impulse825 s
    Thrust49,000 N
    Thrust Power0.2 GW
    Mass Flow6 kg/s
    Total Engine Mass10,000 kg
    T/W0.50
    FuelFission:
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power50 kg/MW
    Resuable Nuclear Shuttle
    Propulsion SystemNERVA
    Exhaust Velocity8,000 m/s
    Specific Impulse815 s
    Thrust344,000 N
    Thrust Power1.4 GW
    Mass Flow43 kg/s
    Wet Mass170,000 kg
    Dry Mass30,000 kg
    Mass Ratio5.67 m/s
    ΔV13,877 m/s
    Widmer Mars Mission
    Propulsion SystemNERVA
    Exhaust Velocity8,000 m/s
    Specific Impulse815 s
    Thrust580,000 N
    Thrust Power2.3 GW
    Mass Flow72 kg/s
    Wet Mass400,000 kg
    Dry Mass150,000 kg
    Mass Ratio2.67 m/s
    ΔV7,847 m/s
    HELIOS 2nd Stage
    Propulsion SystemNTR Solid
    Exhaust Velocity7,800 m/s
    Specific Impulse795 s
    Thrust981,000 N
    Thrust Power3.8 GW
    Mass Flow126 kg/s
    Wet Mass100,000 kg
    Dry Mass6,800 kg
    Mass Ratio14.71 m/s
    ΔV20,968 m/s
    Atomic V-2
    Propulsion SystemNTR Solid
    Exhaust Velocity8,980 m/s
    Specific Impulse915 s
    Thrust1,050,000 N
    Thrust Power4.7 GW
    Mass Flow117 kg/s
    Total Engine Mass4,200 kg
    T/W25
    Wet Mass42,000 kg
    Dry Mass17,000 kg
    Mass Ratio2.47 m/s
    ΔV8,122 m/s
    Specific Power1 kg/MW

    Nuclear Engine for Rocket Vehicle Applications. The first type of NTR-SOLID propulsion systems. It used reactor fuel rods surrounded by a neutron reflector. Unfortunately its thrust to weight ratio is less than one, so no lift-offs with this rocket. The trouble was inadequate propellant mass flow, the result of trying to squeeze too much hydrogen through too few channels in the reactor.

    NERVA Derivative

    NERVA Deriv
    Exhaust Velocity8,085 m/s
    Specific Impulse824 s
    Thrust334,061 N
    Thrust Power1.4 GW
    Mass Flow41 kg/s
    Total Engine Mass10,100 kg
    T/W3
    FuelFission:
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power7 kg/MW

    DUMBO

    General Dumbo
    Thrust Power14.0-4.6 GW
    Exhaust velocitySee Table
    Thrust3,500,000 n
    Engine mass5 tonne
    T/W >1.0yes
    Dumbo (H2)
    Exhaust Velocity8,093 m/s
    Specific Impulse825 s
    Thrust3,500,000 N
    Thrust Power14.2 GW
    Mass Flow432 kg/s
    Total Engine Mass5,000 kg
    T/W71
    FuelFission:
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Dumbo Model AEngine mass0.7 tonneThrust400,000 nPropellant mass flow52 kg/secExhaust velocity7,700 m/secEngine Height0.6 mEngine Radius0.3 mEngine Volume0.2 m3T/W58Dumbo Model BEngine mass2.8 tonneThrust3,560,000 nPropellant mass flow460 kg/secExhaust velocity7,700 m/secEngine Height0.6 mEngine Radius1.0 mEngine Volume1.8 m3T/W130Dumbo Model CEngine mass2.1 tonneThrust400,000 nPropellant mass flow48 kg/secExhaust velocity8,300 m/secEngine Height0.6 mEngine Radius0.4 mEngine Volume0.3 m3T/W20

    This was a competing design to NERVA. It was shelved political decision that, (in order to cut costs on the atomic rocket projects) required both projects to use an already designed NEVRA engine nozzle. Unfortunately, said nozzle was not compatible with the DUMBO active cooling needs. Dumbo does, however, have a far superior mass flow to the NERVA, and thus a far superior thrust. Dumbo actually had a thrust to weight ratio greater than one. NASA still shelved DUMBO because [a] NERVA used off the shelf components and [b] they knew there was no way in heck that they could get permission for nuclear lift-off rocket so who cares if T/W < 0? You can read more about Dumbo in this document.

    Note that the "engine mass" entry for the various models does not include extras like the mass of the exhaust nozzle, mass of control drums, or mass of radiation shadow shield.

    Pebble Bed

    Pebble Bed
    Exhaust Velocity9,530 m/s
    Specific Impulse971 s
    Thrust333,617 N
    Thrust Power1.6 GW
    Mass Flow35 kg/s
    Total Engine Mass1,700 kg
    T/W20
    FuelFission:
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power1 kg/MW

    Particle bed / nuclear thermal rocket AKA fluidized-bed, dust-bed, or rotating-bed reactor. In the particle-bed reactor, the nuclear fuel is in the form of a particulate bed through which the working fluid is pumped. This permits operation at a higher temperature than the solid-core reactor by reducing the fuel strength requirements . The core of the reactor is rotated (approximately 3000 rpm) about its longitudinal axis such that the fuel bed is centrifuged against the inner surface of a cylindrical wall through which hydrogen gas is injected. This rotating bed reactor has the advantage that the radioactive particle core can be dumped at the end of an operational cycle and recharged prior to a subsequent burn, thus eliminating the need for decay heat removal, minimizing shielding requirements, and simplifying maintenance and refurbishment operations.

    Cermet

    Cermet NERVA
    Exhaust Velocity9,120 m/s
    Specific Impulse930 s
    Thrust445,267 N
    Thrust Power2.0 GW
    Mass Flow49 kg/s
    Total Engine Mass9,000 kg
    T/W5
    FuelFission:
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power4 kg/MW
    Cermet NERVA
    Cermet NERVA
    Exhaust Velocity9,810 m/s
    Specific Impulse1,000 s
    Thrust134,400 N
    Thrust Power0.7 GW
    Mass Flow14 kg/s
    Total Engine Mass32,546 kg
    T/W0.42
    Frozen Flow eff.73%
    Thermal eff.96%
    Total eff.70%
    FuelFission:
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power49 kg/MW

    The NERVA (Nuclear Engine for Rocket Vehicle Application) system captures the neutronic energy of a nuclear reaction using a heat exchanger cooled by water or liquid hydrogen. The exchanger uses thin foil or advanced dumbo fuel elements with cermet (ceramic-metal) substrates, jacketed by a beryllium oxide neutron reflector.

    The chamber temperature is limited to 3100K for the extended operational life of the solid fuel elements, which can be fission, fusion, or antimatter. At this temperature, the disassociation of molecular H2 to H significantly boosts specific impulse at chamber pressures below 10 atm.

    A propellant tank pressurized to 2 atm expels the LH2 coolant into the exchanger without the need for turbopumps. This open-cycle coolant is expanded through a hydrogen-cooled nozzle of refractory metal to obtain thrust.

    The efficiencies are 96% thermal, 76% frozen-flow (mainly H2 dissociation, less recombination in the nozzle), and 96% nozzle. A 940 MWth heat exchanger yields a thrust of 134 kN, and a specific impulse of 1 ksec, at a power density of 340 MW/m3.

    Altseimer, et al., “Operating Characteristics and Requirements for the NERVA Flight Engine,” AIAA Paper 70-676, June 1970.

    From High Frontier by Philip Eklund

    LANTR

    LANTR
    LANTR NERVA mode
    Exhaust Velocity9,221 m/s
    Specific Impulse940 s
    Thrust67,000 N
    Thrust Power0.3 GW
    Mass Flow7 kg/s
    RemassLiquid Hydrogen
    LANTR LOX mode
    Exhaust Velocity6,347 m/s
    Specific Impulse647 s
    Thrust184,000 N
    Thrust Power0.6 GW
    Mass Flow29 kg/s
    RemassHydrogen + Oxygen
    LANTR Both
    FuelFission:
    Uranium 235
    ReactorSolid Core
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    SpecialLow-High Gear
    Nuclear DC-X NERVA
    High Gear
    Exhaust Velocity9,810 m/s
    Specific Impulse1,000 s
    Thrust/Engine1,112,000 N
    Thrust5,560,000 N
    Thrust Power27.3 GW
    Mass Flow567 kg/s
    T/W3
    RemassHydrogen
    Specific Power7 kg/MW
    Low Gear
    Exhaust Velocity5,900 m/s
    Specific Impulse601 s
    Thrust/Engine3,336,000 N
    Thrust16,680,000 N
    Thrust Power49.2 GW
    Mass Flow2,827 kg/s
    T/W9
    RemassWater
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power4 kg/MW
    Both
    Number Thrustersx5
    Total Engine Mass199,600 kg
    FuelFission:
    Uranium 235
    SpecialLow-High Gear
    Wet Mass460,000 kg

    LOX-augmented Nuclear Thermal Rocket. One of the systems that can increase thrust by lowering Isp. This concept involves the use of a "conventional" hydrogen (H2) NTR with oxygen (O2) injected into the nozzle. The injected O2 acts like an "afterburner" and operates in a "reverse scramjet" mode. This makes it possible to augment (and vary) the thrust (from what would otherwise be a relatively small NTR engine) at the expense of reduced Isp

    Bi-Modal NTR

    Bimodal NTR Solid (NASA)
    Propulsion SystemNTR Solid Bimodal
    Exhaust Velocity8,980 m/s
    Specific Impulse915 s
    Thrust/Engine66,667 N
    Number Thrustersx3
    Thrust200,000 N
    Thrust Power0.9 GW
    Mass Flow22 kg/s
    Total Engine Mass6,672 kg
    T/W3
    FuelFission:
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    SpecialBimodal
    Wet Mass80,000 kg
    Dry Mass26,830 kg
    Mass Ratio2.98 m/s
    ΔV9,811 m/s
    Specific Power7 kg/MW

    A useful refinement is the Bimodal NTR.

    Say your spacecraft has an honest-to-Johnny NERVA nuclear-thermal propulsion system. Typically it operates for a few minutes at a time, then sits idle for the rest of the entire mission. Before each use, one has to warm up the reactor, and after use the reactor has to be cooled down. Each of these thermal cycles puts stress on the engine. And the cooling process consists of wasting propellant, flushing it through the reactor just to cool it off instead of producing thrust.

    Meanwhile, during the rest of the mission, your spacecraft needs electricity to run life support, radio, radar, computers, and other incidental things.

    So make that reactor do double duty (that's where the "Bimodal" comes in) and kill two birds with one stone. Refer to above diagram. Basically you take a NERVA and attach a power generation unit to the side. The NERVA section is the "cryogenic H2 propellant tank", the turbopump, and the thermal propulsion unit. The power generation section is the generator, the radiator, the heat exchanger, and the compressor.

    Warm up your reactor once, do a thrust burn, stop the propellant flow and use the heat exchanger and radiator to partially cool the reactor to power generation levels, and keep the reactor warm for the rest of the mission while generating electricity for the ship.

    This allows you to get away with only one full warm/cool thermal cycle in the entire mission instead of one per burn. No propellant is wasted as coolant since the radiator cools down the reactor. The reactor supplies needed electricity. And as an added bonus, the reactor is in a constant pre-heated state. This means that in case of emergency one can power up and do a burn in a fraction of the time required by a cold reactor.

    Pretty ingenious, eh?

    And the Pratt & Whitney company went one step better. They took the Bimodal NTR concept and merged it with the LANTR concept to make a Trimodal NTR. Called the Triton, it can be used in LANTR mode when thrust is more important than specific impulse, NTR mode when specific impulse is more important than thrust, and in power generation mode while coasting.

    Dual-mode Fission
    Dual-mode Fission
    Exhaust Velocity9,810 m/s
    Specific Impulse1,000 s
    Thrust124,700 N
    Thrust Power0.6 GW
    Mass Flow13 kg/s
    Total Engine Mass33,000 kg
    T/W0.39
    Thermal eff.94%
    Total eff.94%
    FuelFission:
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power54 kg/MW
    SpecialBimodal
    Thermal Electrical eff.19%
    Electrical Power60 MWe

    When struck by a thermal neutron, a fissile nuclide splits into two fragments plus energy. For example, the fission of the 235U atom produces 165 MeV of energy plus 12 MeV of neutral radiation (gammas and a couple of fast neutrons). The fast neutrons must be thermalized by a low Z moderator (a surrounding blanket of about 80 cm of D2O, Be, liquid or gas D2, or CD4), which returns enough thermal neutrons to the core to sustain the chain reaction. (Thermal neutrons diffuse through the reactor like a low pressure gas.) Alternatively, a molybdenum neutron reflector can be used. Much of a reactor’s mass is constant, regardless of power level. Therefore, nuclear power sources are more attractive at higher power levels.

    The 650 MWth system illustrated is dual mode, which can either generate electricity, or directly exhaust coolant for thrust. It uses a fast reactor with fuel tubes interspersed with cooling tubes. The coolant is lithium, which for electrical power is passed to a potassium boiler at 1650 K. The potassium vapor is passed to a static (AMTEC) or dynamic (turbine) heat engine for power generation (60 MWe), or heats hydrogen in a heat exchanger for thrust (125 kN at a specific impulse of 1 ks). The thermal efficiency is 19% if closed-cycle (for power generation) or 94% if open-cycle (for thrust).

    From High Frontier by Philip Eklund
    Dual-mode Pebble Bed
    Pebble-bed Fission Reactor
    Exhaust Velocity9,810 m/s
    Specific Impulse1,000 s
    Thrust172,700 N
    Thrust Power0.8 GW
    Mass Flow18 kg/s
    Total Engine Mass58,000 kg
    T/W0.30
    Thermal eff.94%
    Total eff.94%
    FuelFission:
    Uranium 233
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power68 kg/MW
    SpecialBimodal
    Electrical Power60 MWe

    This is a graphite-moderated, gas-cooled, nuclear reactor that uses spherical fuel elements called "pebbles". These tennis ball-sized pebbles are made of pyrolytic graphite (which acts as the moderator), interspersed with thousands of micro fuel particles of a fissile material (such as 235U).

    In the reactor illustrated, 360,000 pebbles are placed together to create a 120 MWth reactor. The spaces between the pebbles form the "piping" in the core for the coolant, either propellant or inert He/Xe gas.

    The design illustrated can is dual mode. It can operate either as a generator for 60 MWe of electricity, or act as a solid-core thruster using hydrogen propellant/coolant expelled at a specific impulse of 1 ksec. When used as a thruster, it offers a slight increase in specific impulse but significant acceleration benefits over traditional fission reactors. Moreover, the high temperatures (up to 1900 K) allow higher thermal efficiencies (up to 50%).

    From High Frontier by Philip Eklund

    MITEE

    MInature ReacTor EnginE. MITEE is actually a family of engines. These are small designs, suitable for launching on existing boosters. You can find more details here.

    Basic
    Basic MITEE
    Exhaust Velocity9,810 m/s
    Specific Impulse1,000 s
    Thrust14,000 N
    Thrust Power68.7 MW
    Mass Flow1 kg/s
    Total Engine Mass200 kg
    T/W7
    FuelFission:
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power3 kg/MW

    The baseline design is a fairly conventional NTR. Unlike earlier designs it keeps the fuel elements in individual pressure tubes instead of a single pressure vessel, making it lighter and allowing slightly higher temperatures and a bit better exhaust velocity.

    Monatomic H
    Monatomic-H MITEE
    Exhaust Velocity12,750 m/s
    Specific Impulse1,300 s
    Thrust2,350 N
    Thrust Power15.0 MW
    Mass Flow0.18 kg/s
    Total Engine Mass200 kg
    T/W1
    FuelFission:
    Uranium 235
    ReactorSolid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power13 kg/MW

    This advanced design works at a lower chamber pressure so that some of the H2 propellant disassociates into monatomic hydrogen, although the chamber temperature is only slightly greater. The drawback is that this reduces the mass flow through the reactor, limiting reactor power.

    Hybrid
    HybridMITEE
    Exhaust Velocity17,660 m/s
    Specific Impulse1,800 s
    Thrust1,700 N
    Thrust Power15.0 MW
    Mass Flow0.10 kg/s
    Total Engine Mass10,000 kg
    T/W0.02
    FuelFission:
    Uranium 235
    ReactorSolid Core
    RemassSingle-H
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power666 kg/MW

    The use of individual pressure tubes in the reactor allows some fuel channels to be run at high pressure while others are run at a lower pressure, facilitating a hybrid electro-thermal design. In this design cold H2 is heated in the high pressure section of the reactor, is expanded through a turbine connected to a generator, then reheated in the low pressure section of the reactor before flowing to the nozzle. The electricity generated by the turbine is used to break down more H2 into monatomic hydrogen, increasing the exhaust velocity. Since this is a once through system there is no need for radiators so the weight penalty would not be excessive.

    Liquid Core

    Liquid Core 1
    Exhaust Velocity16,000 m/s
    Specific Impulse1,631 s
    Thrust7,000,000 N
    Thrust Power56.0 GW
    Mass Flow438 kg/s
    Total Engine Mass70,000 kg
    T/W10
    FuelFission:
    Uranium 235
    ReactorLiquid Core
    RemassWater
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power1 kg/MW
    Liquid Core 2
    Exhaust velocity14,700 to 25,500 m/s

    Nuclear thermal rocket / liquid core fission. Similar to an NTR-GAS, but the fissionable core is merely molten, not gaseous. A dense high temperature fluid contains the fissionable material, and the hydrogen propellant is bubbled through to be heated. The propellant will be raised to a temperature somewhere between the melting and boiling point of the fluid. Candidates for the fluid include tungsten (boiling 6160K), osmium (boiling 5770K), rhenium (boiling 6170K), or tantalum (boiling 6370K).

    Liquid core nuclear thermal rockets have a nominal core temperature of 5,250 K (8,990°F).

    The reaction chamber is a cylinder which is spun to make the molten fluid adhere to the walls, the reaction mass in injected radially (cooling the walls of the chamber) to be heated and expelled out the exhaust nozzle.

    Starting up the engine for a thrust burn will be complicated and tricky, shutting it down even more so. Keeping the fissioning fluid contained in the chamber instead of escaping out the nozzle will also be a problem.

    LARS

    LARS
    Exhaust Velocity19,620 m/s
    Specific Impulse2,000 s
    Thrust20,000 N
    Thrust Power0.2 GW
    Mass Flow1 kg/s
    Total Engine Mass1,000 kg
    T/W2
    FuelFission:
    Uranium 235
    ReactorLiquid Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power5 kg/MW
    Luna
    Propulsion SystemLARS
    Exhaust Velocity10,300 m/s
    Specific Impulse1,050 s
    Thrust11,000,000 N
    Thrust Power56.6 GW
    Mass Flow1,068 kg/s
    Total Engine Mass9,000 kg
    T/W125
    FuelFission:
    Uranium 235
    RemassWater
    Thrust DirectorNozzle
    Wet Mass226,000 kg
    Dry Mass45,000 kg
    Mass Ratio5.02 m/s
    ΔV16,623 m/s

    Nuclear thermal rocket / liquid annular reactor system. A type of NTR-LIQUID. You can find more details here

    Vapor Core

    Vapor Core
    Thrust Power1.6 GW
    Exhaust velocity9,800 to
    11,800 m/s
    Thrust330,000 n
    Propellant mass flow30 kg/sec
    Reactor thermal power1,400 to
    1,800 MW
    Total engine mass6.83 tonne
    Fuel element mass total1.35 tonne
    Forward reflector mass0.60 tonne
    Aft reflector mass0.51 tonne
    Radial reflector mass2.47 tonne
    Radiation shield mass0.9 tonne
    Total reactor mass5.83 tonne
    Misc. engine
    component mass
    0.9 tonne
    T/W5
    FuelFission:
    Uranium Hexafluoride
    ReactorVapor Core
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power4 kg/MW

    This is sort of an intermediate step in learning how to design a full-blown Gas Core Nuclear Thermal Rocket. It is basically a solid core NTR where the solid nuclear fuel elements are replaced by chambers filled with uranium235 tetrafluoride vapor. The engine is admirably compact with a nicely low critical mass, and an impressive thrust-to-weight ratio of 5-to-1. However the specific impulse / exhaust velocity is only slightly better than a solid core.

