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At least 793 records · Page 44

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NASA NTRS1991

Chemical approaches to carbon dioxide utilization for manned Mars missions

Use of resources available in situ is a critical enabling technology for a permanent human presence in space. A permanent presence on Mars, e.g., requires a large infrastructure to sustain life under hostile conditions. As a resource on Mars, atmospheric CO2 is as follows: abundant; available at all points on the surface; of known presence; chemically simple; and can be obtained by simple compression. Many studies focus on obtaining O2 and the various uses for O2 including life support and fuel; discussion of CO, the coproduct from CO2 fixation revolves around its uses as a fuel, being oxidized back to CO2. Several new proposals are studied for CO2 fixation through chemical, photochemical, and photoelectrochemical means. For example, the reduction of CO2 to hydrocarbons such as acetylene (C2H2) can be accomplished with H2. C2H2 has a theoretical vacuum specific impulse of approx. 375 secs. Potential uses were also studied of CO2, as obtained or further reduced to carbon, as a reducing agent in metal oxide processing to form metals or metal carbides for use as structural or power materials; the CO2 can be recycled to generate O2 and CO.

Hepp, Aloysius F.↗

NASA NTRS1991

Evaluation of on-board hydrogen storage methods f or high-speed aircraft

Hydrogen is the fuel of choice for hypersonic vehicles. Its main disadvantage is its low liquid and solid density. This increases the vehicle volume and hence the drag losses during atmospheric flight. In addition, the dry mass of the vehicle is larger due to larger vehicle structure and fuel tankage. Therefore it is very desirable to find a fuel system with smaller fuel storage requirements without deteriorating the vehicle performance substantially. To evaluate various candidate fuel systems, they were first screened thermodynamically with respect to their energy content and cooling capacities. To evaluate the vehicle performance with different fuel systems, a simple computer model is developed to compute the vehicle parameters such as the vehicle volume, dry mass, effective specific impulse, and payload capacity. The results indicate that if the payload capacity (or the gross lift-off mass) is the most important criterion, only slush hydrogen and liquid hydrogen - liquid methane gel shows better performance than the liquid hydrogen vehicle. If all the advantages of a smaller vehicle are considered and a more accurate mass analysis can be performed, other systems using endothermic fuels such as cyclohexane, and some boranes may prove to be worthy of further consideration.

Akyurtlu, Ates↗

NASA NTRS1990

The Liquid Annular Reactor System (LARS) propulsion

A concept for very high specific impulse (greater than 2000 seconds) direct nuclear propulsion is described. The concept, termed the liquid annular reactor system (LARS), uses liquid nuclear fuel elements to heat hydrogen propellant to very high temperatures (approximately 6000 K). Operating pressure is moderate (approximately 10 atm), with the result that the outlet hydrogen is virtually 100 percent dissociated to monatomic H. The molten fuel is contained in a solid container of its own material, which is rotated to stabilize the liquid layer by centripetal force. LARS reactor designs are described, together with neutronic and thermal-hydraulic analyses. Power levels are on the order of 200 megawatts. Typically, LARS designs use seven rotating fuel elements, are beryllium moderated, and have critical radii of approximately 100 cm (core L/D approximately equal to 1.5).

Powell, James↗

NASA NTRS1990

An approach to space power

Fusion offers the potential for a very high specific power, providing a large specific impulse that can be traded-off with thrust for mission optimization. Thus fusion is a leading candidate for missions beyond the moon. A new approach is discussed for space fusion power, namely Inertial Electrostatic Confinement (IEC). This method offers a high power density in a relatively small, simple device. It appears capable of burning aneutronic fuels which are most desirable for space applications and is well suited for direct conversion. An experimental device to test the concept is described.

Miley, G. H.↗

NASA NTRS1990

Pulsed plasmoid electric propulsion

A method of electric propulsion is explored where plasmoids such as spheromaks and field reversed configurations (FRC) are formed and then allowed to expand down a diverging conducting shell. The plasmoids contain a toroidal electric current that provides both heating and a confining magnetic field. They are free to translate because there are no externally supplied magnetic fields that would restrict motion. Image currents in the diverging conducting shell keep the plasmoids from contacting the wall. Because these currents translate relative to the wall, losses due to magnetic flux diffusion into the wall are minimized. During the expansion of the plasma in the diverging cone, both the inductive and thermal plasma energy are converted to directed kinetic energy producing thrust. Specific impulses can be in the 4000 to 20000 sec range with thrusts from 0.1 to 1000 Newtons, depending on available power.

