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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 775 records · Page 43
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

Xenon ion propulsion for orbit transfer

For more than 30 years, NASA has conducted an ion propulsion program which has resulted in several experimental space flight demonstrations and the development of many supporting technologies. Technologies appropriate for geosynchronous stationkeeping, earth-orbit transfer missions, and interplanetary missions are defined and evaluated. The status of critical ion propulsion system elements is reviewed. Electron bombardment ion thrusters for primary propulsion have evolved to operate on xenon in the 5 to 10 kW power range. Thruster efficiencies of 0.7 and specific impulse values of 4000 s were documented. The baseline thruster currently under development by NASA LeRC includes ring-cusp magnetic field plasma containment and dished two-grid ion optics. Based on past experience and demonstrated simplifications, power processors for these thrusters should have approximately 500 parts, a mass of 40 kg, and an efficiency near 0.94. Thrust vector control, via individual thruster gimbals, is a mature technology. High pressure, gaseous xenon propellant storage and control schemes, using flight qualified hardware, result in propellant tankage fractions between 0.1 and 0.2. In-space and ground integration testing has demonstrated that ion propulsion systems can be successfully integrated with their host spacecraft. Ion propulsion system technologies are mature and can significantly enhance and/or enable a variety of missions in the nation's space propulsion program.

Rawlin, V. K.↗
NASA NTRS1990

Metallized Propellants for the Human Exploration of Mars

Advanced chemical propulsion using Metallized Propellants (MP) can lead to significant reductions in launch mass for piloted Mars missions. MP allow the propellant density or the specific impulse I(sub sp) of the propulsion system, or both, to increase. It can reduce the propellant mass and the propulsion system dry mass. Detailed mass-scaling equations and estimates of the I(sub sp) for several MP combinations are presented. The most significant savings with MP are derived from increasing the payload delivered to Mars. For the mass in low Earth orbit (LEO), a metallized Mars transfer vehicle can deliver 20 to 22 percent additional payload. This 20-percent payload increase reduces the total number of Mars flights and therefore significantly reduces the number of Space Transportation System-Cargo launches for the entire Mars architecture. Using MP to reduce the mass in LEO per flight is not as effective as increasing the payload delivery capacity. The mass saving per flight, while delivering the same payload with a higher I(sub sp) system, is much smaller. Using MP in all of the Mars propulsion systems would produce a modest 3.3 percent LEO mass saving. This translates into a saving of 38,000 kg over the mass required with O2/H2 propulsion. A Mars excursion vehicle using Earth- or space-storable propellants for the ascent can be an alternative to storing cryogenic H2 on Mars. A space-storable system using oxygen/monomethyl hydrazine/aluminum (O2/MMH/Al) would deliver the lowest mass penalty over O2/H2. For lower-energy expedition missions the LEO mass penalty for using metallized O2/MMH/Al would be only 3 to 5 percent.

Palaszewski, Bryan A.↗
NASA NTRS1990

Hypersonic vehicle simulation model: Winged-cone configuration

Aerodynamic, propulsion, and mass models for a generic, horizontal-takeoff, single-stage-to-orbit (SSTO) configuration are presented which are suitable for use in point mass as well as batch and real-time six degree-of-freedom simulations. The simulations can be used to investigate ascent performance issues and to allow research, refinement, and evaluation of integrated guidance/flight/propulsion/thermal control systems, design concepts, and methodologies for SSTO missions. Aerodynamic force and moment coefficients are given as functions of angle of attack, Mach number, and control surface deflections. The model data were estimated by using a subsonic/supersonic panel code and a hypersonic local surface inclination code. Thrust coefficient and engine specific impulse were estimated using a two-dimensional forebody, inlet, nozzle code and a one-dimensional combustor code and are given as functions of Mach number, dynamic pressure, and fuel equivalence ratio. Rigid-body mass moments of inertia and center of gravity location are functions of vehicle weight which is in turn a function of fuel flow.

Shaughnessy, John D.↗
NASA NTRS1990

Laboratory simulation of the rocket motor thrust as a follower force

Ground tests of solid propellant rocket motors have shown that metal-containing propellants produce various amounts of slag (primarily aluminum oxide), which is trapped in the motor case causing a loss of specific impulse. Although not yet definitely established, the presence of a liquid pool of slag also may contribute to nutational instabilities that have been observed with certain spin-stabilized, upper-stage vehicles. Because of the rocket's axial acceleration - absent in the ground tests - estimates of in-flight slag mass have been very uncertain. Yet such estimates are needed to determine the magnitude of the control authority of the systems required for eliminating the instability. A test rig with an eccentrically mounted hemispherical bowl was designed and built that incorporates a follower force that properly aligns the thrust vector along the axis of spin. A program that computes the motion of a point mass in the spinning and precessing bowl was written. Using various rpm, friction factors, and initial starting conditions, plots were generated showing the trace of the point mass around the inside of the fuel tank. The apparatus will be used extensively during the 1990 to 1991 academic year and incorporate future design features such as a variable nutation angle and a film height measuring instrument. Data obtained on the nutational instability characteristics will be used to determine order-of-magnitude estimates of control authority needed to minimize the sloshing effect.

Source record↗
NASA NTRS1990

High vacuum facility for hydrazine thruster testing

An ongoing modification is described of a large vacuum chamber to accommodate the ignition of an arcjet hydrazine thruster while maintaining a vacuum level of 1 x 10(exp -5) torr or less. The vacuum facility consists of a 20 ft stainless steel vacuum tank with an internal LN2 shroud, four 35 in. cryopumps and an 8 in. turbopump. To maintain a vacuum level of 1 x 10(exp -5) torr or less, 900 sq ft of liquid helium (LHe) shroud surface was installed to maintain the vacuum level and pumping requirements. A vacuum level of 1 x 10(exp -5) torr or less will allow the hydrazine thrust to exit the thruster nozzle and radiate into a space type environment so that the plume flow field can be analyzed and compared to the analytical model density distribution profile. Some other arcjet thruster characteristics measured are the electromagnetic interference (EMI) and exhaust contamination. This data is used to evaluate if the arcjet thruster with its high specific impulse in comparison to current chemical propulsion thruster can be used for the next generation of communication satellites.

Neary, Patrick F.↗
NASA NTRS1991

High power arcjet

The activities of the High Power Arcjet Project (HIPARC) from August 1990 to January 1991 are discussed. In this period the HIPARC thruster was ignited for the first time. Power levels up to 140 kW with a mass flow rate of 300 mg/s hydrogen were reached. Specific impulse values of more than 1300 s were shown to be possible. Tests were performed with the baseline thruster version only, which has a 6 mm throat diameter and a conical nozzle with a 20 degree half angle. Measurement data summing up all tests carried out until now is included. All measuring methods are described, including a check on possible error sources.

Goelz, T. M.↗
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.↗