    In other words, the system is not to be developed because it has fantastic performance, but because it will be an educational step to building a system that does.

    The uranium fuel is kept physically separate from the hydrogen propellant, so the exhaust is not radioactive.

    A 330,000 newton thrust NVTR would have a core with almost 4,000 fuel elements, with a core radius of 120 cm, core height of 150 cm, and 1,800 MW. Criticality can be achieved with smaller cores: a core volume five times smaller with radius of 60 cm, height of 120 cm, and power of 360 MW.

    Data is from Conceptual Design of a Vapor Core Reactor Rocket Engine for Space Propulsion by E.T. Dugan, N.J. Diaz, S.A. Kuras, S.P. Keshavmurthy, and I. Maya (1996).

    Reflectors
    SideCompositionThicknessMass
    ForwardBeryllium oxide15 cm0.60 tonne
    AftC-C Composite25 cm0.51 tonne
    RadialBeryllium oxide15 cm2.47 tonne
    CORE: 2000 fuel elements
    Radius0.5 m
    Height1.5 m
    Fuel channel per element12 to 32
    Hydrogen channel per element12 to 32
    Critical mass20 kg
    Hydrogen pressure100 atm
    UF4 pressure100 atm
    Fuel center temperature4,500 K
    Design Values
    Pump Flowrate (Total)75.20 lbm/s
    Pump Discharge Pressure3,924 psia
    Pump Efficiency80.01%
    Turbopump RPM70,000 RPM
    Turbopump Power (each)9,836 HP
    Turbine Inlet Temperature481 deg-R
    Turbine Pressure Ratio1.69
    Turbine Flow Rate (each)33.77 lbm/s
    Reactor Thermal Power1,769 MW
    Fuel Element and Reflector Power1,716 MW
    Nozzle Chamber Temperature5,580 deg-R
    Chamber Pressure (Nozzle Stagnation)1,500 psia
    Nozzle Expansion Area Ratio500:1
    Vacuum Specific Impulse (Delivered)997.8 sec
    Heat Loads
    Nozzle-con (total)30.05 MW
    Nozzle-div (total)22.97 MW
    Reflector (total)35.0 MW
    Typical NVTR Engine Parameters
    Nozzle Area Ratio500
    Fuel Pressure100 atm
    Average Fuel Temperature4000 K
    Maximum Element Heat Flux420 W/cm2
    Nomial Element Length150 cm
    Fuel Volume Fraction0.15
    Coolant Volume Fraction0.15
    Moderator Volume Fraction0.70
    Fuel Element Power0.9 MWt
    Element Heat Transfer Area2141 cm2
    Reactor Core L/D1.5
    Fuel Channel Diameter0.142 cm
    Fuel Channel Sectional Area0.0158 cm2
    Total Fuel Channel Area Per Element0.505 cm2
    Fuel Element Sectional Area3.464 cm2
    Element Diameter (across flats)2.2 cm
    Coolant Channel Diameter0.142 cm
    Coolant Channel Sectional Area0.0158 cm2
    Total Coolant Channel Area Per Element0.505 cm2
    Core Volume1.2 m3
    Core Volume Density1,500 MW/m3
    Fuel Element Mass, Total1.35 MT
    Forward Reflector Mass0.60 MT
    Aft Reflector Mass0.51 MT
    Radial Reflector Mass2.47 MT
    Radiation Shield Mass0.90 MT
    Total Reactor Mass5.83 MT
    Misc. Engine Components Mass0.9 MT
    Total Engine Mass6.83 MT
    Engine F/W5.0

    Gas Core

    Closed Cycle

    Gaseous Core NTR closed 1
    Exhaust Velocity20,405 m/s
    Specific Impulse2,080 s
    Thrust445,000 N
    Thrust Power4.5 GW
    Mass Flow22 kg/s
    Total Engine Mass56,800 kg
    T/W0.80
    FuelFission:
    Uranium Hexafluoride
    ReactorGas Core
    Closed-Cycle
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power13 kg/MW
    Gaseous Core NTR closed 2
    Thrust Power0.6 to 231 GW
    Exhaust velocity10,800 to 31,400 m/s
    Thrust117,700 to 14,700,000 n
    Engine mass30 to 300 tonne
    Engine T/W0.4 to 5.0
    Operating Pressure400 to 1600 atm
    NASA report nuclear lightbulb
    Thrust Power3.7 GW
    Engine Power4.6 GW
    Exhaust velocity18,300 m/s
    Thrust409,000 n
    Engine mass32 tonne
    Engine T/W1.3
    Operating Pressure500 atm
    Propellant mass flow22.3 kg/s
    Liberty Ship
    Propulsion SystemNuclear Lightbulb
    Exhaust Velocity30,000 m/s
    Specific Impulse3,058 s
    Thrust/Engine5,340,000 N
    Number Thrustersx7
    Thrust37,380,000 N
    Thrust Power560.7 GW
    Mass Flow1,246 kg/s
    Total Engine Mass378,000 kg
    T/W10
    FuelFission:
    Uranium Hexafluoride
    Wet Mass2,700,000 kg
    Dry Mass1,600,000 kg
    Mass Ratio1.69 m/s
    ΔV15,697 m/s
    Specific Power0.67 kg/MW

    Closed-cycle gaseous core fission / nuclear thermal rocket AKA "Nuclear Lightbulb". Similar to an open-cycle gas core fission rocket, but the uranium plasma is confined in a fused quartz chamber. It is sort of like a child's classic Easy-Bake Oven. Except that there is propellant instead of cake mix and the light bulb is full of fissioning uranium instead of electricity.

    You can read more about this on the Unwanted Blog in the posts here, here, and here.

    The good news is that unlike the open-cycle GCNR it does not spray glowing radioactive death there is no uranium escaping in the exhaust. The bad news is that the maximum exhaust velocity is halved, as is the Delta-V. Yes, I did ask some experts if it was possible to make some kind of hybrid that could "shift gears" between closed and open cycle. Sorry, there would be no savings over just having two separate engines.

    The maximum exhaust velocity is halved because you are trying to have it both ways at once. The higher the propellant's temperature, the higher the exhaust velocity and rocket Delta-V. Unfortunately solid core nuclear reactors have a distressing habit of vaporizing at high temperatures (as do all material objects). Gas core reactors are attempting to do an end run around this problem, by having the reactor start out as high temperature vapor. But adding the quartz chamber is re-introducing solid material components into the engine, which kind of defeats the purpose. The only thing keeping this from utter failure is the fact that quartz does not heat up as much due to the fact it is transparent.

    Even with the restrictions, it seems possible to make a closed-cycle GCNR with a thrust to weight ratio higher than one. This would allow using the awesome might of the atom to boost truely massive amounts of payload into Terra orbit, without creating a radioactive wasteland with every launch. See the GCNR Liberty Ship for an example. The Liberty Ship can boost in one launch more payload than any given Space Shuttle does in the shuttles entire 10 year operating life. Then the Liberty Ship can land and do it again.

    The ideal solution would be to somehow constrain the uranium by something non-material, such as a mangnetohydrodynamic force field or something like that. Alas, currently such fields can only withstand pressures on the order of the breeze from a flapping mosquito, not the 500 atmospheres of pressure found here. But since researchers are working along the same lines in their attempts to make a fusion reactor, this may change. And I did find a brief reference to something called an "MHD choke" in reference to slowing the escape of uranium into the exhaust stream of an open-cycle gas core rocket. I'm still trying to find more information on that.

    The high pressure is to ensure the uranium vapor is dense enough to sustain a fission reaction.

    VCR light bulb fission
    VCR light bulb fission
    Exhaust Velocity19,620 m/s
    Specific Impulse2,000 s
    Thrust56,400 N
    Thrust Power0.6 GW
    Mass Flow3 kg/s
    Total Engine Mass72,566 kg
    T/W0.08
    FuelFission:
    Uranium Hexafluoride
    ReactorGas Core
    Closed-Cycle
    RemassSeeded Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power131 kg/MW

    Most fission reactors avoid meltdown, but the vapor core reactor (VCR) runs so hot (25000 K) that its core vaporizes.

    At this temperature, the vast majority of the electromagnetic emissions are in the hard ultraviolet range. A “bulb” transparent to this radiation, made of internally-cooled a-silica, bottles the gaseous uranium hexafluoride, while letting the fission energy shine through.

    The operating pressure is 1000 atm. The UF6 fuel is prevented from condensing on the cooled wall by a vortex flow field created by the tangential injection of a neon “buffer” gas near the inside of the transparent wall.

    In a generator mode, the UV uses photovoltaics to generate electricity. In a propulsion mode, the UV heats seeded hydrogen propellant, which exits at a specific impulse of 2000 seconds.

    From High Frontier by Philip Eklund
    NASA Report

    The pictured engine is the "reference engine" described in the report Studies of Specific Nuclear Light Bulb and Open-Cycle Vortex-Stabilized Gaseous Nuclear Rocket Engines. I will give the high-lights from the report, but for all the boring nitty-gritty details you'll have to read the report yourself. Please note that as always, "down" is the direction the thrust is going (to the right in the blueprint) and "up" is the opposite direction (to the left). I apologize for the use of imperial units instead of metric. The report is in imperial, and it is too much of a pain to change everything.

    The important statistics: Specific impulse is 1870 seconds (18,300 m/s), thrust 409,000 newtons, engine mass 32,000 kg, thrust-to-weight ratio 1.3.

    The other stats. The total volume of the reaction chambers (cavities) is 170 cubic feet. There are seven cavities, each six feet long. The cavity pressure is 500 atmospheres. The specific impulse is 1870 seconds (report says can theoretically be from 1500 to 3000 seconds). The total propellant flow (including seed and nozzle transpirant coolant flow) is 49.3 pounds per second. The thrust is 92,000 pounds. The engine power is 4600 megawatts. The engine weight is 70,000 pounds. If one can design a variable-throat-area nozzle (instead of fixed-area) this will result in a major decrease in the required chamber pressure during startup.

    Configuration

    The basic configuration is seven separate unit cavities surrounded by moderator-reflector material in between each cavity (beryllium oxide) and surrounding the entire cavity array (graphite). Each cavity is 6.0 feet long and the total volume of all seven cavities is 169.8 cubic feet. The cavity pressure is 500 atmospheres due to criticality and fuel density considerations.

    Lightbulbs

    In each lightbulb, a critical mass of gaseous uranium creates thermal radiation. The thermal radiation can pass through the transparent quartz crystal walls of the lightbulb, but the uranium vapor cannot. This means no lethal uranium enters the exhaust. Hydrogen propellant flowing over the lightbulb is heated to high temperatures by the thermal radiation and is expelled out the rocket nozzles, producing thrust. The hydrogen is "seeded" with tungsten dust because it too is ordinarily transparent to thermal radiation. The seeding makes it opaque, and allows it to be heated. Seven "lightbulbs" are used instead of one, since that increases the total lightbulb radiating area by about 2.2 times.

    Transparent quartz walls

    The transparent quartz wall of the lightbulb contains lots of coolant channels. This is because the quartz is mostly transparent to thermal radiation, but not totally. And fissioning uranium produces an awful lot of thermal radiation. I told you that nuclear lightbulb designers were trying to have it both ways. The coolant channels are marked "circumferential coolant tubes" in the diagram below.

    Inside a lightbulb

    Inside the lightbulb, neon buffer gas is used to create a vortex ring to suspend the gaseous nuclear fuel (a "radial inflow" vortex). The vortex ring looks like an elongated donut (I know it looks like two separate blobs above, that's due to the fact the diagram is a cross-section). One of the important jobs done by the neon buffer gas is to prevent the 42,000°R uranium plasma from making contact with the lightbulb walls. This would be very bad, as the walls would be instantly vaporized. The neon passes along the lightbulb walls, bends round the end caps, then travels down the long axis of the lightbulb (right down the center of the vortex ring). When it reaches the fore end cap, it is removed from the lightbulb through a port (marked "thru-flow" in diagram above).

    The removed neon is very hot, and contains unburnt uranium and fission products. It is cooled by mixing with low-temperature neon, which condenses the unburnt uranium vapor into hot liquid uranium. The liquid uranium is separated from the neon by a centrifuge and sent back into the vortex (at point marked "fuel injection"). The neon is cooled further then it too is sent back into the vortex (at point marked "buffer gas injection"). While examining the blueprint, I noticed that the centrifuges, and indeed the entire uranium fuel delivery system is conspicuous by its absence. Probably classified.

    Note that the centrifuges is a neat solution to the problem of fission fragments clogging up the fuel. In essence, this design has its own built-in nuclear fuel reprocessing plant. Of course the nasty fission fragments will have to be stored and eventually disposed of.

    Lightbulb dimensions

    The total volume inside all the lightbulbs is 84.9 cubic feet, which is 12.1 cubic feet per lightbulb. The radius of the uranium fuel containing region is 85% of the radius of the transparent wall. While the fissioning uranium fuel has a core temperature of 42,000° Rankine, the outer surface is only at 15,000° Rankine.

    Propellant flow in a lightbulb

    The propellant is assumed to exit with a temperature of 80% of the fuel temperature, or 12,000° Rankine. This is because the quartz transparent walls will reflect about 15% of the thermal radiation back inside. By some compilcated reasoning that you will find in the report, the total thermal radiation from the lightbulbs is 4.37 x 106 BTU/sec. The hydrogen propellant has an "enthalpy" of 1.033 x 105 BTU/pound at 12,000°R. So by dividing the two, one discovers that the entire engine can support a propellant flow rate of 42.3 pounds per second, which means 6.07 lb/sec for each of the seven cavities.

    If that last paragraph confused you, let me explain. As a simple example, if a pound of hydrogen at 5°R contains 2 BTUs ("enthalpy"), and the engine puts out 6 BTU per second, then obviously the engine can heat up 6 / 2 = 3 pounds of hydrogen per second. Why do we care? If you multiply the propellant flow rate by the exhaust velocity you will discover the engine's thrust value. And that's a number we do care about.

    The tungsten dust that the propellant is seeded with has a particle diameter of 0.05 microns. The seed density is 1.32 x 10-2 lb/ft3, which is about 3.9 percent of the inlet propellant density. This can probably be reduced if tungsten dust was in the form of thin flat plates instead of spherical particles.

    The hydrogen propellant enters the pressure shells from the fore end (see "Primary Circuit Inlet" in pressure shell diagram below). A bit is bled off from small H2 flow ports in order to pressurize the interior of the shells, circulating to provide coolant to the engines and machinery. But most of it is fed into the turbopump, then injected into the cavities. Since the fore end of each cavity is almost blocked off by the butt end of the lightbulb, there is only a narrow rim to inject the hydrogen.

    In the diagram to the right, you can see how the propellant is fed from the pink pipe into the pink-and-gold wedge-shaped injectors. I presume there are three injectors per cavity, spraying into the clear area between the transparent wall's coolant manifolds and buffer gas injectors.

    Uranium fuel

    The total fissioning uranium in all seven vortexes be about 25.2 pounds of uranium (about 3.6 pounds per cavity). You would ordinarily need more to ensure nuclear criticality, but the required amount is brought down by the beryllium oxide neutron reflector encasing each cavity. The average uranium fuel density is 0.409 lb/ft3. The total density of the neon-uranium mix inside the vortex is about 0.56 lb/ft3. A unit of neon gas will spend about 3.8 seconds inside the cavity. A unit of uranium will spend about 19 seconds inside the cavity. This implies a uranium fuel flow rate of 0.19 lb/sec per cavity.

    According to my slide rule, if the array of seven cavities is producing 4,600 megawatts, it means that the array is burning a miniscule total of 0.055 grams (0.00012 pounds) of uranium fuel per second (0.0079 grams per cavity per second). It still needs the full 3.6 pounds per cavity to be present in order to burn the fraction of a gram.

    The theoretical maximum specific impulse possible is 2230 seconds. Due to this designs incomplete expansion, transpiration coolant flow in the nozzle, presence of tungsten seeding, and friction losses the specific impulse is reduced to 84% or 1870 seconds. Total propellant flow (allowing for tungsten seeds and transpiration cooling) is 49.3 lb/sec. This would result in a thrust of 92,000 pounds force. For complicated reasons you can find in the report, this implies that the exhaust nozzles are 0.0875 feet in diameter at the throat expanding to 2.04 feet diameter at the exit.

    Uranium refueling

    Careful readers may have noticed how the description avoids mentioning the details on how one gets the uranium into the lightbulbs. This is because it is quite a difficult problem, and each of the proposed solutions has drawbacks. The basic problem is old reliable: all the atomic fireworks inherent in 235U will happen if you merely let too much of it accumulate in one place. You have to store it diffuse and somehow bring it together in the lightbulb.

    Method #1 Store it as uranium hexafluoride gas. This would be in large tanks of low pressure (i.e., low density) and with the tanks full of neutron absorbing foam. Pump enough into the lightbulb, a chain reaction will start, and well before the reaction reaches 13,000°R the uranium will have separated from the fluorine.

    The problem is that now you have the insanely dangerous task of dealing with 13,000°R fluorine gas. At room temperature the blasted stuff will violently react with any element in the known universe except helium and neon. A temperature of 13,000°R makes it about 13,000 times as deadly. It will explosively corrode away anything solid in its path like molten lead on facial tissue. Chemist Derek Lowe sarcastically notes that "At seven hundred freaking degrees, fluorine starts to dissociate into monoatomic radicals, thereby losing its gentle and forgiving nature." You can read more about the suicidal risk of dealing with hot fluorine in his amusing blog post.

    Method #2 Store it as sub-critical chunks of uranium, melt them, and inject the molten uranium into the lightbulb. Uranium melts at 1403°K, which is difficult but not impossible. The plan is to somehow turn the molten uranium into a sort of aerosol mist suspended in hot neon.

    The problem is that the molten uranium wants to plate itself all over the melter and the aerosol spray equipment. Which is annoying if the material in question is something like lead, but disasterous if the material is radioactive and fissionable.

    Method #3 is to store the uranium cold as finely divided dust. As dust it is pumpable, injectable, and it will not plate over everything. Inside the lightbulb the uranium dust will be rapidly heated to vaporization by the nuclear reaction. This method does not have any major problems, except for the common problem of how to protect the transparent wall from being vaporized by the heat.

    Again, the uranium delivery system seems to be totally missing from the blueprint. The only bit present is the short stub of the injector at the top of each lightbulb.

    Pressure shells

    The entire engine is encased in two nested pressure shells constructed of filament-wound fiberglass. The inside of the inner shell is pressurized to 500 atmospheres. Hydrogen propellant enters through a 0.5 foot diameter duct at the fore end (aka "Primary Circuit Inlet"). There are seven 0.4 foot diameter holes in the aft end for the engine nozzles, one at zero degrees off-axis, the other six at 60°. The pressure shell can be separated into two parts along the flange at the point of maximum diameter, to allow an engineer or waldo manipulator access to the engine interior. This point is also where the rear structural grid protrudes from the interior, this is where the engine bolts onto the structural frame of the spacecraft to transmit the engine thrust.

