Space Transport and Engineering Methods/Thermal Engines

Strictly speaking combustion engines with expansion nozzles are also thermal engines since they rely on hot gases. For discussion purposes we placed that group on the previous page, and on this one look at non-combustion thermal engines. It is possible in theory to combine both into a single device, but this is not usually considered for practical design reasons. Performance is always important, and light molecules have higher exhaust velocity at a given temperature. Therefore the tendency is to use Hydrogen if practical since it is the lightest molecule. Liquid Hydrogen requires extremely cold temperatures (14K or -435F), so storage for long periods is not practical for a small tank. Larger systems have better surface to volume ratios and can use active cooling to keep it liquid.

Electro-thermal Engines
Electro-thermal methods convert externally supplied electric power to heating of the propellant.

36 Electric-Rail Rocket
Alternate Names: Electrothermal Ramjet

Type: Heated Gas Flow by Power Line

Description: High voltage electricity supplied by rails is shorted through a tungsten heat exchanger in the engine. This heats Hydrogen carried by the vehicle traveling between the rails. The rails are assumed to be set up on an incline. Performance would be the same as a nuclear thermal rocket, about 9 km/s exhaust velocity, since both use heated Hydrogen limited by the melting point of the engine components. It requires very high power levels for large vehicles: 44 MW / ton / g acceleration. Since power is coming from outside the vehicle, it can generate enough thrust for launch from Earth. Since the Hydrogen fuel is only stored for a short period, insulation and boiloff are not major problems. This method would compete with other ground accelerator type systems such as gas guns or electromagnetic coils.

Status: Concept only at this time.

Variations:

References:


 * Wilbur, P. J.; Mitchell, C. E.; Shaw, B. D. "Electrothermal Ramjet", AIAA paper number 82-1216 presented at AIAA/SAE/ASME 18th Joint Propulsion Conference, Cleveland, OH, 21-23 June 1982.

37 Resistojet
Alternate Names:

Type: Heated Gas Flow by Photovoltaic Array

Description: In this method sunlight generates electricity, which is used to heat gas passed over or through a heating element, often after catalytic decomposition of a storable fuel to a lighter gas. In principle it is similar to the Electric rail rocket, but is used for smaller thrusters in orbit with attached solar array as power source. This limits the thrust, so it is not powerful enough to use for launch. Compared to chemical thrusters it gets about 50% better exhaust velocity

Status: This method was used to extend the operating life of communications satellites, since they had large solar arrays for their main job. The recent tendency is to use ion thrusters, which have even better performance.

Variations:

References:


 * Louviere, Allen J. et al "Water-Propellant Resistojets for Man- Tended Platforms", NASA Technical Memorandum 100110, 1987.

Photo-thermal Engines
This group uses direct electromagnetic radiation (photons) from natural or artificial sources to heat the propellant.

38 Solar-Thermal Engine
Alternate Names:

Type: Heated Gas Flow by Solar Flux

Description: In this method sunlight is concentrated by a reflector or lens, and then heats an absorber. The absorber transfers heat to a working fluid, usually Hydrogen. The Hydrogen is then expanded through a nozzle. If the absorber is in the form of pipes, the exhaust velocity is limited to about 9 km/s. If the absorber is in the form of a particle bed, which does not require mechanical strength, refractory carbides can be used. Tantalum hafnium carbide, formula Ta4HfC5, is hypothesized to have a melting point of 4488 K (7619 F) and would set an upper limit to the particle bed method. There are obvious difficulties in testing that material as no other container could hold it when melted. The bed is rotated to keep the particles from being blown out, and Hydrogen flow is from the outside in, then out a nozzle. Sunlight is focused on the inside surface which is then the hottest point. Hydrogen dissociates above 3000K to single atoms, leading to exhaust velocities slightly above 10 km/s.

Solar concentrators can be very low mass, and use all of the solar spectrum. So they can reach higher power levels than electro-thermal thrusters. They are more suited to main propulsion rather than orbit maintenance as electro-thermal typically is. The direction of the Sun is usually not the same as the direction of thrust, and changes over time. So solar-thermal systems need a way to point the concentrators. One way to handle this is to roll the vehicle about the thrust axis plus pivot the concentrators about a perpendicular axis. The concentrators are typically large, so need to be assembled or unfolded in orbit.

Status: Components have been tested by the US Air Force. Ion and plasma thrusters have become preferred because of their 3-5 times higher exhaust velocity.

Variations:

References:


 * Gartrell, C. F. "Future Solar Orbital Transfer Vehicle Concept", IEEE Transactions on Aerospace Electronic Systems, vol AES-19 no 5 pp 704-10, 1983.