Bourque, Robert F.↗

NASA NTRS1990

Laser driven launch vehicles for continuous access to space

The availability of megawatt laser systems in the next century will make laser launch systems from ground to orbit feasible and useful. Systems studies indicate launch capabilities of 1 ton payload per gigawatt laser power. Recent research in ground to orbit laser propulsion has emphasized laser supported detonation wave thrusters driven by repetitively pulsed infrared lasers. In this propulsion concept each laser repetition cycle consists of two pulses. A lower energy first pulse is used to vaporize a small amount of solid propellant and then after a brief expansion period, a second and higher energy laser pulse is used to drive a detonation wave through the expanded vapor. The results are reported of numerical studies comparing the detonation wave properties of various candidate propellants, and the simulation of thruster performance under realistic conditions. Experimental measurements designed to test the theoretical predictions are also presented. Measurements are discussed of radiance and opacity in absorption waves, and mass loss and momentum transfer. These data are interpreted in terms of specific impulse and energy conversion efficiency.

Rollins, C. J.↗

NASA NTRS1990

Comparison of super-high-energy-propulsion-systems based on metallic hydrogen propellant for ES to LEO space transportation

The application is studied of metallic H2 as a rocket propellant, which contains a specific energy of about 52 kcal/g in theory yielding a maximum specific impulse of 1700 s. With the convincing advantage of having a density 14 times that of conventional liquid H2/liquid O2 propellants, metallic H2 could satisfy the demands of advanced launch vehicle propulsion for the next millennium. Provided that there is an atomic metallic state of H2, and that this state is metastable at ambient pressure, which still is not proven, the results are given of the study of some important areas, which concern the production of metallic H2, the combustion, chamber cooling, and storage. The results show that the use of metallic H2 as rocket propellant could lead to revolutionary changes in space vehicle philosophy toward small size, small weight, and high performance single stage to orbit systems. The use of high metallic H2 mass fractions results in a dramatic reduction of required propellant volume, while gas temperatures in the combustion chamber exceed 5000 K. Furthermore, it follows, that H2 (liquid or slush) is the most favorable candidate as working fluid. Jet generated noise due to high exhaust velocities could be a problem.

Thierschmann, M.↗

NASA NTRS1990

Space Station auxiliary thrust chamber technology

The objective was to establish a technical data base to support future development of GO2/GH2 flight thrusters for a Space Station Auxiliary Propulsion System. Specific issues of concern were thruster performance and cycle life. To address these issues, NASA funded Aerojet to design, fabricate, and altitude test two 25-lbf GO2/GH2 thrusters. The first thruster was designed to operate at a nominal mixture ratio (O/F) of 4.0 and expansion area ratio (epsilon) of 100:1. It was tested over a range of O/F from 2.0 to 8.0, achieving a range of specific impulse (Isp) from 440 to 310 lbf-sec/Ibm. The second thruster was optimized for a nominal O/F of 8.0 at a lower nozzle expansion area ratio, epsilon, of 30:1. This second thruster was tested over an O/F range of 3.0 to 9.5, achieving an Isp range of 416 to 3323 lbf-sec/Ibm, respectively. At O/F = 8.0, the Isp was 360 lbf-sec/Ibm, as predicted.

Robinson, Philip J.↗

NASA NTRS1991

Carbon monoxide and oxygen combustion experiments: A demonstration of Mars in situ propellants

The feasibility of using carbon monoxide and oxygen as rocket propellants was examined both experimentally and theoretically. The steady-state combustion of carbon monoxide and oxygen was demonstrated for the first time in a subscale rocket engine. Measurements of experimental characteristic velocity, vacuum specific impulse, and thrust coefficient efficiency were obtained over a mixture ratio range of 0.30 to 2.0 and a chamber pressures of 1070 and 530 kPa. The theoretical performance of the propellant combination was studied parametrically over the same mixture ratio range. In addition to one dimensional ideal performance predictions, various performance reduction mechanisms were also modeled, including finite-rate kinetic reactions, two-dimensional divergence effects and viscous boundary layer effects.