    If you look at the large blueprint, you will see that parts of the rear structural grid penetrate the cavities to support the end-caps of the quartz lightbulbs.

    Coolant system

    The plumbing for the coolant system is rather complicated (translation: I don't understand it all). Click for larger image. You can use this diagram along with the large blueprint to attempt to puzzle out what all the pipes are for. Basically the propellant enters the system through the "Primary circuit inlet" (at lower left of plumbing diagram, and in the pressure shell diagram above) and leaves the system via the "Propellant injection" arrow, where the propellant is heated by the lightbulbs in the cavity and jets out the exhaust nozzles. In between, the propellant frantically threads its way over every single other engine component in a desperate attempt to cool them off.

    Propellant flow overview

    In the blueprints you can see how the pipes that feed the propellant injectors are originally fed from horns over the graphite moderators. Which is exactly as per the plumbing diagram.

    This is my best guess at how the hydrogen propellant flows through the engine. It may be incorrect, use at your own risk. It starts with the green arrow at the left. This is the Primary circuit inlet at the nose of the engine, where the propellant enters the pressure shell. The pipe splits several ways (probably six ways, one for each outer cavity) and enters the base of the turbopump (arrows change color to Yellow).

    Pipe runs to the inner shell, then I hypothesize that there is a connection between the two bumps on the inner shell. Propellant runs to the inner pipe array just on top of the cavities, then it is injected into coolant channels in the beryllium oxide moderator around the tie rods. After cooling the beryllium, it spurts out and enters the base of the graphite moderator surrounding the hexagonal beryllium array (arrows change color to orange). It passes through coolant channels in the graphite, and emerges at the top into the collector horns. There it enters the outer pipe array above the inner pipe array.

    This feeds the three wedge shaped propellant injectors on each cavity. This injects the propellant around the edge of the transparent light bulbs (arrows change color to red). The propellant shoots aft while being heated by the thermal radiation from the light bulbs. The hot propellant then jets out the exhuast nozzles and thrust occurs.

    Cross sections

    Here are a set of cross sections through the cavities. The one on the left is zoomed in on the cavity interior, the other two gradually zoom out.

    Open Cycle

    Coaxial
    Coaxial
    Exhaust Velocity17,658 m/s
    Specific Impulse1,800 s
    Thrust17,800,000 N
    Thrust Power0.2 TW
    Mass Flow1,008 kg/s
    Total Engine Mass127,000 kg
    T/W14
    FuelFission:
    Uranium Hexafluoride
    ReactorGas Core
    Open-Cycle
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power1 kg/MW
    NASA-Lewis
    Thrust Power0.495GW
    Exhaust velocity22,000 m/s
    Thrust45,000 n
    Engine mass66 tonne
    T/W0.68

    Gaseous core coaxial flow fission / nuclear thermal rocket.

    Circa 1960 NASA-Lewis concept for a gas core nuclear rocket engine. Specific Impulse 2200 seconds (exhaust velocity 22,000 m/s). Thrust 45,000 newtons. Thrust to weight ratio 0.68 (engine mass 66,000 kilograms), reactor diameter 5 meters, overall reactor length 5 meters. The fuel would reach 20,000 degrees R, while the propellant would get to 10,000 degrees R. From The Unwanted Blog.

    Open Cycle
    Open Cycle
    Propulsion SystemGas Core NTR
    Exhaust Velocity35,000 m/s
    Specific Impulse3,568 s
    Thrust3,500,000 N
    Thrust Power61.2 GW
    Mass Flow100 kg/s
    Total Engine Mass200,000 kg
    FuelFission:
    Uranium Hexafluoride
    ReactorGas Core
    Open-Cycle
    RemassLiquid Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power3 kg/MW
    Engine mass30-200 tonne
    T/W 11.9 to 1.8
    Open Cycle 2
    Propulsion SystemGas Core NTR
    Exhaust Velocity50,000 m/s
    Specific Impulse5,097 s
    Thrust5,000,000 N
    Thrust Power0.1 TW
    Mass Flow100 kg/s
    FuelFission:
    Uranium Hexafluoride
    RemassLiquid Hydrogen
    Specific Power2 kg/MW
    Engine mass30-200 tonne
    T/W 17.0 to 2.5
    Open Cycle 3
    Thrust Power GW
    Exhaust velocity25,000 to 69,000 m/s
    Thrust19,600 to 108,000 n
    Engine mass40 to 110 tonne
    T/W0.05 to 0.10
    Operating Pressure400 to 2000 atm
    Open Cycle MAX
    Exhaust Velocity98,000 m/s
    Specific Impulse9,990 s
    Thrust3,000,000 N
    Thrust Power0.15 TW
    Mass Flow31 kg/s
    Total Engine Mass15,000 kg
    T/W20
    FuelFission:
    Uranium Hexafluoride
    Tanker
    Propulsion SystemGas Core NTR
    Open-cycle
    Exhaust Velocity35,316 m/s
    Specific Impulse3,600 s
    Thrust3,500,000 N
    Thrust Power61.8 GW
    Mass Flow99 kg/s
    FuelFission:
    Uranium Hexafluoride
    RemassLiquid Hydrogen
    Wet Mass433,000 kg
    Dry Mass268,000 kg
    Mass Ratio1.62 m/s
    ΔV16,943 m/s

    Gaseous core fission / nuclear thermal rocket AKA consumable nuclear rocket, plasma core, fizzer, cavity reactor rocket. The limit on NTR-SOLID exhaust velocities is the melting point of the reactor. Some engineer who obviously likes thinking "outside of the box" tried to make a liability into a virtue. They asked the question "what if the reactor was already molten?"

    Gaseous uranium is injected into the reaction chamber until there is enough to start a furious chain reaction. Hydrogen is then injected from the chamber walls into the center of this nuclear inferno where is flash heats and shoots out the exhaust nozzle.

    The trouble is the uranium shoots out the exhaust as well.

    The reaction is maintained in a vortex tailored to minimize loss of uranium out the nozzle. Fuel is uranium hexaflouride (U235F6), propellant is hydrogen. However, in one of the designs, U235 is injected by gradually inserting into the fireball a long rod of solid uranium. The loss of uranium in the exhaust reduces efficiency and angers environmentalists.

    Open-cycle gas core nuclear thermal rockets have a nominal core temperature of 21,000 K (37,340°F).

    In some designs the reaction chamber is spun like a centrifuge. This encourages the heavier uranium to stay in the chamber instead of leaking into the exhaust. This makes for a rather spectacular failure mode if the centrifuge's bearings seize.

    You can find more details here.

    If used for lift off it can result in a dramatic decrease in the property values around the spaceport, if not the entire county. An exhaust plume containing radioactive uranium is harmless in space but catastrophic in Earth's atmosphere.

    Amusingly enough, this is the best match for the propulsion system used in the TOM CORBETT: SPACE CADET books. However the books are sufficiently vague that it is possible the Polaris used a nuclear lightbulb. According to technical advisor Willy Ley, "reactant" is the hydrogen propellant, but the books imply that reactant is the liquid uranium.

    Porous Wall Gas Core Engine

    GAS CORE FISSION THERMAL ROCKETS

    The temperature limitations imposed on the solid core thermal rocket designs by the need to avoid material melting can be overcome, in principle, by allowing the nuclear fuel to exist in a high temperature (10,000 — 100,000 K), partially ionized plasma state. In this so-called "gaseous- or plasma-core" concept, an incandescent cylinder or sphere of fissioning uranium plasma functions as the fuel element. Nuclear heat released within the plasma and dissipated as thermal radiation from its surface is absorbed by a surrounding envelope of seeded hydrogen propellant that is then expanded through a nozzle to provide thrust. Propellant seeding (with small amounts of graphite or tungsten powder) is necessary to insure that the thermal radiation is absorbed predominantly by the hydrogen and not by the cavity walls that surround the plasma. With the gas core rocket (GCR) concept Isp values ranging from 1500 to 7000 s appear to be feasible [Ref. 26]. Of the various ideas proposed for a gas core engine, two concepts have emerged that have considerable promise: an open cycle configuration, where the uranium plasma is in direct contact with the hydrogen propellant, and a closed-cycle approach, known as the "nuclear light bulb engine" concept, which isolates the plasma from the propellant by means of a transparent, cooled solid barrier.

    Porous Wall Gas Core Engine

    The "open cycle," or "porous wall," gas core rocket is illustrated in Fig. 9. It is basically spherical in shape and consists of three solid regions: an outer pressure vessel, a neutron reflector/moderator region and an inner porous liner. Beryllium oxide (BeO) is selected for the moderator material because of its high operating temperature and its compatibility with hydrogen. The open cycle GCR requires a relatively high pressure plasma (500 — 2000 atm; 1 atm = 1.013 × 105 N/m2 ) to achieve a critical mass. At these pressures the gaseous fuel is also dense enough for the fission fragment stopping distance to be comparable to or smaller than the dimensions of the fuel volume contained within the reactor cavity. Hydrogen propellant, after being ducted through the outer reactor shell, is injected through the porous wall with a flow distribution that creates a relatively stagnant non-recirculating central fuel region in the cavity. A small amount of fissionable fuel (1/4 to 1 % by mass of the hydrogen flow rate) is exhausted, however, along with the heated propellant.

    Because the uranium plasma and hot hydrogen are essentially transparent to the high energy gamma rays and neutrons produced during the fission process, the energy content of this radiation (~7—10% of the total reactor power) is deposited principally in the solid regions of the reactor shell. It is the ability to remove this energy, either with an external space radiator or regeneratively using the hydrogen propellant, that determines the maximum power output and achievable Isp for the GCR engines. To illustrate this point, an open cycle engine with a thrust rating of 220 kN (50,000 lbf) is considered. We assume that 7% of reaction energy Prx reaches the solid, temperature-limited portion of the engine and that the remainder is converted to jet power at an isentropic nozzle expansion efficiency of ηj. Based on the realtionships between Isp, reactor power, and propellant flow rate (ṁp) given below.

    (ed note: elsewhere in this website, ṁ is called "m-dot")

    0.93·Prx(MW) = 4.9×10-6·F(N)·Isp(s) / ηj

    0.93·Prx(MW) = 4.9×10-5·ṁp(kg/s)·Isp2(s) / ηj

    a 5000 s engine generating 7500 MW of reactor power will require a flow rate of 4.5 kg/s at rated thrust. If the hydrogen is brought into the cavity at a maximum overall operating temperature of 1400 K, no more than 1.2% of the total reactor power (~17% of the neutron and gamma power deposited in the reactor structure) can be removed regeneratively (ṁp cp ΔT ≈ 90 MW). Total removal requires either (1) operating the sold portions of the engine at unrealistically high temperatures (>11,000 K at ṁp = 4.5 kg/s) or (2) increasing the propellant flow rate substantially to 36.8 kg/s (at 1400 K), which reduces the engine's Isp to 1750 s. "Closed cooling cycle" space radiator systems have been proposed [Ref. 27] as a means of maintaining the GCR's operational flexibility. With such a system, adequate engine cooling is possible even during high Isp operation when the hydrogen flow is reduced. Calculations performed by NASA/Lewis Research Center [Ref. 28] indicate that specific impulses ranging from 3000 to 7000 s could be attained in radiator-cooled, porous wall gas core engines.

    The performance and engine characteristics for a 5000 s class of open cycle GCRs are summarized in Table 4 for a range of thrust levels. The diameter of the reactor cavity and the thickness of the external reflector/moderator region are fixed at 2.44 m and 0.46 m, respectively, which represents a near-optimum engine configuration. The engine weight (Mw) is composed primarily of the pressure vessel (Mpv); radiator (Mrad); and moderator (Mmod).

    Table 4
    Characteristics of 5000 s Porous Wall Gas Core Rocket Engines
    Thrust


    F(kN)
    Engine
    Power

    Prx(MW)
    Radiator
    Power

    Prad(MW)a
    Engine
    Weight

    Mw(mT)
    Pressure
    Vessel
    Weight
    Mpv(mT)b
    Radiator
    Weight

    Mrad(mT)c
    Moderator
    Weight

    Mmod(mT)d
    Alpha


    αp(kW/kg)
    Thrust
    to
    Weight
    F/Mw
    2275043.552.3106.33610.34.3×10-2
    4415008761.61312.63617.57.3×10-2
    11037502188618323631.30.13
    220750043512324633643.80.18
    44015,000870193311263655.90.23
    1. For a hydrogen cavity inlet temperature of 1400 K and a heat deposition rate that is 7% of the reactor power, the ratio of radiated to total reactor power is a constant equal to 5.8%.
    2. The weight of the spherical pressure vessel is based on a strength-to-density value of 1.7×l05 N-m/kg [Ref. 29] which Is characteristic of high strength steels.
    3. Used in these estimates is a radiator specific mass of 145 kg/MW [Ref. 28] which is based on a heat rejection temperature of 1225 K and a radiator weight per unit surface area of 19 kg/m2
    4. Density of BeO is 2.96 mT/m3.

    By fixing the engine geometry in Table 4 the mass of the BeO moderator remains constant at 36 mT. However, the pressure vessel and radiator weights are both affected by the thrust level. While the radiator weight increases in proportion to the extra power that must be dissipated at higher thrust, the reason for the increase in pressure vessel weight is slightly more subtle. For a constant Isp engine an increase in thrust is achieved by increasing both the reactor power and hydrogen flow rate. In order to radiatively transfer this higher power to the propellant, the uranium fuel temperature increases, necessitating an increase in reactor pressure to maintain a constant critical mass in the engine. Accommodating this increased pressure leads to a heavier pressure vessel. (In going from 22 kN to 440 kN, the engine pressure rises from 570 atm to 1780 atm).

    As Table 4 illustrates, the moderator is the major weight component at lower thrust levels (<110 kN) while the radiator becomes increasingly more important at higher thrust. At thrust levels of 220 kN and above, the radiator accounts for more than 50% of the total engine weight. There is therefore a strong incentive to develop high temperature (~1500 K) liquid metal heat pipe radiators that could provide significant weight reductions in the higher thrust engines.

    Table 4 also shows an impressive range of specific powers (alphas) and engine thrust-to-weight ratios for the thrust levels examined. The F/Mw ratio for the 22 kN engine is over two orders of magnitude higher than the 5000 s nuclear-powered MPD electric propulsion system proposed in the Pegasus study [Ref. 30]. For manned Mars missions the higher acceleration levels possible with the GCR can lead to significant (factor of 5) reductions in trip time compared to the Pegasus system.

    Vortex Confined
    Vortex Confined
    Exhaust Velocity19,620 m/s
    Specific Impulse2,000 s
    Thrust50,400 N
    Thrust Power0.5 GW
    Mass Flow3 kg/s
    Total Engine Mass114,116 kg
    T/W0.04
    Frozen Flow eff.75%
    Thermal eff.70%
    Total eff.53%
    FuelFission:
    Uranium Hexafluoride
    ReactorGas Core
    Vortex Confined
    RemassSeeded Hydrogen
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorNozzle
    Specific Power231 kg/MW

    The hotter the core of a thermodynamic rocket, the better its fuel economy. If it gets hot enough, the solid core vaporizes.

    A vapor core rocket mixes vaporous propellant and fuel together, and then separates the propellant out so it can be expelled for thrust. Energy is efficiently transferred from fuel to propellant by direct molecular collision, radiative heat, and direct reaction fragment deposition.

    The open-cycle arrangement illustrated accomplishes this by spinning the plasma mixture in a vortex maintained by tangential injection of preheated propellant from the reactor walls. The denser material is held to the outside of the cylindrical reactor vessel by centrifugal force. The fuel is subsequently cooled in a heat exchanger and recirculated for re-injection at the forward end of the reactor, while the propellant is exhausted at high velocity.

    The plasma source can be fission, antimatter, or fusion.

    For fission reactions, the outer annulus of the vortex is high-density liquid uranium fuel, and the low-density propellant is bubbled through to the center attaining temperatures of up to 18500 K. A BeO moderator returns many reaction neutrons to the vortex. Prompt feedback actuators maintain a critical fuel mass in spite of the turbulent flow of water or hydrogen propellant. Since the core has attained meltdown, reaction rates must be maintained by fuel density variation rather than with control rods or drums.

    For antimatter reactions, swirling liquid tungsten (about 4 cm thick) is used instead of uranium, for absorbing anti-protons.

    For fusion reactions, it is the propellant that is cooler and higher in density, and thus it is the reacting fuel ball that resides at the center of the vortex.

    N. Diaz of INSPI, 1990.

    From High Frontier by Philip Eklund
    Nuclear Salt Water
    NSWR
    20% UTB
    Exhaust Velocity66,000 m/s
    Specific Impulse6,728 s
    Thrust12,900,000 N
    Thrust Power425.7 GW
    Mass Flow195 kg/s
    Total Engine Mass33,000 kg
    T/W40
    FuelFission:
    Uranium Tetrabromide
    ReactorGas Core
    Open-Cycle
    RemassWater
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorPusher Plate
    Specific Power0.08 kg/MW
    90% UTB
    Exhaust Velocity4,700,000 m/s
    Specific Impulse479,103 s
    Thrust13,000,000 N
    Thrust Power30.6 TW
    Mass Flow3 kg/s

    This concept by Dr. Zubrin is considered far-fetched by many scientists. The fuel is a 2% solution of 20% enriched Uranium Tetrabromide in water. A Plutonium salt can also be used.

    Just to make things clear, there are two percentages here. The fuel is a 2% solution of uranium tetrabromide and water. That is, 2 molecules of uranium tetrabromide per 100 molecules of water.

    But the uranium tetrabromide can be 20% enriched. This means that out of every 100 atoms of uranium (or molecules of uranium tetrabromide), 20 are fissionable Uranium-235 and 80 are non-fissionable uranium. If it is 90% enriched, then 90 atoms are Uranium-235 and 10 atoms are non-fissionable. As a side note, 90% enriched is considered "weapons-grade".


    The fuel tanks are a bundle of pipes coated with a layer of boron carbide neutron damper. The damper prevents a chain reaction. The fuel is injected into a long cylindrical plenum pipe of large diameter, which terminates in a rocket nozzle. Free of the neutron damper, a critical mass of uranium soon develops. The energy release vaporizes the water, and the blast of steam carries the still reacting uranium out the nozzle.

    It is basically a continuously detonating Orion type drive with water as propellant. Although Zubrin puts it like this:

    As the solution continues to pour into the plenum from the borated storage pipes, a steady-state conditions of a moving detonating fluid can be set up within the plenum.

    He also notes that it is preferable to subject your spacecraft to a steady acceleration (as with the NSWR) as opposed to a series of hammer-blow accelerations (as with Orion).

    The controversy is over how to contain such a nuclear explosion. Zubrin maintains that skillful injection of the fuel can force the reaction to occur outside the reaction chamber. He says that the neutron flux is concentrated on the downstream end due to neutron convection. Other scientists are skeptical.

    Naturally in such a spacecraft, damage to the fuel tanks can have unfortunate results (say, damage caused by hostile weapons fire). Breach the fuel tubes and you'll have a runaway nuclear chain reaction on your hands. Inside your ship.

    The advantage of NSWR is that this is the only known propulsion system that combines high exhaust velocity with high thrust. The disadvantage is that it combines many of the worst problems of the Orion and Gas Core systems. For starters, using it for take-offs will leave a large crater that will glow blue for several hundred million years, as will everything downwind in the fallout area.