39 Laser-Thermal Engine
Alternate Names:

Type: Heated Gas Flow by Laser

Description: A laser beam from an external source is passed through a window into a chamber. It is then absorbed by a heat exchanger or is focused to create a laser-sustained plasma within the gas flow. Hot gas is then expelled through a nozzle. By using a powerful energy source external to the propellant, exhaust velocity of around 10 km/s can be reached with high thrust to mass. One method of doing this on Earth is with large, ground-based lasers. Alternately the lasers can be at ground level, and a directing mirror is located on top of a large tower. A vacuum pipe connects the two. The extra height avoids atmospheric distortion and allows more distance to the horizon. Use of laser propulsion only in an upper stage would allow smaller lasers than are required for a first stage system. Even so, powerful enough lasers are not available yet, which limits the use of the method. Other methods of supplying energy to a vehicle are likely to be less expensive.

Status: Concept only at present.

Variations:

References:


 * Abe, T.; Shimada, T. "Laser Assisted Propulsion System Experiment on Space Flyer Unit", 38th International Astronautical Federation Conference paper number IAF-87-298, 1987.
 * Abe, T.; Kuriki, K. "Laser Propulsion Test Onboard Space Station", Space Solar Power Review vol 5 no 2 pp 121-5, 1985.
 * Jones, L. W.; Keefer, D. R. "NASA's Laser Propulsion Project", Astronautics and Aeronautics, v 20 no 9 pp 66-73, 1982.

40 Laser Detonation-Wave Engine
Alternate Names:

Type: Plasma via Laser

Description: In this method the propellant is a solid block with a flat bottom. A first laser pulse evaporates a layer of propellant. A second, larger, pulse creates a plasma detonation wave, which shocks and heats the propellant layer. The layer expands against the base of the solid block of remaining propellant. The pulse pattern is repeated as soon as the plasma dissipates and the laser can reach the block again. Because no fuel tank overhead or engine is needed, the vehicles are potentially very cheap, but it requires a powerful laser to function. For example, a 10 kg vehicle accelerated at 2 g's using a 20 km/s exhaust velocity plasma wave would require a pulsed laser with an average power of 2 MW, while large industrial pulsed lasers are about 600 W. Since the propellant temperature is not limited by any container, it can be hotter than other thermal methods, and so have better performance.

10 kg is a feasible size for this type of vehicle, if the laser can be kept focused on it till it reaches orbit velocity. In theory this would deliver about 6 kg to orbit. For larger payloads, such as to carry passengers, the laser would need to scale to GigaWatt power levels, which has led to the common saying within space propulsion circles of "there is nothing wrong with laser propulsion except the lack of GigaWatt lasers". Similar power levels are generally required of all launch methods carrying metric ton or larger payloads from Earth. For example, the Space Shuttle Main Engines, of which the Space Shuttle Orbiter used three, each had a power of 9.2 GW.

Status: Some detonation experiments have been carried out at small scale in a laboratory.

Variations:

References:


 * Kare, J.T. "SDIO/DARPA Workshop on Laser Propulsion, Volume 1: Executive Summary" Lawrence Livermore National Laboratory report number DE87-003254, 1987.

41 Microwave Thermal Engine
Alternate Names:

Type: Heated Gas Flow via Microwaves

Description: For this method microwaves from an external source are absorbed by a heat exchanger or concentrated by a waveguide into the engine. Hydrogen flows through the engine, absorbs the energy, and then exits by a nozzle. A large phased microwave array on the ground can focus onto a rocket-sized area over a range of hundreds of kilometers. Given a way to couple the microwave energy to a working fluid such as Hydrogen, this type of propulsion could provide significant launch vehicle velocities. High power microwave amplifiers exist in a variety of forms with efficiencies up to 75% and power levels up to one megawatt. As compared to Laser Thermal the chief advantage is the availability of high power microwave sources at relatively low cost. A disadvantage is the much larger wavelength of microwaves vs lasers, so maintaining focus over a distance is harder.


 * Design Example: 10 meter diameter receiver, 5 cm wavelength, 1 km phased array, range = 200 km.

Status: Concept only so far.

Variations:

References:

Nuclear Thermal Engines
This group use a nuclear reactor to heat the propellant. They vary in the physical state of the reactor core (solid, liquid, or gas).

42 Solid Core Nuclear Engine
Alternate Names:

Type: Heated Gas Flow by Fission Reactor

Description: Hydrogen is heated by flowing through a critical Nuclear Reactor, then expelled through a nozzle at high velocity. The low molecular weight of Hydrogen gas allows a higher exhaust velocity, about 9 km/s, than combustion rockets. Advantages include high power levels and high total stored energy.

Issues for Nuclear Thermal include:
 * For single missions the energy of the reactor has barely been tapped by the time the hydrogen is consumed, leaving a now radioactive core to work around if refueling for another mission.
 * Radiation shielding is needed for crew and cargo. To some degree that is mitigated by the need for shielding from the natural space radiation environment.
 * Any type of nuclear device raises extensive safety and environmental issues, even if not activated until in orbit, and even if not really warranted for technical reasons. One way around this issue is to mine and use the fissionable materials away from Earth. For example, parts of the Lunar surface have Thorium concentrations of 10 parts per million. If used as reactor fuel, that can provide a net energy of 350 MJ/kg of unprocessed Lunar soil, or about 7 times the energy density of gasoline on Earth. The processed fuel of course will be 100,000 times higher energy density, but the unprocessed ore energy density is an indicator of the feasibility of mining it.