Linne, Diane L.↗

NASA NTRS1991

Nuclear rocket propulsion. NASA plans and progress, FY 1991

NASA has initiated planning for a technology development project for nuclear rocket propulsion systems for space explorer initiative (SEI) human and robotic missions to the moon and Mars. An interagency project is underway that includes the Department of Energy National Laboratories for nuclear technology development. The activities of the project planning team in FY 1990 and 1991 are summarized. The progress to date is discussed, and the project plan is reviewed. Critical technology issues were identified and include: (1) nuclear fuel temperature, life, and reliability; (2) nuclear system ground test; (3) safety; (4) autonomous system operation and health monitoring; and (5) minimum mass and high specific impulse.

Clark, John S.↗

NASA NTRS1989

Unconventional propellant performance

Theoretical calculations are made of rocket performance of exotic fuels at various operating conditions such as chamber pressure, pressure ratios, and oxidizer-to-fuel ratios. By exotic fuels, it is meant using materials that may not normally be used as fuels here on Earth due to low performance. The majority of the work is currently being done on a VAX using the CET 86 program by Gordon and McBride. The CET 86 program will be installed on our IBM PS/2 Model 80 to reduce computer costs. The JPL Solid Propellant Theoretical Performance program for the IBM was also used, it handles solid fuels better than the CET 86 program. The CET 86 program by Gordon and McBride was set on the VAX and results were obtained for a number of the propellant combinations being analyzed. A sample of these results is included. Several programs were written to reduce the large amounts of data generated to include only the information needed, such as specific impulse, temperature, etc. This work involves analyzing propellants at various operating conditions and determining which will give acceptable performance and yet be easily storable for long periods of time and be readily available in space missions.

Ramohalli, Kumar↗

NASA NTRS1990

A figure-of-merit approach to extraterrestrial resource utilization

A concept is developed for interrelated optimizations in space missions that utilize extraterrestrial resources. It is shown that isolated (component) optimizations may not result in the best mission. It is shown that substantial benefits can be had through less than the best propellants, propellant combinations, propulsion hardware, and actually, some waste in the traditional sense. One ready example is the possibility of discarding hydrogen produced extraterrestrially by water splitting and using only the oxygen to burn storable fuels. The gains in refrigeration and leak-proof equipment mass (elimination) outweigh the loss in specific impulse. After a brief discussion of this concept, the synthesis of the four major components of any future space mission is developed. The four components are: orbital mechanics of the transportation; performance of the rocket motor; support systems that include power; thermal and process controls, and instruments; and in situ resource utilization plant equipment. This paper's main aim is to develop the concept of a figure-of-merit for the mission. The Mars Sample Return Mission is used to illustrate the new concept. At this time, a popular spreadsheet is used to quantitatively indicate the interdependent nature of the mission optimization. Future prospects are outlined that promise great economy through extraterrestrial resource utilization and a technique for quickly evaluating the same.

Ramohalli, K.↗

NASA NTRS1991

Development of an electrostatic propulsion engine using sub-micron powders as the reaction mass

Asteroid sample return missions would benefit from development of an improved rocket engine. Chemical rockets achieve their large thrust with high mass consumption rate (dm/dt) but low exhaust velocity; therefore, a large fraction of their total mass is fuel. Present day ion thrusters are characterized by high exhaust velocity, but low dm/dt; thus, they are inherently low thrust devices. However, their high exhausy velocity is poorly matched to typical mission requirements and therefore, wastes energy. A better match would be intermediate between the two forms of propulsion. This could be achieved by electrostatically accelerating solid powder grains, raising the possibility that interplanetary material could be processed to use as reaction mass. An experiment to study the charging properties of sub-micron sized powder grains is described. If a suitable material can be identified, then it could be used as the reaction mass in an electrostatic propulsion engine. The experiment employs a time of flight measurement to determine the exhaust velocity (v) of various negatively charged powder grains that were charged and accelerated in a simple device. The purpose is to determine the charge to mass ratio that can be sustained for various substances. In order to be competitive with present day ion thrusters, a specific impulse (v/g) of 3000 to 5000 seconds is required. Preliminary results are presented. More speculatively, there are some mission profiles that would benefit from collection of reaction mass at the remote asteroid site. Experiments that examine the generation of sub-micron clusters by electrostatic self-disruption of geologically derived material are planned.