    Zubrin calculates that the 20% enriched uranium tetrabromide will produce a specific impulse of about 7000 seconds (69,000 m/s exhaust velocity), which is comparable to an ion drive. However, the NSWR is not thrust limited like the ion drive. Since the NSWR vents most of the waste heat out the exhaust nozzle, it can theoretically produce jet power ratings in the thousands of megawatts. Also unlike the ion drive, the engine is relatively lightweight, with no massive power plant required.

    Zubrin suggests that a layer of pure water be injected into the plenum to form a moving neutron reflector and to protect the plenum walls and exhaust nozzle from the heat. One wonders how much protection this will offer.

    Zubrin gives a sample NSWR configuration. It uses as fuel/propellant a 2% (by number) uranium bromide aqueous solution. The uranium is enriched to 20% U235. This implies that B2 = 0.6136 cm-2 (the material buckling, equal to vΣfa)/D) and D = 0.2433 cm (diffusion coefficent).

    Radius of the reaction plenum is set to 3.075 centimeters. this implies that A2 = 0.6117 cm-2 and L2 = 0.0019. Since exponential detonation is desired, k2 = 2L2 = 0.0038 cm-2. Then k = U / 2D = 0.026 cm-1 and U = 0.03.

    If the velocity of a thermal neutron is 2200 m/s, this implies that the fluid velocity needs to be 66 m/s. This is only about 4.7% the sound speed of room temperature water so it should be easy to spray the fuel into the plenum chamber at this velocity.

    The total rate of mass flow through the plenum chamber is about 196 kg/s.

    Complete fission of the U235 would yield about 3.4 x 1012 J/kg. Zubrin assumes a yield of 0.1% (0.2% at the center of the propellant column down to zero at the edge), which would not affect the material buckling during the burn. This gives an energy content of 3.4 x 109 J/kg.

    Assume a nozzle efficiency of 0.8, and the result is an exhaust velocity of 66,000 m/s or a specific impules of 6,7300 seconds. The total jet power is 427 gigawatts. The thrust is 12.9 meganewtons. The thrust-to-weight ratio will be about 40, which implies an engine mass of about 33 metric tons.

    For exponetial detonation, kz has to be about 4 at the plenum exit. Since k = 0.062 cm-1, the plenum will have to be 65 cm long. The plenum will be 65 cm long with a 3.075 cm radius, plus an exhaust nozzle.

    Zubrin then goes on to speculate about a more advanced version of the NSWR, suitable for insterstellar travel. Say that the 2% uranium bromide solution used uranium enriched to 90% U235 instead of only 20%. Assume that the fission yield was 90% instead of 0.1%. And assume a nozzle efficency of 0.9 instead of 0.8.

    That would result in an exhaust velocity of a whopping 4,725,000 m/s (about 1.575% c, a specific impulse of 482,140 seconds). In a ship with a mass ratio of 10, it would have a delta V of 3.63% c. Now you're talkin...

    Ken Burnside: In my game universe, the engineers call the pumps that feed Uranium Tetrabromide solution into the reaction chamber "Wileys", reputedly after the engineer who first made them safe to use and maintain.

    More than likely, it's after the coyote of the same name...


    Winchell Chung: An appropriate name for what are basically atomic squirt-guns.

    From a thread in SFConSim-l (2002)
    Zubrin NSWR
    Zubrin NSWR
    Exhaust Velocity78,480 m/s
    Specific Impulse8,000 s
    Thrust8,696,900 N
    Thrust Power0.3 TW
    Mass Flow111 kg/s
    Total Engine Mass495,467 kg
    T/W1.79
    Frozen Flow eff.80%
    Total eff.80%
    FuelFission:
    Uranium Tetrabromide
    ReactorGas Core
    Open-Cycle
    RemassWater
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorPusher Plate
    Specific Power1.45 kg/MW

    The illustration shows the vision of Robert Zubrin: a rocket riding on a continuous controlled nuclear explosion just aft of a nozzle/reaction chamber.

    The propellant is water, containing dissolved salts of fissile uranium or plutonium. These fuel-salts are stored in a tank made from capillary tubes of boron carbonate, a strong structural material that strongly absorbs thermal neutrons, preventing the fission chain reaction that would otherwise occur.

    To start the engine, the salt-water is pumped from the fuel tank into an absorber-free cylindrical nozzle. The salt-water velocity is adjusted as it exits the tank so that the thermal neutron flux peaks sharply in the water-cooled nozzle.

    At critical mass (around 50 kg of salt water), the continuous nuclear explosion produces 427 GWth, obtaining a thrust of 8600 kN and a specific impulse of 8 ksec at a thermal efficiency of 99.8% (with open-cycle cooling). Overall efficiency is 80%.

    Robert Zubrin, "Nuclear Salt Water Rockets: High Thrust at 10,000 sec ISP," Journal of the British Interplanetary Society 44, 1991.


    You need much more propellant than fuel, 22,000 times more in the case of the Zubrin without open cycle cooling, and 44,000 times more if open cycle cooling is used.

    The Zubrin drive exhaust (without open cycle cooling) contains 108 kg/sec of water, but only about 5 grams/sec of uranium.

    (This is from a quick calculation: mass flow equals the Zubrin thrust (8.7 meganewtons) divided by the exit velocity (80 km/sec) = 108 kg/sec. But the fissioning energy can be estimated from the Zubrin total power of 427 GW divided by the energy content of Uranium 235 of 83 TJ/kg.)


    Dr. Zubrin responded, and he defends the performance of the Zubrin drive as depicted in the game (as high thrust & high specific impulse rocket with low mass and low radiators).

    1). In U235 fission, only about 2% of the energy goes into neutrons (unlike D-T fusion).

    2). The design uses a pusher plate or open nozzle, like an Orion drive. Or magnetic confinement (since most of the energy is released as a plasma). Therefore, the opportunity to absorb heat is low.

    3) Many of the neutrons that are intercepted would sail through the pusher plate, rather than be absorbed as waste heat.

    4) No lithium should be in the outer water, because this would poison the fission reactions.

    5). Because the design does not use a heat engine cycle, the radiators could be far hotter than ones in the game. He suggested graphite at 2500 K°. That would drop the required radiating area by a factor of 40 (2.5 to the fourth power), which means that the radiator could be the first wall itself.

    Dr. Zubrin went on to say the chief disadvantage is the expense of the fuel (like He3-D and antimatter drives).

    Philip Eklund, from a discussion on the High Frontier Yahoo group about the NSWR drive in High Frontier

    "So anyway, we were passing through the outer Kirkwood Gap, totally the a** end of nowhere. I'm trying to catch some rack time, XO has the conn, nice boring trip to Europa." The CO of U.S.N.A.S. Saskatchewan tipped back another shot of Scotch and continued his story. "Totally routine, right? No problems at all. So then, all of the sudden, the whole ship gets racked. Meteoroid. Big one, too, maybe a centimeter across."

    The captains seated around the table, two Americans including Fitzthomas, an Indian, three Chinese, and the South African, all clucked and groaned.

    "Well, we got lucky and it missed the crew compartment, but by the time I get to command the chief engineer is screaming over the intercom that it holed tank one, busted three tubes, and we've got nuke juice pooling and we have to dump the tank. Problem is, we're running at top speed and if we dump the tank, we don't have enough propellant to stop at Europa. We'd have to ride all the way out to Neptune, sling around, and hope someone from the inner solar system has dispatched a tanker to intercept us on the return trip, and we don't have near enough consumables for that."

    "So what did you do, mate?" said the South African.

    "I told the chief he had to fix the tank or we'd all starve before we could stop the damn ship. Well, he screams some more that we don't have time, and I tell him his choice is fix the tank or die real fast in a runaway, because we're not going to die slow in the void. So he grabs a crew, stuffs them into suits, and crawls out onto the tank. They punch some holes in it to let the juice drain instead of pool, but it's still leaking like a f***** and the water's evaporating and leaving uranium crusted all over everything. So he radios command and says, 'It's still leaking, and all this uranium crud is going to accumulate into a critical mass somewhere, so we still have to drop the tank.' Meteoroid busted open three valves, you see. No way to stop the leak. And I tell him again, that's no good, and by now astrogation has confirmed it and the XO has tallied up the consumables and I know for sure we don't have enough for an unscheduled trip to Neptune.

    "So he says something about how he's not a miracle worker, and I tell him he damn well has to be. Lo and behold, he and his crew go ahead and do something crazy and it works."

    "What was that?" said the South African.

    "They take torches to the tank. The plug up the broken pipes as best they can, and then they go ahead and cut away the smashed cells. Just cut it off and jettison it into space, and suddenly the propellant that's still leaking is leaking right into space. We have lousy flow through the tank and the braking burn is going to be real tricky, but we can make Europa. I put the chief up for a commendation medal for figuring that out on the fly and saving our asses."

    The other captains nodded their approval at the chief's quick thinking. Good chiefs prevented accidents; great ones prevented disasters.

    "Is the chief's name Mr. Scott, by any chance?" said one of the Chinese captains.


    Commander George Allen, New Jersey's full blooded Cherokee XO, drifted into the command deck from astrogation, where he'd been monitoring the final approach to Hektor. He took his place at the copilot station and put on his headset. Fitzthomas toggled his direct channel to Allen's station.

    "How was the approach?"

    "We wasted too much propellant before the chain reaction started. I think Pennai should inspect the nozzles and pumps before we get underway again."

    "What does Pennai say about it?"

    Pause. "Pennai thinks the fuel is dirty."

    "Is it?"

    "It was certified 90% enriched at Roosevelt Station."

    "Is there any way to test it here?"

    "No sir. Not without a centrifuge."

    "How does Pennai know, then?"

    "Some engineering technobabble about neutron flux and reaction rate. I couldn't follow a tenth of what she said."

    Fitzthomas considered that. "Have her inspect the pumps and nozzle alignment. If they pass, then we might have a fuel problem."


    "Captain," said Allen, "Thought you'd like to know: Pennai just inspected the entire fuel line. Everything there is in order."

    "So what are you telling me, George?"

    "I think we have dirty fuel."

    "What's her recommendation?"

    "She wants to drain the tanks and top up with the good stuff. But I can't—"

    "Write that request, I know. The CO has to. Where's Pennai now?"

    "Racked out. She has the midwatch tonight."

    "After her watch tonight, she has four days of leave."

    "Sir, she's supposed to be OOW all day Wednesday."

    "I'll take that shift. She was right, we were wrong. She deserves to be rewarded. When I get back I'll write up a request and have it to the fuelmaster by tomorrow AM."


    "Do you have to return to your ship?"

    "Yeah. Dirty fuel, God damn it. Wait until I get my hands on the fuelmaster at Roosevelt."


    (Admiral Castro said) "Anyway, I saw your chief engineer's report. I passed it back to Fleet. The fuelmaster at Roosevelt Station is going to have a lousy day tomorrow. There's also a bulletin going out to the entire fleet. Everyone who tanked up at Roosevelt near the same time you did should keep a close eye on his reaction rate. Your Lieutenant Pennai might be up for a commendation letter in her file."


    Duvalier left Ortiz main engineering and vaulted down the access tube to the reactor room. The tube ran down the ship's spine, surrounded by megaliters of water enriched with uranium salts in highly complex tanks made of neutron absorbing material. In his head, he knew the tube was the safest part of the ship, shielded from the worst the universe could throw at it by dozens of meters of water. In his head, he knew the fuel, so long as it didn't pool into a critical mass somewhere in the thousands of kilomters of pipes on all sides of him, emitted only low intensity alpha rays which couldn't penetrate his own skin, let alone the aluminum skin of the pressure tube. It was all perfectly safe, so far as anything in space could be safe. He knew that in his head.

    His b***s, however, hadn't gotten the memo. His testicles tried to crawl up into his body every time he climbed through the hatch.

    From The Last Great War by Matthew Lineberger (not yet published)

    Fission Fragment

    Fission Fragment

    George Chapline
    Exhaust velocity980,000 m/s

    All of the other nuclear thermal rockets generate heat with nuclear fission, then transfer the heat to a working fluid which becomes the reaction mass. The transfer is always going to be plagued by inefficiency, thanks to the second law of thermodynamics. What if you could eliminate the middleman, and use the fission heat directly with no transfer?

    That what the fission fragment rocket does. It uses the hot split atoms as reaction mass. The down side is that due to the low mass flow, the thrust is minuscule. But the up side is that the exhaust velocity is 5% the speed of light! 15,000,000 kilometers per second, that's like a bat out of hell. With that much exhaust velocity, you could actually have a rocket where less than 50% of the total mass is propellant (i.e., a mass ratio below 2.0).

    Dr. Chapline's design use thin carbon filaments coated with fission fuel (coating is about 2 micrometers thick). The filaments radiated out from a central hub, looking like a fuzzy vinyl LP record. These revolving disks were spun at high speed through a reactor core, where atoms of nuclear fuel would undergo fission. The fission fragments would be directed by magnetic fields into an exhaust beam.

    The drawback of this design is that too many of the fragments fail to escape the fuel coat (which adds no thrust but does heat up the coat) and too many hit the carbon filaments (which adds no thrust but does heat up the filaments). This is why the disks spin at high speed, otherwise they'd melt.

    Dusty Plasma
    550AU
    Thrust22 N
    Thrust Power0.2 GW
    Mass Flow1.00e-06 kg/s
    T/W2.49e-04
    Specific Power55 kg/MW
    0.5LY
    Thrust344 N
    Thrust Power2.6 GW
    Mass Flow2.30e-05 kg/s
    T/W4.00e-03
    Specific Power3 kg/MW
    All
    Exhaust Velocity15,000,000 m/s
    Specific Impulse1,529,052 s
    Total Engine Mass9,000 kg
    FuelFission:
    Uranium 235
    ReactorGas Core
    MHD Choke
    RemassReaction
    Products
    Remass AccelFission-Fragment
    Thrust DirectorMagnetic Nozzle

    Rodney Clark and Robert Sheldon solve the problem with their Dusty plasma bed reactor (report).

    You take the fission fuel and grind it into dust grains with an average size of 100 nanometers (that is, about 1/20th the thickness of the fuel coating in dr. Chapline's design). This does two things [A] most of the fragments escape and [B] the dust particles have such a high surface to volume ratio that heat (caused by fragments which fail to escape) readily dissipates, preventing the dust particles from melting.

    The dust is suspended in the center of a reaction chamber whose walls are composed of a nuclear moderator. Power reactors will use beryllium oxide (BeO) as a moderator, but that is a bit massive for a spacecraft. The ship will probably use lithium hydride (LiH) for a moderator instead, since is only has one-quarter the mass. Probably about six metric tons worth. The dust is suspended electrostatically or magnetically by a containment field generator. The dust is heated up by radio frequency (RF) induction coils. The containment field generator will require superconductors, which will probably require a coolant system of its own.

    The dust particles are slow and are relatively massive, while the fission fragments are fast and not very massive at all. So the magnetic field can be tailored so it holds the dust but allows the fission fragments to escape. Magnetic mirrors ensure that fragments headed the wrong way are re-directed to the exhaust port.

    One valuable trick is that you can use the same unit for thrust or to generate electricity. Configure the magnetic field so that the fragments escape "downward" through the exhaust port and you have thrust. Flip a switch to change the magnetic field so that the fragments escape upward into deceleration and ion collection electrodes and you generate electricity. As a matter of fact, it is go efficient at generrating electricity that researchers are busy trying to adapt this for ground based power plants. But I digress.

    The dust is only sufficient for a short period of critical nuclear reaction so it must be continuously replenished. The thermal energy released by fission events plus heat from collisions between fission fragments and dust grains create intense heat within the dust cloud. Since there is no core cooling flow, the reactor power is limited to the temperature at which the dust can radiatively cool itself without vaporizing. The interior of the reaction chamber walls will protected by a mirrored (95% reflection) heat shield attached to a heat radiator. The outer moderator layer will have its own heat shield.

    Clark and Sheldon roughed out a propulsion system. It had six tons for the moderator, 2 tons for radiators and liquid metal cooling, 1 ton for magnets, power recovery, and coils, for a grand total of 9 tons. The reaction chamber will be about 1 meter in diameter and 10 meters long. The moderator blanket around the chamber will be about 40 centimeters thick. The thrust is a function the size of the cloud of fissioning dust, and is directly related to the power level of the reactor. There is a limit to the maximum allowed power level, set by the coolant system of the reaction chamber. Clark and Sheldon estimate that only about 46% of the fission fragments provide thrust while the rest are wasted. See the report for details.

    In the table, the 550AU engine is for a ten year journey to the Solar gravitational lensing point at 550 astronomical units (so you can use the sun as a giant telescope lens). The 0.5LY engine is for a thirty year trip to the Oort cloud of comets. These are constant acceleration brachistochrone trajectories, the 550AU mission will need a reactor power level of 350 MW and the 0.5LY mission will need 5.6 GW. Don't forget that the engine power is only 46% efficient, that's why the table thrust values are lower.

    Werka FFRE
    First Generation
    Exhaust Velocity5,170,000 m/s
    Specific Impulse527,013 s
    Thrust43 N
    Thrust Power0.1 GW
    Mass Flow8.00e-06 kg/s
    Total Engine Mass113,400 kg
    T/W3.90e-05
    FuelFission:
    Plutonium 239
    ReactorGas Core
    MHD Choke
    RemassReaction
    Products
    Remass AccelFission-Fragment
    Thrust DirectorMagnetic Nozzle
    Specific Power1,020 kg/MW
    HOPE FFRE
    Propulsion SystemWerka FFRE
    Wet Mass303,000 kg
    Dry Mass295,000 kg
    Mass Ratio1.03 m/s
    ΔV138,336 m/s

    Robert Werka has a more modest and realistic design for his fission fragment rocket engine (FFRE). He figures that a practical design will have an exhaust velocity of about 5,200,000 m/s instead of his estimated theoretical maximum of 15,000,000 m/s. His lower estimate is still around 1.7% the speed of light so we are still talking about sub 2.0 mass ratios. Collisions between fission fragments and the dust particles is responsible for the reduction in exhaust velocity.

    Incidentally the near relativistic exhaust velocity reduces radioactive contamination of the solar system. The particles are traveling well above the solar escape velocity (actually they are even faster than the galactic escape velocity) so all the radioactive exhaust goes shooting out of the solar system at 0.017c.

    The dusty fuel is nanometer sized particles of slightly critical plutonium carbide, suspended and contained in an electric field. A moderator of deuterated polyethylene reflects enough neutrons to keep the plutonium critical, while control rods adjust the reaction levels. The moderator is protected from reaction chamber heat by a heat shield, an inner layer composed of carbon-carbon to reflect infrared radiation back into the core. The heat shield coolant passes through a Braydon cycle power generator to create some electricty, then the coolant is sent to the heat radiator.

    The details of Werka's initial generation FFRE can be found in the diagram below. The reaction chamber is about 5.4 meters in diameter by 2.8 meters long. The magnetic nozzle brings the length to 11.5 meters. The fuel is uranium dioxide dust which melts at 3000 K, allowing a reactor power of 1.0 GW. It consume about 29 grams of uranium dioxide dust per hour (not per second). Of the 1.0 GW of reactor power, about 0.7 GW of that is dumped as waste heat through the very large radiators required.

    The second most massive component is the magnetic mirror at the "top" of the reaction chamber. Its purpose is to reflect the fission fragments going the wrong way so they turn around and travel out the exhaust nozzle. Surrounding the "sides" of the reaction chamber is the collimating magnet which directs any remaining wrong-way fragments towards the exhaust nozzle. The exhaust beam would cause near-instantaneous erosion of any material object (since it is electrically charged, relativistic, radioactive grit). It is kept in bounds and electrically neutralized by the magnetic nozzle cage.