In comparison to ion and plasma thrusters, nuclear thermal has about 3-5 times lower exhaust velocity, but much higher thrust levels. The near-instant burns relative to orbit time for nuclear thermal vs. constant burn for electric thrusters reduces the latter's advantage by 30%, but are still are 2.1-3.5 higher. The choice of method would depend on the importance of fuel mass, which is usually high, so today electric thrusters are usually preferred. In comparison to solar thermal it has about the same performance in terms of exhaust velocity, but higher thrust levels.

History The two major American reactor development efforts in the 1960s were KIWI and NERVA. Together with the $328 million spent on technology development, $90 million spent on the Nuclear Rocket Development Station in Nevada, and $153 million on other test facilities, almost $1.4 billion (in then-year dollars) was spent on nuclear rocket development from 1955 to 1972 (see Appendix 2:Reference Data ). Although considerable engine testing was done, the problem of solid core fuel damage at high operating temperatures, which are desirable for performance, was not solved.

Status: Nuclear rockets reached the testing stage in the 1960's under the NERVA program. Lack of actual need in a mission and rising worries about anything nuclear led to a halt in development. Since then only minor studies have been done.

Variations:


 * LOX-Augmented Nuclear Thermal Rocket Propulsion This injects Oxygen after the Hydrogen is heated by the reactor core . This increases thrust by about a factor of 3, which is useful for initial launch, then transitions to pure Hydrogen later for higher efficiency. Exhaust velocity is lowered by about 1/3 when adding Oxygen due to the higher molecular weight of the resulting exhaust.
 * Particle Bed Nuclear Engine - Although the nuclear rocket program was stopped a number of years ago, more recent work at Brookhaven National Laboratories on fluidized particle bed reactors warrants their consideration for launch vehicles. The small particle size (.3 mm) allows high heat transfer rates to the working fluid, hydrogen, and hence potentially high thrust to weight ratios. The smaller particles also potentially solves the fuel damage problem, as there is less scope for cracking in a fine powder. Exhaust velocity is increased slightly to 10 km/s.

References:


 * Thomas, Ulrich "Nuclear Ferry - Cislunar Space Transportation Option of the Future", Space Technology (Oxford) v 7 no 3 pp 227-234, 1987.
 * Holman, R.R.; Pierce, B. L. "Development of NERVA reactor for Space Nuclear Propulsion", presented at AIAA/ASME/SAE/ASEE 22nd Joint Propulsion Conference, Huntsville, Alabama, 16-18 Jun 1986, AIAA paper number 86-1582, 1986.
 * Thom, K. et al "Physics and Potentials of Fissioning Plasmas for Space Power and Propulsion", Acta Astronautica vol 3 no 7-8 pp 505-16, Jul. -Aug. 1976.
 * DiStefano, E. "Space Nuclear Propulsion - Future Applications and Technology", 2nd Symposium on Space Nuclear Power Systems, Albequerque, New Mexico, 14 January 1985, pp 331-342, 1987.

43 Liquid Core Nuclear
Alternate Names:

Type: Heated Gas Flow by Fission Reactor

Description: In order to attain higher performance than a solid core rocket, the reactor core is raised to a high enough temperature to become liquid. Hydrogen is bubbled through the liquid, then exhausted out a nozzle. The Hydrogen is first used to cool the reactor container, and so the temperature limit is governed by that rather than the melting point of the core. Expected exhaust velocity is up to 13-15 km/s, but development and testing of this type of engine will be difficult, as test failures can easily squirt the core fluid out the nozzle.

Status: Currently a concept only.

Variations:

References:

44 Gas Core Nuclear
Alternate Names:

Type: Heated Gas Flow by Fission Reactor

Description: In this version the reactor core is hot enough to be in gaseous form. The Hydrogen flow is seeded with an absorbent material to directly absorb the thermal radiation from the core. The core is kept from leaking out the nozzle either by a transparent container (nuclear light bulb), a flow vortex, which uses the density difference between uranium and hydrogen, or by magnetic separation, which uses the ionization difference between the uranium and the hydrogen. Expected performance ranges from 15-20 km/s for a quartz container, up to 30-50 km/s for a flow vortex. The latter is in the range of ion or plasma thrusters, but development and testing would be as difficult as for liquid core due to the possibility of ejecting the core during tests. A solid core reactor producing electricity for an electric thruster would have the same performance as gas core nuclear, with fewer development problems.

Status: Currently a concept only.

Variations:

References:


 * Wikipedia article: Nuclear Lighbulb