Herbert, F.↗

NASA NTRS1991

Design issues for propulsion systems using metallized propellants

Metallized propellants are liquid propellants that contain metal particles. These particles are suspended in a gelled fuel or oxidizer. Aluminum is used as the metal additive. The addition of metal to conventional propellants can increase their specific impulse and their density over conventional propellants, and consequently, the payload delivered on Mars and lunar transportation vehicles, Earth-to-Orbit vehicles and upper stages for robotic planetary missions. Gelled fuels also provide increased safety during accidental propellant leakage or spills. To take full advantage of these potential performance increases, there are changes that must be made to the vehicle design. The differences are discussed between metallized propellant and traditional liquid propellants and their effect on the propulsion system design. These differences include the propellant density, mixture ratio, engine performance, and propellant rheology. Missions related to the Space Exploration Initiative are considered as design examples to illustrate these issues.

Palaszewski, Bryan↗

NASA NTRS1991

Performance and optimization of a derated ion thruster for auxiliary propulsion

The characteristics and implications of use of a derated ion thruster for north-south stationkeeping (NSSK) propulsion are discussed. A derated thruster is a 30 cm diameter primary propulsion ion thruster operated at highly throttled conditions appropriate to NSSK functions. The performance characteristics of a 30 cm ion thruster are presented, emphasizing throttled operation at low specific impulse and high thrust-to-power ratio. Performance data and component erosion are compared to other NSSK ion thrusters. Operations benefits derived from the performance advantages of the derated approach are examined assuming an INTELSAt 7-type spacecraft. Minimum ground test facility pumping capabilities required to maintain facility enhanced accelerator grid erosion at acceptable levels in a lifetest are quantified as a function of thruster operating condition. Approaches to reducing the derated thruster mass and volume are also discussed.

Patterson, Michael J.↗

NASA NTRS1991

Mass comparisons of electric propulsion systems for NSSK of geosynchronous spacecraft

A model was developed and exercised to allow wet mass comparisons of three axis stabilized communication satellites delivered to geosynchronous transfer orbit. The mass benefits of using advanced chemical propulsion for apogee injection and north-south stationkeeping (NSSK) functions or electric propulsion (hydrazine arcjets and xenon ion thrusters) for NSSK functions are documented. A large derated ion thrusters is proposed which minimizes thruster lifetime concerns and qualification test times when compared to those of smaller ion thrusters planned for NSSK applications. The mass benefits, which depend on the spacecraft mass and mission duration, increase dramatically with arcjet specific impulse in the 500 to 600 s range, but are nearly constant for the derated ion thruster operated in the 2300 to 3000 s range. For a given mission, the mass benefits with an ion system are typically double those of the arcjet system; however, the total thrusting time with arcjets is less than 1/3 that with ion thrusters for the same thruster power. The mass benefits may permit increases in revenue producing payload or reduce launch costs by allowing a move to a smaller launch vehicle.

Rawlin, Vincent K.↗

NASA NTRS1991

Medium power hydrogen arcjet performance

An experimental investigation was performed to evaluate hydrogen arcjet operating characteristics in the range of 1 to 4 kW. A series of nozzles were operated in modular laboratory thrusters to examine the effects of geometric parameters such as constrictor diameter and nozzle divergence angle. Each nozzle was tested over a range of current and mass flow rates to explore stability and performance. In the range of mass flow rates and power levels tested, specific impulse values between 650 and 1250 sec were obtained at efficiencies between 30 and 40 percent. The performance of the two larger half angle (20, 15 deg) nozzles was similar for each of the two constrictor diameters tested. The nozzles with the smallest half angle (10 deg) were difiicult to operate. A restrike mode of operation was identified and described. Damage in the form of melting was observed in the constrictor region of all the nozzle inserts tested. Arcjet ignition was also difficult in many tests and a glow discharge mode that prevents starting was identified.

Curran, Francis M.↗

NASA NTRS1991

MPD thruster technology

MPD (MagnetoPlasmaDynamic) thrusters demonstrated between 2000 and 7000 seconds specific impulse at efficiencies approaching 40 percent, and were operated continuously at power levels over 500 kW. These demonstrated capabilities, combined with the simplicity and robustness of the thruster, make them attractive candidates for application to both unmanned and manned orbit raising, lunar, and planetary missions. To date, however, only a limited number of thruster configurations, propellants, and operating conditions were studied. The present status of MPD research is reviewed, including developments in the measured performance levels and electrode erosion rates. Theoretical studies of the thruster dynamics are also described. Significant progress was made in establishing empirical scaling laws, performance and lifetime limitations and in the development of numerical codes to simulate the flow field and electrode processes.

Myers, Roger M.↗