    Fission Sail

    Fission Sail

    Antimatter-Driven Sail

    The sail is made of graphite and carbon-carbon fiber, infused with a tiny amount of uranium. It is subjected to a misting of antiprotons. These induce uranium atoms to fission, with the recoil pushing the sail. Since this is nuclear powered, the sail does not have to be kilometers in diameter, five meters will do. 30 miligrams of antiprotons could push the sail to the Kuiper Belt.

    Pulse

    Orion

    Fission Orion
    Exhaust Velocity43,000 m/s
    Specific Impulse4,383 s
    Thrust263,000 N
    Thrust Power5.7 GW
    Mass Flow6 kg/s
    Total Engine Mass200,000 kg
    T/W0.13
    FuelFission:
    Uranium 235
    ReactorPulse Unit
    RemassTungsten
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorPusher Plate
    Specific Power35 kg/MW
    Fusion Orion
    Exhaust Velocity73,000 m/s
    Specific Impulse7,441 s
    Thrust292,000 N
    Thrust Power10.7 GW
    Mass Flow4 kg/s
    Total Engine Mass200,000 kg
    T/W0.15
    FuelDeuterium-Deuterium
    Fusion
    Specific Power19 kg/MW
    1959 Orion 1st Gen
    Thrust Power1,600 GW
    Exhaust velocity39,000 m/s
    Thrust80,000,000 n
    Engine mass1,700 tonne
    T/W >1.0yes
    1959 Orion 2nd Gen
    Thrust Power24,000 GW
    Exhaust velocity120,000 m/s
    Thrust400,000,000 n
    Engine mass3,250 tonne
    T/W >1.0yes
    ORION USAF 10m
    Exhaust Velocity32,900 m/s
    Specific Impulse3,354 s
    Thrust2,000,000 N
    Thrust Power32.9 GW
    Mass Flow61 kg/s
    Total Engine Mass107,900 kg
    T/W2
    Wet Mass475,235 kg
    Dry Mass180,975 kg
    Mass Ratio2.63 m/s
    ΔV31,763 m/s
    Specific Power3 kg/MW
    ORION 4K ton battleship
    Exhaust Velocity39,000 m/s
    Specific Impulse3,976 s
    Thrust80,000,000 N
    Thrust Power1.6 TW
    Mass Flow2,051 kg/s
    Total Engine Mass1,700,000 kg
    T/W4.80
    Specific Power1.09 kg/MW
    ΔV 10 km/s
    Wet Mass4,000,000 kg
    Dry Mass3,100,000 kg
    Mass Ratio1.29 m/s
    ΔV9,941 m/s
    ΔV 21 km/s
    Wet Mass4,000,000 kg
    Dry Mass2,353,000 kg
    Mass Ratio1.70 m/s
    ΔV20,694 m/s
    ΔV 30 km/s
    Wet Mass4,000,000 kg
    Dry Mass1,852,000 kg
    Mass Ratio2.16 m/s
    ΔV30,031 m/s
    ORION 10k ton adv
    Exhaust Velocity120,000 m/s
    Specific Impulse12,232 s
    Thrust400,000,000 N
    Thrust Power24.0 TW
    Mass Flow3,333 kg/s
    Total Engine Mass3,250,000 kg
    T/W13
    Specific Power0.14 kg/MW
    ΔV 10 km/s
    Wet Mass10,000,000 kg
    Dry Mass9,199,000 kg
    Mass Ratio1.09 m/s
    ΔV10,019 m/s
    ΔV 15.5 km/s
    Wet Mass10,000,000 kg
    Dry Mass8,772,000 kg
    Mass Ratio1.14 m/s
    ΔV15,722 m/s
    ΔV 20 km/s
    Wet Mass10,000,000 kg
    Dry Mass8,403,000 kg
    Mass Ratio1.19 m/s
    ΔV20,880 m/s
    ΔV 30 km/s
    Wet Mass10,000,000 kg
    Dry Mass7,813,000 kg
    Mass Ratio1.28 m/s
    ΔV29,616 m/s
    ΔV 100 km/s
    Wet Mass10,000,000 kg
    Dry Mass4,348,000 kg
    Mass Ratio2.30 m/s
    ΔV99,944 m/s
    Orion MAX
    Exhaust Velocity9,800,000 m/s
    Specific Impulse998,981 s
    Thrust8,000,000 N
    Thrust Power39.2 TW
    Mass Flow0.82 kg/s
    Total Engine Mass8,000 kg
    T/W102
    FuelProton-Proton
    Fusion
    RemassTungsten
    Specific Power2.04e-04 kg/MW

    Orion AKA "old Boom-boom" is the ultimate consumable nuclear thermal rocket, based on the "firecracker under a tin can" principle. Except the tin can is a spacecraft and the firecracker is a nuclear warhead.

    This concept has the spacecraft mounted with shock absorbers on an armored "pusher plate". A stream of small (5 to 15 kiloton) fission or fusion explosives are detonated under the plate to provide thrust. While you might find it difficult to believe that the spacecraft can survive this, you will admit that this will give lots of thrust to the spacecraft (or its fragments). On the plus side, a pusher plate that can protect the spacecraft from the near detonation of nuclear explosives will also provide dandy protection from any incoming weapons fire. On the minus side I can hear the environmentalists howling already. It will quite thoroughly devastate the lift-off site, and give all the crew bad backs and fallen arches. And they had better have extra-strength brassieres and athletic supporters.

    Mathematician Richard Courant viewed an Orion test and said "Zis is not nuts, zis is super-nuts."


    This section is about the Orion propulsion system. If you want all the hot and juicy details about various versions of Orion spacecraft go here.

    Please note that Orion drive is pretty close to being a torchship, and is not subject to the Every gram counts rule. It is probably the only torchship we have the technology to actually build today.


    If you want the real inside details of the original Orion design, run, do not walk, and get a copies the following issues of of Aerospace Projects Review: Volume 1, Number 4, Volume 1, Number 5, and Volume 2, Number 2. They have blueprints, tables, and lots of never before seen details.

    If you want your data raw, piled high and dry, here is a copy of report GA-5009 vol III "Nuclear Pulse Space Vehicle Study - Conceptual Vehicle Design" by General Atomics (1964). Lots of charts, lots of graphs, some very useful diagrams, almost worth skimming through it just to admire the diagrams.


    The sad little secret about Orion is that the mission it is best suited for is boosting heavy payloads into orbit. Which is exactly the mission that the enviromentalist and the nuclear test ban treaty will prevent. Orion has excellent thrust, which is what you need for lift-off and landing. Unfortunately its exhaust velocity is pretty average, which is what you need for efficient orbit-to-orbit maneuvers.

    Having said that, there is another situation where high thrust is desirable: a warship jinking to make itself harder to be hit by enemy weapons fire. It is also interesting to note that the Orion propulsion system works very well with the bomb-pumped laser weapons system.


    Each pulse unit is a tiny nuclear bomb, encased in a "radiation case" that has a hole in the top. A nuclear blasts is initially mostly x-rays. The radiation case is composed of a material that his opaque to x-rays (depleted uranium). The top hole thus "channels" the flood of x-rays in an upwards direction (at least in the few milliseconds before the bomb vaporizes the radiation case).

    The channeled x-rays then strike the "channel filler" (beryllium oxide). The channel filler transforms the atomic fury of x-rays into an atomic fury of heat.

    Lying on top of the channel filler is the disc of propellant (tungsten). The atomic fury of heat flashes the tungsten into a jet of ionized tungsten plasma, traveling at high velocity (in excess of 1.5 × 105 meters per second). This crashes into the pusher plate, accelerating the spacecraft. It crashes hard. You will note that there are two stages of shock absorbers between the pusher plate and the spacecraft, preventing instant crew death.

    The ratio of beryllium oxide to tungsten is 4:1.

    The jet is confined to a cone about 22.5 degrees (instead of in all directions). The detonation point is positioned such that the 22.5 cone exactly covers the diameter of the pusher plate. The idea is that the wider the area of the cone, the more spread out the impulse will be, and the larger the chance that the pusher plate will not be utterly destroyed by the impulse.

    It is estimated that 85% of the energy of the nuclear explosion can be directed in the desired direction. The pulse units are popped off at a rate of about one per second. A 5 kiloton charge is about 1,152 kg. The pulses are so brief that there is no appreciable "neutron activation", that is, the neutron from the detonations do not transmute parts of the spacecraft's structure into radioactive elements. This means astronauts can exit the spacecraft and do maintenance work shortly after the pulse units stop detonating.

    The device is basically a nuclear shaped charge. A pulse unit that was not a shaped charge would of course waste most of the energy of the explosion. Figure that 10% at best of the energy of a non-shaped-charge explosion would actually hit the pusher plate, what a waste of perfectly good plutonium.

    Each charge accelerates the spacecraft by roughly 12 m/s. A 4,000 ton spacecraft would use 5 kiloton charges, and a 10,000 ton spacecraft would use 15 kiloton charges. For blast-off, smaller charges of 0.15 kt and 0.35 kt respectively would be used while within the Terra's atmosphere. The air between the charge and the pusher plate amplifies the impulse delivered (it is extra propellant), so if you are not in airless space you can get away with a smaller kt yield.


    Even though only a fraction of the pulse unit's mass is officially tungsten propellant, you have to count the entire mass of the pulse unit when figuring the mass ratio. The mass of the Orion spacecraft with a full load of pulse units is the wet mass, and the mass with zero pulse units is the dry mass.

    The thrust is not applied constantly, it is in the form of pulses separated by a fixed detonation interval. Generally the interval is from about half a second to 1.5 seconds. This means to figure the "effective" thrust you take the thrust-per-pulse-unit and divide it by the detonation interval in seconds. So if each pulse unit gives 2×106 Newtons, and they are detonated at 0.8 second intervals, the effective thrust is 2×106 / 0.8 = 2.5×106 Newtons

    Obviously the converse is if you have the effective thrust, you multiply it by the detonation interval to find the thrust-per-pulse-unit. So if the effective thrust is 3.5×106 N and the units are detonated at 0.86 second intervals, the thrust-per-pulse-unit is 3.5×106 N * 0.86 = 3.01×106 Newtons


    How much weapons-grade plutonium will each charge require? As with most details about nuclear explosives, specifics are hard to come by. According to GA-5009 vol III , pulse units with 2.0×106 newtons to 4.0×107 newtons all require approximately 2 kilograms per pulse unit, with 1964 technology. It goes on to say that advances in the state of the art could reduce the required amount of plutonium by a factor 2 to 4, especially for lower thrust units. 2.0×106 n is 1 kiloton, I'm not sure what 4.0×107 n corresponds to, from the document I'd estimate it was about 15 kt. Presumably the 2 kg plutonium lower limit is due to problems with making a critical mass, you need a minimum amount to make it explode at all.


    According to Scott Lowther, the smallest pulse units were meant to propel a small ten-meter diameter Orion craft for the USAF and NASA. The units had a yield ranging from one-half to one kiloton. The USAF device was one kiloton, diameter 36 centimeters, mass of 86 kilograms, tungsten propellant mass of 34.3 kilograms, jet of tungsten plasma travels at 150,000 meters per second. One unit would deliver to the pusher plate a total impulse of 2,100,000 newton-seconds. Given the mass of the ten-meter Orion, detonating one pulse unit per second would give an acceleration well over one gee. According to my slide rule, this implies that the mass of the ten-meter Orion is a bit under 210 metric tons.

    Pulse UnitYieldMassDia.HeightPropellant
    (percent)
    Det.
    Interval
    Propellant
    Velocity
    Effective
    Exhaust
    Velocity
    (Isp)
    Thrust
    per unit
    Effective
    Thrust
    NASA 10m Orion
    (vacuum)
    141 kg0.86 s18,200 m/s
    (1,850 s)
    3.0×106 N3.5×106 N
    USAF 10m Orion
    (vacuum)
    1 kg79 kg
    (86 kg?)
    0.33 m0.61 m34.3 kg
    (40%)
    1 s1.5×105 m/s25,800 m/s
    (2,630 s)
    2.0×106 N2.0×106 N
    20m Orion
    (vacuum)
    450 kg0.87 s30,900 m/s
    (3,150 s)
    1.4×107 N1.6×107 N
    4000T Orion
    (atmo)
    0.15 kt1,152 kg0.81 m0.86 m1.1 s1.17×105 m/s42,120 m/s
    (4,300 s)
    8.8×107 N8.0×107 N
    4000T Orion
    (vacuum)
    5 kt1,152 kg0.81 m0.86 m415 kg
    (36%)
    1.1 s1.17×105 m/s42,120 m/s
    (4,300 s)
    8.8×107 N8.0×107 N
    10,000T Orion
    (atmo)
    0.35 kt118,000 m/s
    (12,000 s)
    4.0×108
    10,000T Orion
    (vacuum)
    15 kt118,000 m/s
    (12,000 s)
    4.0×108
    20,000T Orion
    (vacuum)
    29 kt1,150 kg0.8 m
    • Pulse Unit: The type of Orion spacecraft that uses this unit, and whether it is an atmospheric or vacuum type.
    • Yield: Nuclear explosive yield (kilotons)
    • Mass: Mass of the pulse unit
    • Dia.: Diameter of pulse unit
    • Height: Height of pulse unit
    • Propellant (percent): Mass of tungsten propellant in kilograms, as percentage of pulse unit mass in parenthesis.
    • Det. Interval: Time delay interval between pulse unit detonations.
    • Propellant Velocity: The velocity the tungsten propellant plasma travels at. Do not use this for delta V calculations.
    • Effective Exhaust Velocity (Isp): A value for exhaust velocity suitable for delta V calculations. Specific impulse in parenthesis.
    • Thrust per unit: Amount of thrust produced by detonating one pulse unit.
    • Effective Thrust: Thrust per second. Calculated by taking Thrust per unit and dividing by Det. Interval.

    There are some interesting equations in GA-5009 vol III on pages 25 and 26 on the subject of nuclear pulse units. These were developed in the study for the 10 and 20 meter NASA Orion spacecraft, and they heavily rely upon a number of simplifying assumptions. These were for first generation pulse units, with the assumption that second generation units would have better performance. So take these with a grain of salt.

    These equations are only considered valid over the range 3×106 < FE < 2×108

    You are given the amount of thrust you want to get out of the propulsion system: FE and the detonation interval time Dp. From those you calculate the amount of thrust each pulse unit has to deliver Fp:

    Fp = FE / Dp

    From this the specific impulse, nuclear yield, and the mass of the Orion propulsion module.

    Isp = 1 / ((5.30×102 / (Fp * (1 + (2.83×10-3 * Fp1/3)))) + ((4.32×10-2 * (1 + (2.83×10-3 * Fp1/3))) / Fp1/3))

    Ve = Isp * g0

    Y = 9.30×10-10 * Fp4/3

    ME = Fp / (3.6 * g0)

    where

    • FE = effective thrust (newtons)
    • Dp = delay between pulses (seconds)
    • Fp = thrust per pulse (newtons)
    • Isp = effective specific impulse (seconds)
    • Ve = exhaust velocity (m/s)
    • Y = size of nuclear yield in pulse unit (kilotons)
    • ME = mass of Orion propulsion module (kg)
    • g0 = acceleration due to gravity = 9.81 m/s2
    • x1/3 = cube root of x

    The results are close but do not exactly match the values given in the document, but they are better than nothing

    NASA 10-meter Orion
    Given Effective Thrust3.5×106 N
    Given Detonation Delay0.86 s
    ParameterDocument
    Value
    Equation
    Value
    Specific Impulse1,850 s1,830 s
    Yield1 kt0.4 kt
    Propulsion module mass90,946 kg85,245 kg
    NASA 20-meter Orion
    Given Effective Thrust1.6×107 N
    Given Detonation Delay0.87 s
    ParameterDocument
    Value
    Calculated
    Value
    Specific Impulse3,150 s3,082 s
    Yield5 kt3.1 kt
    Propulsion module mass358,000 kg394,223 kg

    For more in depth calculations of an Orion rocket's specific impulse, read page 1 and page 2. But be prepared for some heavy math.


    Tungsten has an atomic number (Z) of 74. When the tungsten plate is vaporized, the resulting plasma jet has a relatively low velocity and diverges at a wide angle (22.5 degrees). Now, if you replace the tungsten with a material with a low Z, the plasma jet will instead have a high velocity at a narrow angle. The jet angle also grows narrower as the thickness of the plate is reduced. This makes it a poor propulsion system, but an effective weapon. Instead of a wall of gas hitting the pusher plate, it is more like a directed energy weapon. The military found this to be fascinating, who needs cannons when you can shoot spears of pure nuclear flame? The process was examined in a Strategic Defense Initiative project called "Casaba-Howitzer", which apparently is still classified. Which is not surprising but frustrating if one is trying to write a science fiction novel or spacecraft combat game.

    NASA has been quietly re-examining ORION, under the new name of "External Pulsed Plasma Propulsion". As George Dyson observed, the new name removes most references to "Nuclear", and all references to "Bombs."

    For details about spacecraft using Orion propulsion, go here.


    Oh, and another thing. ORION is fantastic for boosting unreasonably huge payloads into orbit and it is pretty great for orbit to orbit propulsion. But trying to use it to land is not a very good idea. At least not on a planet with an atmosphere.

    Project Orion
    Project Orion
    Exhaust Velocity19,620 m/s
    Specific Impulse2,000 s
    Thrust2,215,200 N
    Thrust Power21.7 GW
    Mass Flow113 kg/s
    Total Engine Mass203,680 kg
    T/W1
    Frozen Flow eff.39%
    Thermal eff.99%
    Total eff.39%
    FuelFission:
    Curium 245
    ReactorPulse Unit
    RemassWater
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorPusher Plate
    Specific Power9 kg/MW

    This fabled technology converts the impulses of small nuclear detonations into thrust.

    The small shaped-charge bombs each have a mass of 230 kg (including propellant) and a yield of a quarter kiloton (1 terajoule). The fissile material is curium 245, with a critical mass of 4 kg, surrounded by a beryllium reflector. The soft X-rays, UV and plasma from the external detonation vaporize and compress the propellant to a gram per liter, highly opaque to the bomb energies at the temperatures attained (67000 K).

    The propellant, a mixture of water, nitrogen, and hydrogen, interfaces with a pusher plate “nozzle”, which can be either solid or magnetic.

    Shown is a solid plate, which tapers to the edges (to maintain a constant net velocity of the plate given a greater momentum transfer in the center). Pressure on the plate reaches 690 MPa in the center. The impulse shock is absorbed by a set of pneumatic “tires”, followed by gas-filled pistons detuned to the 56 Hz detonation frequency.

    The shock plate system becomes a useful shield if pointed towards the enemy.

    The amount of blast energy utilized for thrust is 7%, and the amount of pulse mass that intercepts the plate is 39%. A 56 TWth design optimized for 1TJ bombs achieves a specific impulse of 2 ksec and a thrust of 2.2 MN.

    Ted Taylor’s classic design, optimized for low yield bombs and 2 ksec specific impulse: “Project Orion”, George Dyson, Henry Holt and Company, 2002.

    From High Frontier by Philip Eklund
    Do Not Land With Orion

    Interesting e-mail conversation I had with Rhys Taylor on the topic of Entry-Descent-Landing (EDL) as relevant to nuclear pulse propulsion.

    I was aware one of the concepts that came out of the 1958 Project Orion involved landing a surface installation and a 100 man crew on the surface of Mars. Two of the early large Orion's would be involved. One would enter a low Mars orbit and completely cancel its orbital velocity while well above the sensible Martian atmosphere. The crew would ride down in a number of smaller landing craft with individual return stages. A large section of the vehicle, the base structure carrying a cargo of surface rovers, scientific gear, and consumables, would separate from the Orion propulsion module and descend propulsively on rockets without undergoing meteoric entry. The propulsion module would be allowed to crash on the surface (presumably this would entail transferring any remaining pulse units to the second Orion remaining in orbit before cancelling its orbital velocity — so only the absolute minimum required number of pulse units would remain to be expended before its uncontrolled descent and crash landing).

    My interest was in regards to soft landing an Orion intact after a controlled descent, and I was unsure of how deep into the atmosphere the nuclear pulse propulsion system could be fired, if it could be fired in descent mode, or if this was even advisable.

    Rhys was kind enough to advise me on these particular points, which to sum up are:

    1. Orion is capable of completely cancelling its orbital velocity.
    2. Descent would be a matter of managing the free-fall velocity of the vehicle.
    3. Inside the atmosphere the pulse unit will generate a many-thousands degree fireball, this is not a problem during launch, or in the vacuum of space, but during descent flying into the fireball would not be a good thing for vehicle and crew.
    4. There is some point at very high altitude where you would have to trade off from nuclear pulse propulsion to rocket powered descent.

    The input Rhys provided went toward this spacecraft designed for my Orion's Arm future history, and will be applied to several related spacecraft to be posted in the near future. 

    Orion Environmental Impact

    Naturally, some people freak out when you tell them about a rocket that rises into orbit by detonating Two! Hundred! Atom! Bombs!. But it actually isn't quite as bad as it sounds.

    First off, these are teeny-tiny atom bombs, honest. The nuclear pulse units used in space will be about one kiloton each, while the Nagasaki device was more like 20 kt. And in any event, the nuclear pulse units used in the atmosphere are only 0.15 kt ( about 1/130th the size of the Nagasaki device). This is because the atmosphere converts the explosion x-rays into "blast", increasing the effectiveness of the pulse unit so you can lower the kilotonnage.

    So we are not talking about zillions of 25 megaton city-killer nukes scorching the planet and causing nuclear winter.


    Some environmentalists howl that Orion should never be used for surface-to-orbit boosts, due to the danger of DUNT-dunt-Dunnnnnnnn Deadly Radioactive Fallout. However, there is a recent report that suggests ways of minimizing the fallout from an ORION doing a ground lift-off (or a, wait for it, "blast-off" {rimshot}). Apparently if the launch pad is a large piece of armor plate with a coating of graphite there is little or no fallout.

    By which they mean, little or no ground dirt irradiated by neutrons and transformed into deadly fallout and spread the the four winds.

    There is another problem, though, ironically because the pulse units use small low-yield nuclear devices.

    Large devices can be made very efficient, pretty much 100% of the uranium or plutonium is consumed in the nuclear reaction. It is much more difficult with low-yield devices, especially sub-kiloton devices. Some of the plutonium is not consumed, it is merely vaporized and sprayed into the atmosphere. Fallout, in other words. You will need to develop low-yield devices with 100% plutonium burn-up, or use fusion devices (with 100% burn-up fission triggers or with laser inertial confinement fusion triggers).

    The alternative is boosting the Orion about 90 kilometers up using a non-fallout chemical rocket. Which more or less defeats the purpose of using an Orion engine in the first place. Remember that Orions are best at boosting massive payloads into orbit.

    Most of the fallout will fall within 80 kilometers of the launch site. You can also reduce the fallout by a factor of 10 if you launch from near the Magnetic Pole.


    When fissionables like plutonium undergo fission, their atoms are split which produces atomic energy. The split atoms are called fission fragments.

    The good news is that they have very short half-lives, e.g., in 50 days pretty much all of the Strontium 94 has decayed away (because 50 days is 58,000 St94 half-lives).

    The bad news is that they have very short half-lives, this means they are hideously radioactive. Radioactive elements decay by emitting radiation, shorter half-life means more decays per second means a higher dose of radiation per second.

    The fragments that come screaming out of the detonation aimed at the sky are no problem. They are moving several times faster than Terra's escape velocity, you will never see them again (Terra's escape velocity is 11.2 km/s, the fragments are travelling like a bat out of hell at 2,000 km/s). The ones aimed towards Terra are a problem. The fragments can be reduced by using fusion instead of fission pulse units. The fragments can also be reduced by designing the pulse units to trade thrust in favor of directing more of the fragments skyward.


    A more sophisticated objection to using Orion inside an atmosphere is the sci-fi horror of EMP melting all our computers, making our smart phones explode, and otherwise ruining anything using electricity. But that actually is not much of a problem. EMP is not a concern unless the detonation is larger than one megaton or so, Orion propulsion charges are only a few kilotons (one one-thousandth of a megaton). Ben Pearson did an analysis and concluded that Orion charges would only have EMP effects within a radius of 276 kilometers (the International Space Station has an orbital height of about 370 kilometers). So just be sure your launch site is in a remote location, which you probably would have done anyway.


    Naturally watching an Orion blast-off is very bad for your eyes, defined as instant permanent blindness. This is called "eyeburn". While the Orion is below 30 km you definitely need protective goggles or you might be blinded. Above 90 km your eyesight it safe. In between 30 and 90 is the gray area, where prudent people keep their protective goggles on.


    Detonating pulse units in space near Terra can create nasty artificial radiation belts. The explosion can pump electrons into the magnetosphere, creating the belt.

    There are two factors: detonation altitude from Terra's surface, and magnetic latitude in Terra's magnetic field. If the detonation is within 6,700 kilometers of Terra's surface (i.e., further than 2 Terran radii from Terra's center) and at a magnetic latitude from 0° to 40°, the radiation belt can last for years. Above 2 Terran radii the radiation belt will last for only weeks, and from latitude 80° to 90°, the radiation belt will last for only a few minutes.

    The military discovered this the hard way with the Starfish Prime nuclear test. The instant auroras were very pretty. The instant EMP was very scary, larger than expected (but the test was using a 1.4 megaton nuke, not a 0.001 megaton pulse unit). The artificial radiation belt that showed up a few days later was a very rude surprise. About one-third of all low orbiting satellites were eventually destroyed by the radiation belt.

    The radiation belts are harmless to people on Terra, but astronauts in orbit and satellites are at risk.


    There are three classes of pulse unit failure modes. Note that in this analysis the USAF had given up and had decided to boost the Orion on top of a chemical rocket.

    Class I - Pad Abort
    Typically occurs when the chemical booster burns or explodes on the pad. There will be no nuclear explosion. The pulse units contain chemical explosives, but there is much more explosive potential in the chemical booster fuel. Even if all the pulse units exploded simultaneously there would only be a 1 psi overpressure out to 300 meters and shrapnel hazard out to 2,000 meters.

    A chemical booster burn could aerosolize radioactive plutonium from booster units and create a downrange fallout hazard. The solution is to put the launch pad over a pool of water about 10 meters deep. In event of fire, collapse the pad into the pool. The fire would be extinguished and any escaped plutonium will be contained in the water. Many of the pulse units can be recovered and reused.
    Class II - Failure to Orbit
    The trouble is that the thousands of nuclear pulse units will fall down, probably into uncontrolled territory. As with Class I there will be no nuclear explosion, the chemical explosion will be impressive but not too huge, and there is a danger of radioactive fallout. All in what could very well be a foreign country.

    In addition, it will be scattering thousands of containers of weapons grade plutonium in convenient form to cause nuclear weapon proliferation. Or the pulse units could be used as is as impromptu terrorist devices. Though I'm sure the devices will contain fail-safes seven ways to Sunday, the same way nuclear warheads are in order to deal with the possibility of them falling into the Wrong Hands.

    Probably the best solution is to command all of the nuclear charges to detonate simultaneously while the spacecraft is at high altitude. This will make one heck of a fireworks display, and may cause an EMP, but nuclear devices in questionable hands is to be avoided at all costs.
    Class III - Misfire
    If a given pulse unit fails to detonate, the command can be resent repeatably, and/or there can be an automatic on-board destruct system. Otherwise the unit could survive reentry (due to the tungsten propellant plate) causing some damage to the country it hit and causing a foreign policy nightmare to the nation owning the Orion spacecraft.

    By about 1963 General Atomic had given up on designing an Orion to lift off from Terra's surface under nuclear power. They put together three plans for using chemical rocket boosters to get the Orion into orbit. Again this is throwing away the big advantage of the Orion, its ability to boost massive payloads.

    Mode I
    A fully loaded and fully fueled Orion is boosted to an altitude of 90 kilometers and 900 m/s by a chemical rocket. There it stages, and the Orion proceeds into orbit or into mission trajectory under nuclear power. The disadvantage is it requires a subobital start-up of the Orion engine. The Orion engine will need a thrust greater than the mass of the spacecraft, the standard was T/W of 1.25. But high thrust is never a problem with Orion.
    Mode II
    An empty Orion is loaded with just enough pulse units. It is boosted to an altitude of 90 kilometers and 900 m/s by a chemical rocket. There it stages, and the Orion proceeds into orbit. A second chemical booster rendezvous with the Orion to deliver the payload and a full load of pulse units.
    This was the worst plan. It combines the disadvantage of Mode I (by requiring suborbital start-up of the Orion engine) with the disadvantage of Mode III (by requiring orbital assembly).
    Mode III
    The Orion is boosted into orbit piecemeal as payload on a series of chemical boosters. The Orion is assembled in orbit, then departs on its mission under nuclear power. The main advantage is it avoids the possibility of the entire Orion spacecraft crashing to Terra in the event of a propulsion failure. The second advantage is it allowed a lower thrust Orion unit to be used, but with Orion thrust is never a problem. The main disadvantage is that orbital assembly is time consuming and difficult.

    Zeta-Pinch

    Zeta pinch is a type of plasma confinement system that uses an electrical current in the plasma to generate a magnetic field that compresses it. The compression is due to the Lorentz force.

    Zeta-Pinch Fission

    Mini-Mag Orion
    Mini-Mag Orion
    Exhaust Velocity157,000 m/s
    Specific Impulse16,004 s
    Thrust1,870,000 N
    Thrust Power0.1 TW
    Mass Flow12 kg/s
    Total Engine Mass199,600 kg
    T/W0.95
    FuelFission:
    Curium 245
    Specific Power1 kg/MW
    FuelFission:
    Curium 245
    ReactorZeta-Pinch
    RemassReaction
    Products
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle
    Mini-Mag Orion (DRM-1)
    Exhaust Velocity93,164 m/s
    Specific Impulse9,497 s
    Thrust642,000 N
    Thrust Power29.9 GW
    Mass Flow7 kg/s
    Total Engine Mass119,046 kg
    T/W0.55
    Wet Mass731,924 kg
    Dry Mass250,300 kg
    Mass Ratio2.92 m/s
    ΔV99,967 m/s
    Specific Power4 kg/MW
    Mini-Mag Orion (DRM-3)
    Exhaust Velocity93,000 m/s
    Specific Impulse9,480 s
    Thrust642,000 N
    Thrust Power29.9 GW
    Mass Flow7 kg/s
    Total Engine Mass199,600 kg
    T/W0.33
    Wet Mass788,686 kg
    Dry Mass157,723 kg
    Mass Ratio5.00 m/s
    ΔV149,686 m/s
    Specific Power7 kg/MW

    The Mini-MagOrion is a sort of micro-fission Orion propulsion system. The idea was to make an Orion with weaker (and more reasonably sized) explosive pulses, using pulse charges that were not self contained (the full Orion pulse units were nothing less than nuclear bombs). Subcritical hollow spheres of curium-245 are compressed by a Z-pinch magnetic field until they explode. The sacrificial Z-pinch coil in each pulse charge is energized by an a huge external capacitor bank mounted in the spacecraft. So the pulse units are not bombs.

    The explosion is caught by a superconducting magnetic nozzle.

    More details are in the Realistic Designs section.

    Z-pinch Microfission
    Z-pinch Microfission
    Z-pinch Microfission
    Exhaust Velocity156,960 m/s
    Specific Impulse16,000 s
    Thrust8,500 N
    Thrust Power0.7 GW
    Mass Flow0.05 kg/s
    Total Engine Mass193,333 kg
    T/W4.00e-03
    Frozen Flow eff.74%
    Thermal eff.90%
    Total eff.67%
    FuelFission:
    Curium 245
    ReactorZeta-Pinch
    RemassReaction
    Products
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle
    Specific Power290 kg/MW

    Electrodynamic zeta-pinch compression can be used to generate critical mass atomic bombs at very low yields. These detonations can be used to generate impulsive power or thrust.

    Exotic fission material (245Cm) is utilized to reduce the required compression ratio. The explosion of each low yield (335 GJ) atomic bomb energizes and vaporizes a set of low mass transmission lines, used to pump either another high current Z-pinch, or a bank of nanotube-enhanced ultracapacitors.

    Each bomb uses 40 grams of Cm fissile material and 60 grams of Be reflector material, with an aspect ratio of 5. A DT diode is used as a neutron emitter. The mylar transmission lines have a mass of 15 kg, and are replaced after each shot.

    The design illustrated is rated for a shot every 5.5 minutes, equivalent an output of 1000 MWth. If utilized for thrust, this provides 7.7 kN at a specific impulse of 17 ksec.

    Ralph Ewig & Dana Andrews, “Mini-MagOrion Micro Fission Powered Orion Rocket”, Andrews Space & Technology, 2002.

    From High Frontier by Philip Eklund
    n-Li6 Microfission
    n-Li6 Microfission
    n-6Li Microfission
    Exhaust Velocity156,960 m/s
    Specific Impulse16,000 s
    Thrust20,000 N
    Thrust Power1.6 GW
    Mass Flow0.13 kg/s
    Total Engine Mass106,667 kg
    T/W0.02
    Frozen Flow eff.87%
    Thermal eff.90%
    Total eff.78%
    FuelFission:
    Lithium 6
    ReactorUltracold Neutron
    Catalyzed
    RemassReaction
    Products
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle
    Specific Power68 kg/MW

    The minimum explosive yield for fission bombs is about a quarter kiloton. Thus, rockets that fly using atomic explosions, such as Project Orion, require huge shock absorbers.

    The pulse energy can be brought down to microfission levels by the use of exotic particles. A n-6Li microfission thruster brings the lithium isotope 6Li to spontaneous microfission by interaction with particles with very large reaction cross sections such as ultracold neutrons. No “critical mass” is required. This clean reaction produces only charged particles (T and He), each at about 2 MeV.

    The system illustrated uses a 5-meter magnetic nozzle to transfer the microexplosion energy to the vehicle. This magnetic impulse transfer is borrowed from the MagOrion concept (combination of Orion and the magnetic sail).

    A fuel reaction rate of 60 mg/sec yields 3720 MWth. At a pulse repetition rate of one 224 GJ (0.05 kT) detonation each minute, the thrust is 12.8 kN at a 12 ksec specific impulse. A hydraulic fixture oscillates at a tuned frequency to provide a constant acceleration to the spacecraft. The combined frozen-flow and nozzle efficiencies are 21%, and the thermal efficiency is 96%.

    Ralph Ewig’s “Mini-magOrion” concept, modified for n-6Li fission, http://www.andrews-space.com/images/videos/PAPERS/Pub-MMOJPLTalk.pdf

    From High Frontier by Philip Eklund
    Ultracold Neutrons

    Neutrons are normally unstable particles, with a half life of 12 minutes.

    When polarized and ultra-cooled (using vibrators or turbines), they form a dineutron or tetraneutron phase. These “molecules” are believed to be stable and storable in total internal reflection bottles, lined with diamond-like carbon as the neutron reflector.

    Ultracold neutrons (UCN) have a huge quantum mechanical wavelength as a consequence of their slow movement (typically 0.4 μm @ 1 m/sec), and thus can spontaneously initiate fission reactions such as n-235U or n-6Li.

    If the neutron source is a nuclear reactor, the neutrons must be cooled from 2 MeV to 2 meV using a heavy water moderator, and then in a UCN turbine to 0.2 IeV.

    Robert L. Forward, “Alternate Propulsion Energy Sources”, 1983.

    From High Frontier by Philip Eklund

    Zeta-Pinch Fusion

    HOPE Z-Pinch
    Propulsion SystemZ-Pinch Fusion
    Exhaust Velocity189,780 m/s
    Specific Impulse19,346 s
    Thrust38,120 N
    Thrust Power3.6 GW
    Mass Flow0.20 kg/s
    Total Engine Mass95,138 kg
    T/W0.04
    FuelDeuterium-Tritium fusion
    + Lithium6 fission
    ReactorZeta-Pinch
    RemassReaction
    Products
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle
    Wet Mass888,720 kg
    Dry Mass552,000 kg
    Mass Ratio1.61 m/s
    ΔV90,380 m/s
    Specific Power26 kg/MW
    Firefly Starship
    2013 design
    ΔV2.698×107 m/s
    (0.09c)
    Wet Mass17,800 metric tons
    Dry Mass2,365 metric tons
    Mass Ratio7.526
    Payload150 metric tons
    PropulsionZ-Pinch DD Fusion
    Exhaust Velocity1.289×107 m/s
    Thrust1.9×106 N
    Acceleration0.11 m/s
    (0.01 g)
    Accel time4 years
    Coast time93 years
    Decel time1 years

    Medusa

    Medusa
    Exhaust velocity490,000 m/s
    to 980,000 m/s

    Medusa is driven by the detonation of nuclear charges like Orion, except the charges are set off in front of the spacecraft instead of behind. The spacecraft trails behind a monstrously huge parachute shaped sail (about 500 meters). The sail intercepts the energy from the explosion. Medusa performs better than the classical Orion design because its pusher plate intercepts more of the bomb's blast, its shock-absorber stroke is much longer, and all its major structures are in tension and hence can be quite lightweight. It also scales down better. The nuclear charges will be from 0.025 kilotons to 2.5 kilotons.

    The complicated stroke cycle is to smooth out the impulses from each blast, transforming it from a neck-braking jerk into a prolonged smooth acceleration.

    Jondale Solem calculates that the specific impulse is a function of the mass and yield of the nuclear charges, while the thrust is a function of the yield and explosion repetition rate. In this case, the mass of the nuclear charge is the mass of "propellant".

    Remarkably the mass of the spinnaker (sail) is independent of the size of its canopy or the number or length of its tethers. This means the canopy can be made very large (so the bomb blast radiation does not harm the canopy) and the tethers can be made very long (so the bomb blast radiation does not harm the crew). The mass of the spinnaker is directly proportional to the bomb yield and inversely proportional to the number of tethers.

    Inertial Confinement

    IC-Fusion
    Exhaust Velocity10,000,000 m/s
    Specific Impulse1,019,368 s
    Thrust100,000,000 N
    Thrust Power500.0 TW
    Mass Flow10 kg/s
    Total Engine Mass1,000,000 kg
    T/W10
    FuelProton-Proton
    Fusion
    Specific Power2.00e-03 kg/MW

    A pellet of fusion fuel is bombarded on all sides by strong pulses from laser or particle accelerators. The inertia of the fuel holds it together long enough for most of it to undergo fusion.

    D-D Fusion Inertial
    D-D Fusion Inertial
    Exhaust Velocity78,480 m/s
    Specific Impulse8,000 s
    Thrust3,200 N
    Thrust Power0.1 GW
    Mass Flow0.04 kg/s
    Total Engine Mass243,333 kg
    T/W1.00e-03
    Frozen Flow eff.50%
    Thermal eff.50%
    Total eff.25%
    FuelDeuterium-Deuterium
    Fusion
    ReactorInertial Confinement
    Laser
    RemassGraphite
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorAblative Nozzle
    Specific Power1,938 kg/MW

    A “target” of fusion fuel can be brought to ignition by “inertial confinement”: the process of compressing and heating the fuel with beamed energy arriving from all sides. A snowflake of deuterium, the “heavy” isotope of hydrogen, can be imploded and fused with a combination of lasers and deuterium particle beams.

    The illustrated design uses combined input beam energy of 38 megajoules, arrayed in a ring surrounding the ejected iceball target. This energy operates at 1 Hz to blast a 2 gram ice pellet ejected each second. The outside 99% of the pellet is ablated away within 10 ns, super-compressing the deuterium fuel at the core to a density of a kilogram per cubic centimeter. The T and 3He products are catalyzed to undergo further fusion until all that remains is hydrogen, helium and some neutrons. (Neutrons comprise 36% of the reaction energy.) Fractional burn-up of the fuel (30%) is twice that of magnetic confinement systems, which implies a 40% higher fuel economy. The energy gain factor (Q) is 53.

    For a 500 MWth reactor, 320 MW of charged particles are produced, which can be used directly for thrust or metals refining. About 105 MW of fast neutrons escape to space, but another 75 MW of them are intercepted by the structure. About two thirds of this energy must be rejected as waste heat, but the remainder is thermally used to generate electricity or to breed tritium to be added to the fuel to facilitate the cat D-D pellet ignition.

    When used as a rocket, an ablative nozzle, made of nested layers of whisker graphite whose mass counts as propellant and shadow shield, is employed (much like the ACMF).

    “A Laser Fusion Rocket for Interplanetary Propulsion,” Hyde, R., 34th International Astronautical Conf., AIF Paper 83-396, Budapest, Hungary, Oct. 1983.

    (To keep radiator mass under control, I reduced the pellet repetition rate from 100 Hz to 1 Hz).

    From High Frontier by Philip Eklund
    VISTA
    Propulsion SystemIC Fusion
    Exhaust Velocity170,000 m/s
    Specific Impulse17,329 s
    Thrust240,000 N
    Thrust Power20.4 GW
    Mass Flow1 kg/s
    FuelDeuterium-Tritium
    Fusion
    ReactorInertial Confinement
    Laser
    RemassReaction
    Products
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle
    Wet Mass6,000,000 kg
    Dry Mass1,835,000 kg
    Mass Ratio3.27 m/s
    Width170 m
    Height100 m

    Magneto Inertial Fusion

    Magneto Inertial Fusion
    Both
    Exhaust Velocity50,420 m/s
    Specific Impulse5,140 s
    FuelDeuterium-Deuterium
    Fusion
    ReactorMagneto-Inertial
    Confinement
    RemassLithium
    Remass AccelThermal Accel:
    Reaction Heat
    Low Gear
    Thrust103 N
    Thrust Power2.6 MW
    Mass Flow2.00e-03 kg/s
    Delay between
    Fusion Pulses
    180 seconds
    High Gear
    Thrust13,800 N
    Thrust Power0.3 GW
    Mass Flow0.27 kg/s
    Delay between
    Fusion Pulses
    14 seconds

    There are two main approaches to utilizing nuclear fusion, magnetic confinement and inertial confinement. Magnetic confinement uses titanic magnetic fields, inertial confinement is how fusion bombs explode (a third way would be stars shining by gravitational confinement, but we don't know how to generate artificial gravitational fields). As propulsion systems, both have major drawbacks.

    Magnetic confinement requires huge (read: massive) electromagnets. The technique also has the problem of plasma instabilities (read: fusion plasma has thousands of different ways to wiggle out of the magnetic cage) which so far have defied any solution.

    Inertial confinement works well in bombs, but trying to do it in a small controlled fashion (read: so the fusion reaction does not vaporize everything in a one kilometer radius) has also defied any solution. The compressing laser or particle beams have such low efficiencies that tons of excess power is required. Timing all the beams so they strike at the same instant is a challenge. Also, there is nothing in between the fusion reaction and the chamber walls, leading to severe damage to the walls.

    Both approaches have a problem with getting the fusion reaction energy to heat the propellant. Magnetic confinement tries to use the actual fusion plasma as propellant, resulting in a ridiculously small mass flow and thus a tiny thrust.

    Dr. John Slough and his associates have come up with a new technique that sort of combines the two conventional approaches: magneto inertial fusion (MIF). You can find their published papers on the subject here

    A blob of FRC (field reversed configuration) plasma is created and injected axially into the chamber.

    Simultaneously injected into the chamber is a "liner". The liner is a foil ring composed of lithium, about 0.2 meters in radius. Each liner will have a mass of 0.28 kg (minimum) to 0.41 kg.

    As the liner travels axially down the chamber, electromagnets crush it down into a solid cylinder (the crush speed is about 3 kilometers per second, the cylinder will have a radius of 5 centimeters). This is timed so that the plasma blob (plasmoid) is in the center of the cylinder. The liner compresses the plasmoid and ignites the fusion reaction.

    The lithium stands in between the reaction and the chamber walls, protecting the walls. It also absorbs much of the radiation, protecting the crew. The lithium is also the propellant. Since it is tightly wrapped around the reaction, it is very efficient at getting the fusion reaction energy to heat the propellant. The ionized lithium (plus the burnt fusion fuel) exits through a magnetic nozzle, providing thrust.

    Since this is an open-cycle system, the exhaust acts as the heat radiator, so the spacecraft can get by with only a tiny radiator. The energy to run the magnets can be supplied by solar cell arrays. Since the compression is so efficient, this will work with several types of fusion fuel: D-T, D-D, and D-3He. D-D is probably preferred, since tritium is radioactive with a short half-life, and 3He is rare.

    Please note that if you replace the magnetic nozzle with a magnetohydrodynamic (MHD) generator, the propulsion system is transformed into an electrical power generator. This could be used for ground based fusion power generators.

    Dr. Slough et al worked up two spacecraft for a Mars mission. The first was optimized to have a high payload mass fraction. The second was optimized to have the fastest transit time. Both were capable of a direct abort and return. The "Low Gear" engine is the study author's opinion of an engine easily achievable with current technology (that is, achievable fusion yields). The "High Gear" engine is a bit more speculative, but requiring only modest incremental improvements in technology.

    Fusion Drive Rockets (FDR)
    High Mass Fraction
    EngineLow Gear
    Transit Time90 days
    Initial Mass90 mT
    Payload Mass Fraction65%
    Specific Mass4.3 kg/kW
    Shortest Transit Time
    EngineHigh Gear
    Transit Time30 days
    Initial Mass153 mT
    Payload Mass Fraction36%
    Specific Mass0.38 kg/kW

    Antimatter Bottle

    Antimatter Bottle
    Antimatter Bottle
    Exhaust Velocity78,480 m/s
    Specific Impulse8,000 s
    Thrust34,700 N
    Thrust Power1.4 GW
    Mass Flow0.44 kg/s
    Total Engine Mass180,000 kg
    T/W0.02
    Frozen Flow eff.80%
    Thermal eff.85%
    Total eff.68%
    FuelAntimatter:
    antiprotons
    ReactorAntimatter Catalyzed
    RemassLead
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle
    Specific Power132 kg/MW

    Antimatter fuel can be stored as levitated antihydrogen ice. By illuminating it with UV to drive off the positrons, a bit is electromagnetically extracted and sent to a magnetic bottle.

    There it is collided with 60 g of heavy metal propellant (9 × 1024 atoms of lead or depleted uranium). Each antiproton annihilates a proton or neutron in the nucleus of a heavy atom. The use of heavy metals helps to suppress neutral pion and gamma ray production by reabsorption within the fissioning nucleus. If regolith is used instead of a heavy metal, the gamma flux is trebled requiring far more cooling.

    A pulse of 5 μg of fuel (3 × 1018 antiprotons) contains 900 MJ of energy, and at a repetition rate of 0.8 Hz, a power level of 700 MWth is attained.

    Compared to fusion, antimatter rockets need higher magnetic field strengths: 16 Tesla in the bottle and 50 Tesla in the throat. After 7 ms, this field is relaxed to allow the plasma to escape at 6 keV and 350 atm.

    These high temperatures and pressures cause higher bremsstrahlung X-ray losses than fusion reactors. Furthermore, the antiproton reaction products are short-lived charged pions and muons, that must be exhausted quickly to prevent an increasing amount of reaction power lost to neutrinos. About a third of the reaction energy is X-rays and neutrons stopped as heat in the shields (partly recoverable in a Brayton cycle), another third escapes as neutrinos. Only the final third is charged fragments directly converted to thrust or electricity in a MHD nozzle.

    D.L. Morgan, “Concepts for the Design of an Antimatter Annihilation Rocket,” J. British Interplanetary Soc. 35, 1982. (For use in this game, to keep the radiator mass within reasonable bounds, I reduced the pulse rate from 60 Hz to 0.8 Hz.)

    Robert L. Forward, “Antiproton Annihilation Propulsion”, University of Dayton, 1985.

    From High Frontier by Philip Eklund

    Antimatter catalyzed

    Nuclear fission pulse drives like Orion scale up well, since it is relatively easy to design a bigger bomb than the last one. However, physics seem to prevent the creation of a nuclear device with a yield smaller than about 1/100 kiloton (10 tons, 42 GJ) and a fissionable material mass under 25 kilograms. This is due to critical mass restraints.

    However, if a tiny sub-critical bit of fissionable material is bombarded by a few antiprotons, it will indeed create a tiny nuclear explosion. The antiprotons annihilate protons in uranium atoms, the energy release splits the atoms, creating a shower of neutrons, and a normal chain reaction ensues. Using antiprotons, yields smaller than 1/100 kiloton can be achieved. This can be used to create Antimatter catalyzed nuclear pulse propulsion

    AIM

    AIM
    Exhaust Velocity598,000 m/s
    Specific Impulse60,958 s
    Thrust55 N
    Thrust Power16.4 MW
    Mass Flow1.00e-04 kg/s
    FuelHelium3-Deuterium
    Fusion
    ReactorAntimatter Catalyzed
    RemassReaction
    Products
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorMagnetic Nozzle

    Antiproton-initiated Microfusion. Inertial Confinement Fusion. See here.

    ACMF

    ICAN-II
    Propulsion SystemACMF
    Exhaust Velocity132,435 m/s
    Specific Impulse13,500 s
    Thrust180,000 N
    Thrust Power11.9 GW
    Mass Flow1 kg/s
    Total Engine Mass27,000 kg
    T/W0.68
    FuelFission:
    Uranium 235
    ReactorAntimatter Catalyzed
    RemassSilicon Carbide
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorAblative Nozzle
    Wet Mass707,000 kg
    Dry Mass345,000 kg
    Mass Ratio2.05 m/s
    ΔV95,020 m/s
    Specific Power2 kg/MW

    Antiproton-catalyzed microfission, inertial confinement fission. See here.

    Fuel pellets have 3.0 grams of nuclear fuel (molar ratio of 9:1 of Deuterium:Uranium 235) coated with a spherical shell of 200 grams of lead. The lead shell is to convert the high energy radiation into a form more suited to be absorbed by the propellant. Each pellet produces 302 gigajoules of energy (about 72 tons of TNT) and are fired off at a rate of 1 Hz (one per second). The pellet explodes when it is struck by a beam containing about 1×1011 antiprotons.

    A sector of a spherical shell of 4 meters radius is centered on the pellet detonation point. The shell is the solid propellant, silicon carbide (SiC), ablative propellant. The missing part of the shell constitutes the exhaust nozzle. Each fuel pellet detonation vaporizes 0.8 kilograms of propellant from the interior of the shell, which shoots out the exhaust port at 132,000 meters per second. This produces a thrust of 106,000 newtons.

    The Penn State ICAN-II spacecraft was to have an ACMF engine, a delta-V capacity of 100,000 m/s, and a dry mass of 345 metric tons. The delta-V and exhaust velocity implied a mass ratio of 2.05. The dry mass and the mass ratio implied that the silicon carbide propellant shell has a mass of 362 metric tons. The wet mass and the thrust implied an acceleration of 0.15 m/s2 or about 0.015g. It can boost to a velocity of 25 km/sec in about three days. At 0.8 kilograms propellant ablated per fuel pellet, it would require about 453,000 pellets to ablat the entire propellant shell.

    It carries 65 nanograms of antiprotons in the storage ring. At about 7×1014 antiprotons per nanogram, and 1×1011 antiprotons needed to ignite one fuel pellet, that's enough to ignite about 453,000 fuel pellets.

    The system is very similar to Positron Ablative.

    H-B inertial catalzyed fusion
    H-B cat inertial
    Exhaust Velocity156,960 m/s
    Specific Impulse16,000 s
    Thrust4,700 N
    Thrust Power0.4 GW
    Mass Flow0.03 kg/s
    Total Engine Mass65,089 kg
    T/W7.00e-03
    Frozen Flow eff.86%
    Thermal eff.85%
    Total eff.73%
    FuelHydrogen-Boron
    Fusion
    ReactorAntimatter Catalyzed
    RemassGraphite
    Remass AccelThermal Accel:
    Reaction Heat
    Thrust DirectorAblative Nozzle
    Specific Power176 kg/MW

    The fusion of hydrogen and boron 11 is a clean reaction, releasing only 300 keV alpha particles, which can be magnetically directed. However, the H-B fusion will not proceed at temperatures less than 300 keV unless catalyzed using exotic particles.

    One possibility: replace the electrons in H-B atoms with stable massive leptons such as magnetic monopoles or fractionally-charged particles (the existence of these is hypothetical). The resulting exotic atoms can fuse at “cold” temperatures, allowing the exotic catalysts to be recycled.

    A second possibility is to use antiproton-catalyzed microfission to initiate the H-B fusion. If a hundred billion antiprotons at 1.2 MeV in a 2 nsec pulse are shot at a target of three grams of HB: 235U in a 9:1 molar ratio, the uranium microfission initiates H-B and releases 20 GJ of energy. Operating at a fifth of a hertz, hydrogen and boron 11 reacting at a rate of 145 mg/shot produces 2000 MWth. A shell of 200g of lead about the target thermalizes the plasma from 35 keV average to 1 keV, low enough that this radiation can be optimally transferred to thrust using a magnetic or ablative nozzle at 73% efficiency. The ejected mass per shot is 2.4 kg. The exotic catalysts are recycled. Catalyzed fusion enjoys an excellent thermal efficiency (86%) and thus a good thrust/weight ratio (3.2 milli-g), making it one of the best engines in the game. The specific impulse ranges between 8 and 16 ksec, depending whether spin-polarized free radicals are used as the hydrogen fuel.

    “Antiproton-Catalyzed Microfission/Fusion Propulsion Systems for Exploration of the Outer Solar System and Beyond”, G. Gaidos, et al., Pennsylvania State University, 1998.

    (I used the ICAN-II spacecraft design, modified from cat D-T to cat H-B fuel, and scaled way down from 1 Hz to 0.2 Hz, and 302 GW to 2 GW.)

    From High Frontier by Philip Eklund

    Sail

    Sail propulsion does not carry onboard reaction mass or does not use reaction mass. They are powered by a remote source, either the Sun or a satellite installation with a huge power supply and an equally huge laser/plasma beam.

    Electric Sail

    An E-Sail is a sail powered by solar wind.

    Electric Sail
    Electric Sail
    Thrust per tether0.01 N
    @ 1 AU
    Number tethers30,000
    Total Thrust300 N
    @ 1 AU
    Mass per tether1 kg
    Total Mass30,000 kg

    The solar wind dynamic pressure is about 2 nPa at one AU. An electric sail generates nanothrust from this particle stream in a manner similar to a mag sail, except that electric rather than magnetic fields are used.

    Its geometry employs hundreds of long thin conducting wires, rotating with a period of 20 minutes to keep them in positive tension.

    A solar-powered electron gun (typical power is a few hundred watts) keeps the spacecraft and sail in a high positive potential (up to 20 kV). This electric field surrounds each wire a few tens of meters into the surrounding solar wind plasma. Therefore the solar wind protons "see" the positively-charged wires as rather thick obstacles. It is this multiplication factor that allows sails using the solar wind to outperform those using photon pressure, which is 5000 times stronger.

    Furthermore, the electric sail thrust force varies as (1/r){7/6} from Sol, compared to the photon pressure, which varies as the inverse square distance.

    Each 100 km tether, massing but a kilogram, generates 0.01 N of thrust. Simultaneously it also attracts electrons from the solar wind plasma, which are neutralized by the electron gun. Potentiometers between each tether and the spacecraft control the attitude by fine-tuning the tether potentials. Additionally, the thrust may be turned off by simply switching off the electron gun.

    Each 20 μm tether is redundantly interlinked for robustness against meteoroids.

    Electric sails must avoid magnetospheres, since there is no solar wind inside these zones.

    Pekka Janhunen, “Electric Sail”, 2004. P. Janhunen and A. Sandroos,“Simulation study of solar wind push on a charged wire” 2007.

    From High Frontier by Philip Eklund

    Magnetic Sail

    Magnetic Sail
    Thrust per sail area0.001 N/km2
    Thrust by Sol dist1/R2

    A MagSail is a sail powered by the solar magnetic field.

    Electric Sail
    Thrust110 N
    @ 1 AU
    Sail Mass20,000 kg

    At 1 AU, the solar wind comprises several million protons per cubic meter, spiraling away from the sun at 400 to 600 km/sec (256 μwatts/m2). When such charged particles move through a magnetic field formed by the mag sail, a tremendous loop of wire some 2 km across, they are deflected.

    An unloaded mag sail this size has a thrust of 100 N (at 1 AU) and a mass of 20 tonnes. The wire is superconducting whisker, at 10 kg/km, connected to a central bus and payload via shroud lines. The loop requires multi-layer insulation and reflective coatings to maintain its superconducting temperature of 77 K. Because the sail area is a massless magnetic field, a mag sail has a superior thrust/weight ratio than photon sails.

    Just as with photon sails, lateral motion is possible by orienting the sail at an angle to the thrusting medium. A mag sail also develops thrust from planetary and solar magnetospheres, which decrease as the fourth power of the distance from the magnetosphere source. Field strength is typically 10 μT in Earth’s magnetosphere, or less in the solar magnetosphere.

    The mag sail illustrated is augmented by a spinning disk photon sail attached to its staying lines. It is maneuvered using photonic laser thrusters (propellantless thrust derived from the bouncing of laser photons between two mirrors).

    Zubrin 1988.

    From High Frontier by Philip Eklund

    M2P2

    M2P2
    Thrust per sail area0.001 N/km2
    Thrust by Sol distConstant
    Disk Inflates
    as 1/R2
    Plasma use0.25 kg/Day per N Thrust
    Isp = 35,000

    A Mini-magnetospheric plasma sail (M2P2) is a MagSail inflated by an injection of plasma, powered by the solar wind.

    MagBeam

    A MagBeam is Mini-magnetospheric plasma sail beam-powered by a remote helicon plasma beam installation. Report here. Alternatively the spacecraft can use a plasma magnet instead of a M2P2 to intercept the beam. With the current design, the spacecraft mass cannot be larger than about 10,000 kg (10 metric tons).

    The installation is called a High Power Platform (HPP). The HPP does not have much range, so the spacecraft will require a second HPP at the destination in order to slow down. For a Mars mission the HPP fires for about four hours before the spacecraft is out of range. By that time the spacecraft is travelling at about 20,000 m/s, which is fast enough to get to Mars in 50 days flat. The range is about 1×107 meters (ten thousand kilometers).

    After boosting a spacecraft, the HPP rotates the MagBeam in the opposite direction and uses it as an ion drive to move back into position. Newton's laws still hold, the recoil from the MagBeam is going to push the HPP way off base.

    And I'm quite sure that at short ranges the MagBeam can be used as a weapon. It would also be a nifity thing for a warship to mount, so it can use it to boost missiles to ferocious velocities.

    The main advantages seem to be increased acceleration levels on the spacecraft, and that one HPP propulsion unit can service multiple spacecraft. There are certain maneuvers that are impossible for low acceleration spacecraft, such as sub-orbital to orbital transfers, LEO to GEO transfers, LEO to escape velocity, and fast planetary missions.

    Plasma beams as a general rule have short ranges. However, the system can take advantage of the fact that both the HPP and the spacecraft have magnetic fields. The MagBeam uses magnetic fields to focus the beam and the spacecraft has a MagSail to catch the beam. If they start off close enough to each other, the two magnetic field merge ("magnetic reconnection"), and gradually stretch as the spacecraft moves. This creates a long magnetic tunnel to confine the plasma stream, making the stream self-focusing.

    This will be a problem when the HPP is faced with the task of slowing down an incoming spacecraft, since initially there will be no magnetic link. The spacecraft will have to temporarily inflate its MagSail, which can be done because it is an M2P2. Once the magnetic connection is made the M2P2 can be deflated to normal size.

    Plasma will probably be argon or nitrogen. The beam range will a few thousand kilometers if the HPP or the beam passes through the ionosphere, tens of thousands of kilometers if in the magnetosphere. This is because of the ambient plasma and magnetic fields in the ionosphere.

    Since the spacecraft does not carry its propellant, the standard rocket equation does not apply. Instead:

    HPPe = (0.25 * M * deltaV * Ve ) / HPPeff

    where:

    • HPPe = electrical energy expended by HPP (joules)
    • M = mass of spacecraft (kg)
    • deltaV = delta V applied to spacecraft (m/s)
    • HPPeff = efficiency of HPP at converting electricity into plasma energy (100% = 1.0, currently 0.6)

    Mpb = HPPe / (0.5 * Ve2)

    where:

    • Mpb = mass of propellant expended in HPP beam (kg)
    • HPPe = electrical energy expended by HPP (joules)
    • Ve = velocity of HPP beam (m/s)

    HPPpower = HPPe / Taccel

    where:

    • HPPpower = miminum power level of HPP power plant (watts)
    • HPPe = electrical energy expended by HPP (joules)
    • Taccel = duration of HPP beam usage (sec)

    So if a HPP had to boost a 10,000 kg (10 metric ton) spacecraft to a deltaV of 3,000 m/s (3 km/s) using a plasma beam with a velocity of 19,600 m/s (2000 s) had only 300 seconds (5 minutes) to do so and had an efficiency of 0.6 (60%), then the electrical power used would be 2.5×1010 joules, the power plant would need a level of 82,000,000 watts (82 megawatts), and 127 kilograms of propellant would be expended.

    Photon Sail

    Photon Sail
    Thrust per sail area9 N/km2
    Thrust by Sol dist1/R2

    A Photon Sail is a sail powered by solar photons. Commonly called a "solar sail", but that term does not make it clear if the sail is powered by solar photons, solar magnetic field, or solar wind.

    Kite Photon Sail
    Kite Photon Sail
    Max Thrust182 N
    @ 1 AU
    Useful Thrust69 N
    @ 1 AU
    Mass16,000 kg

    The simplest way to hold a sail out to catch sunlight is to use a rigid structure, much like a kite. The columns and beams of such a structure form a three-axis stabilization, so-named because all three dimensions are rigidly supported.

    Kite sails are easier to maneuver than sails that support themselves by spinning. By tilting the sail so that the light pressure slows the vessel down in its solar orbit will cause an inward spiral towards the sun. Tilting the opposite way will cause an outward spiral.

    The kite sail shown has a has a mast, four booms, and stays supporting a square sail 4 km to a side. At 93% reflectance, it develops a maximum thrust of 182 newtons at 1 AU. Control is provided by 4 steering vanes of 20,000 m2 area each. The unloaded mass is 16,000 kg and the unloaded sail loading is 0.5 g/m2.

    The film is 300 nm aluminum. Its microstructure is formed by DNA scaffolding, which is then coated with aluminum and the DNA baked off. This leaves holes the size of the wavelength of visible light, which makes the film lighter. The perforated film is thermally limited to 600K, and cannot operate in an Earth orbit lower than 1000 km due to air drag.

    Its thrust can augmented by the illumination of the 60 MW laser beam which is standard in this game. Operating at 50 Hz, this beam boils off water coolant replenished through capillary action in the perforated film. Tiny piezoelectric robot sailmakers repair ablated portions of the sail using vapor-deposited aluminum.

    Twice the size of Garvey’s “Large Square Rigged Clipper Sail”, and adding the perforation feature: J. M. Garvey, "Space station options for constructing advanced solar sails capable of multiple mars missions", AIAA Paper 87-1902, AIAA/SAE/ASME 23rd Joint Propulsion Conference,1987.

    From High Frontier by Philip Eklund
    Photon Heliogyro Sail
    Photon Heliogyro Sail
    Thrust140 N
    @ 1 AU
    Number Blades192
    Mass40,000 kg

    A heliogyro is a photon sail consisting of multiple spinning blades. Its blades are rigidified by centrifugal force and pitched to provide attitude control, much like a helicopter.

    Although a spinning design does not need the struts of a kite sail, the centrifugal loads generated must be carried by edge members in the blades. Moreover oscillations are created when the sail’s attitude changes, which need to be restrained by transverse battens. Small sail panels prevent wrinkling from the curvature in edge members between the battens.

    For these reasons, the heliogyro has no mass advantage over a kite sail, but it has the advantage of easier deployment in space.

    The reference design at 1 AU generates 140 newtons maximum thrust from 4 banks of 48 blades each. Each blade has a dimension of 8 × 7500 meters. This thrust is quite low (about 31 lbs), but its game performance is comparable to an electric rocket since its impulse is imparted over a full year rather than a few days.

    The sail film is 1 μm thick with reflective and emissive coatings. Each bank is fixed to a hub so the members co-rotate. The combined film masses 7 tonnes alone, and with the supporting cables masses 40 tonnes.

    Scaled up from the JPL Halley Rendezvous design: Jerome Wright, “Space Sailing”, 1992.

    From High Frontier by Philip Eklund

    Plasma Magnet

    An plasma Magnet is a type of E-sail powered by solar wind.

    Other

    { Beer }

    Beer
    Thrust Power8 × 10-8 GW
    Exhaust velocity83 m/s
    Thrust84 n
    T/W >1.0no

    In The Makeshift Rocket (also known as A Bicycle Built for Brew), the old geezer cobbles together a crude rocket out of hogs-heads of pressurized beer in order to escape to an adjacent asteroid.

    (ed note: I asked Rob Davidoff for an estimate of the performance of beer.)

    Thrust = velocity * mass_flow

    Assume we model the system as the fluid starting from stagnation (V-o = 0) under pressure P_o and accelerating to a vacuum pressure P_2 = 0 at velocity v_1. We can then employ Bernoulli's equation, which says the following once we knock out some irrelevant terms:

    P_o = 0.5 * rho * (V_1)2

    Solve for V_1:

    V_1 = sqrt( 2 * P_o / rho)

    So, what's a reasonable pressure? Sheesh, I dunno. A standard fuel-driven rocket engine operates at about 35 atm for a very low-pressure combustion, let's try that. Using the density of water (1000 kg/m3), I get...84 m/s. Isp of 8.5 seconds or so. The thrust will be this times the mass flow, so 1 kg/s would give 84 Newtons.

    Is this any use? It's pretty crappy, but maybe it's good enough. Say he needs, oh, 150 m/s. That's a mass ratio of 6, which isn't terrible. To lift off from an asteroid, you basically need a T/W of anything non-zero, so it's workable. Of course, keeping beer pressurized to 35 atmospheres was the starting assumption, any maybe that was a little high.

    However, the big issue is the density of the beer. Substitute in an air-like gas with a density of 1.4 kg/m2 instead of 1000, and you get to an Isp of ~220s, instead of 8. That's a lot more like it.

    Rob Davidoff

    Mass Driver

    Mass Driver
    Exhaust Velocity30,000 m/s
    Specific Impulse3,058 s
    Thrust20,000 N
    Thrust Power0.3 GW
    Mass Flow0.67 kg/s
    Total Engine Mass150,000 kg
    T/W0.01
    Thermal eff.90%
    Total eff.90%
    Fuel800MWe input
    RemassRegolith
    Remass AccelElectromagnetic
    Acceleration
    Specific Power500 kg/MW

    Mass drivers: magnetic buckets filled with packed rock dust are accelerated electmagnetically. Buckets are recovered for re-use. Propellant is rock dust or anything else you can stuff into the bucket. Popular with asteroid miners who want to nudge their claims into different orbits. However, their existence may prompt the creation of an Orbital Guard.

    In Gerard O'Neill's plan for L5 colonies, mass drivers were used to deliver raw materials mined on Luna into orbit for colony construction. But instead of the mass driver being mounted on a cargo rocket, it was instead a ground installation near the lunar mine. The buckets were filled not with rock propellant, but instead with cargo cannisters of raw materials. The mass driver shot the cannisters into orbit. The cannisters were intercepted by a "catcher" at the colony site. So instead of needing a fleet of cargo rockets, you just needed a mass driver launcher and a catcher.

    A mass driver is an electromagnetic mass accelerator that is optimized for propulsion. If you optimize it as a weapon instead, you have a coil-gun or rail gun. The weapons still have recoil and can be used as a crude propulsion system.

    Mass Driver
    Mass Driver
    Exhaust Velocity9,810 m/s
    Specific Impulse1,000 s
    Thrust10,400 N
    Thrust Power51.0 MW
    Mass Flow1 kg/s
    Total Engine Mass163,000 kg
    T/W7.00e-03
    Thermal eff.85%
    Total eff.85%
    Fuel60MWe input
    RemassRegolith
    Remass AccelElectromagnetic
    Acceleration
    Specific Power3,195 kg/MW

    An electrodynamic traveling-wave accelerator can be used as either a thruster or a payload launcher.

    The reaction mass or payload is loaded into a lightweight bucket banded by a pair of superconducting loops acting as armatures of a linear-electric guideway. The thruster illustrated accelerates the bucket at 75,000 gee's, utilizing 7 GJ of electromagnetic energy stored inductively in superconducting coils. The trackway length is 390 meters. One 36kg of reaction mass is ejected each minute at 15 km/sec. The bucket is decelerated and recovered. Cryogenic 77 K radiators cool the superconductors.

    A mass-driver optimized for materials transport rather than for propulsion uses a higher ratio of payload mass to bucket mass. With a 54% duty cycle, this system can launch 10 kt/yr of factory products. Coupled with a pointing accuracy in the tens of microradians, this can launch payloads or projectiles to targets millions of kilometers distant. A terrestrial mass driver running up the side of an equatorial mountain can launch payloads at the Earth escape velocity (11 km/sec). Imparted with a launch energy of 76 GJ, a one tonne payload the size and shape of a telephone pole with a carbon cap would burn up only 3% of its mass and lose only 20% of its energy on its way to solar or Earth orbit.

    Gerard K. O’Neill, “The High Frontier: Human Colonies in Space,” 1977.

    From High Frontier by Philip Eklund

    “Anjeä SysCon, this is VS Ardent Voyager, gated in-system from Loxix, identifying. Over.”

    “Ardent Voyager, Anjeä SysCon, we have you arriving at 5173-09-14:7-51-11; squawk ident. Welcome to Imperial space, please specify your intentions. Over.”

    “Anjeä SysCon, Ardent Voyager. Request through-clearance for immediate transit to Conclave System, minimum delta transfers. Over.”

    “Wait one, Ardent Voyager… Voyager, please confirm your hull class and propulsion. Over.”

    “Anjeä SysCon, we are a beehive habitat with reserve mass driver propulsion. Over.”

    “In other words, Ardent Voyager, you’re flying an asteroid and moving by throwing rocks. With regret, please shut down all active drive systems immediately. You are denied transit permission under power. Over.”

    “Anjeä SysCon, we are a diplomatic vessel and have the right of transit to Conclave System. Over.”

    “Ardent Voyager, you have the right of transit, but that doesn’t exempt you from the rules of navigation. Over.”

    “Anjeä SysCon, what’s your problem with us? Nowhere else has refused us transit. Over.”

    “Ardent Voyager, this is a crowded system with too damn many loose rocks anyway, see? We don’t want any accidents, and a drive like yours is a flyin’ invitation to accidents, or a hefty cleanup bill. It’s a miracle you got clearance to transit this far. Over.”

    “Anjeä SysCon, what are we supposed to do, then, just sit here? Over.”

    “Ardent Voyager, hire a tug? Either to finish out your voyage or jump back out-system, but either way, you’re not runnin’ that hazard to navigation anywhere in our sky. SysCon, clear.”

    - overheard on system space-control channel, Anjeä (High Verge)

    Photon

    Photon
    Exhaust Velocity299,792,458 m/s
    Specific Impulse30,559,884 s
    Fuel1.1TWe input
    RemassPhotons
    Remass AccelElectromagnetic
    Acceleration

    The exhaust is not a stream of matter. Instead it is a beam of Electromagnetic radiation, basically a large laser. The advantage is that it has the maximum possible exhaust velocity and thus the highest specific impulse. The disadvantage is the ludicrously high power requirements.

    The momentum of a photon is p = E/c, where E is the energy of the photon. So the thrust delivered by a stream of photons is ∂p/∂t = ∂E/∂t/c. This boils down to:

    F = P/c

    P = F * c

    where:

    • F = thrust in Newtons
    • P = power in watts
    • c = speed of light in a vacuum (3e8 m/s)

    In other words, one lousy Newton of thrust takes three hundred freaking megawatts!!

    Watch the Heat

    From my limited understanding, the basic problem is how to keep the engine from vaporizing.

    Fp = (F * Ve ) / 2

    where

    • Fp = thrust power (watts)
    • F = thrust (newtons)
    • Ve = exhaust velocity (m/s)

    The problem is that at high enough values for exhaust velocity and thrust, the amount of watts in the jet is too much. "Too much" is defined as: if only a fractional percentage of those watts are lost as waste heat, the spacecraft glows blue-white and evaporates. The size of the dangerous fractional percent depends on heat protection technology. There is a limit to how much heat that current technology can deal with, without a technological break-through.

    Jerry Pournelle says (in his classic A STEP FARTHER OUT) that an exhaust velocity of 28,800,000 cm/s corresponds to a temperature of 5 million Kelvin.

    As an exceedingly rough approximation:

    Ae = (0.5 * Am * Av2) / B

    where

    • Ae = particle energy (Kelvin)
    • Am = mass of particle (g) (1.6733e-24 grams for monatomic hydrogen)
    • Av = exhaust velocity (cm/s)
    • B = Boltzmann's constant: 1.38e-16 (erg K-1)

    (note that the above equation is using centimeters per second, not meters per second)

    A slightly less rough approximation:

    Qe = (Ve / (Z * 129))2 * Pw

    where

    • Qe = engine reaction chamber temperature (Kelvin)
    • Ve = exhaust velocity (m/s)
    • Z = heat-pressure factor, varies by engine design, roughly from 1.4 to 2.4 or so.
    • Pw = mean molecular weight of propellant, 1 for atomic hydrogen, 2 for molecular hydrogen

    The interiors of stars are 5 million Kelvin, but few other things are. How do you contain temperatures of that magnitude? If the gadget is something that can be mounted on a ship smaller than the Queen Mary, it has other implications. It is an obvious defense against hydrogen bombs, for starters.

    Larry Niven postulates something like this in his "Known Space" series, the crystal-zinc tube makes a science-fictional force field which reflects all energy. Niven does not explore the implications of this. However, Niven and Pournelle do explore the implications in THE MOTE IN GOD'S EYE. The Langson Field is used in the ship's drive, and as a force screen defense. The Langson field absorbs energy, and can re-radiate it. As a defense it sucks up hostile laser beams and nuclear detonations. As a drive, it sucks up and contains the energy of a fusion reaction, and re-radiates the energy as the equivalent of a photon drive exhaust.

    (And please remember the difference between "temperature" and "heat". A spark from the fire has a much higher temperature than a pot of boiling water, yet a spark won't hurt your hand at all while the boiling water can give you second degree burns. The spark has less heat, which in this context is the thrust power in watts.)

    If one has no science-fictional force fields, as a rule of thumb the maximum heat load allowed on the drive assembly is around 5 MW/m2. This is the theoretical ultimate, for an actual propulsion system it will probably be quite a bit less. For a back of the envelope calculation:

    Rc = 0.12 * sqrt[H]

    where

    • Rc = reaction chamber radius (meters)
    • H = reaction chamber waste heat (megawatts)

    (this equation courtesy of Anthony Jackson)

    Example

    Say your propulsion system has an exhaust velocity of 5.4e6 m/s and a thrust of 2.5e6 N. Now Fp=(F*Ve)/2 so the thrust power is 6.7e12 W. So, 6.7e12 watts divided by 1.0e6 watts per megawatt gives us 6.7e6 megawatts. Plugging this into the equation results in 0.12 * sqrt[6.7e6 MW] = drive chamber radius of 310 meters or a diameter of a third of a mile. Ouch.

    As a first approximation, for most propulsion systems one can get away with using the thrust power for H. Science-fictional technologies can cut the value of H to a percentage of thrust power by somehow preventing the waste heat from getting to the chamber walls.

    Only use this equation if H is above 4,000 MW or so, and if the propulsion system is a thermal type (i.e., fission, fusion, or antimatter).


    Playing with these figures will show that enclosing a thermal torch drive inside a reaction chamber made of matter appears to be a dead end. Unless you think a drive chamber a third of a mile in diameter is reasonable.

    Therefore, the main strategy is to try and direct the drive energy with magnetic fields instead of metal walls. The metal framework lets the heat escape instead of vaporizing the nozzle. The magnetic field cannot be vaporized since it is composed of energy instead of matter.

    This is gone into in more detail here in the Torchship section.


    And don't forget the Kzinti Lesson.

    Calculating the performance of a spaceship can be complicated. But if the ship is powerful enough, we can ignore gravity fields. It is then fairly easy. The ship will accelerate to a maximum speed and then turn around and slow down at its destination. Fusion or annihilation-drive ships will probably do this. They will apply power all the time, speeding up and slowing down.(ed note: a "brachistochrone" trajectory)

    In this simple case, all the important performance parameters can be expressed on a single graph. This one is drawn for the case when 90% of the starting mass is propellant. (ed note: a mass ratio of 10) Jet velocity (exhaust velocity) and starting acceleration are the graph scales. Distance for several bodies are shown. Mars varies greatly; I used 150 million kilometers. Trip times and specific power levels are also shown. "Specific power" expresses how much power the ship generates for each kilogram of its mass, that is, its total power divided by its mass. The propellant the ship will carry is not included in the mass value.

    An example: Suppose your ship can produce 100 kW/kg of jet power. You wish to fly to Jupiter. Where the 100 kW/kg and Jupiter lines cross on the graph, read a jet velocity of 300,000 m/s (Isp = 30,000) and an initial acceleration of nearly 0.01g. Your trip will take about two months.

    The upper area of the graph shows that high performance is needed to reach the nearest stars. Even generation ships will need, in addition to very high jet velocities, power on the order of 100 kW/kg. The space shuttle orbiter produces about 100 kW/kg with its three engines. The high power needed for starflight precludes its attainment with means such as electric propulsion.

    Gordon Woodcock

    Atomic Rockets notices

    Welcome to the improved Atomic Rockets!

    Atomic Rockets

    Support Atomic Rockets

    Support Atomic Rockets on Patreon