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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 721 records · Page 40
All sourcesNASA NTRSDOE OSTI
NASA NTRS1986

Low power dc arcjet operation with hydrogen/nitrogen/ammonia mixtures

The effect of gas composition and ambient pressure on arcjet operation was determined. Arcjet operation in different facilities was also compared to determine the validity of tests in small facilities. Volt-ampere characteristics were determined for an arcjet using hydrogen/nitrogen mixtures (simulating both ammonia and hydrazine), hydrogen/nitrogen/ammonia mixtures, and pure ammonia as propellants at various flow rates. The arcjet had a typical performance of 450 sec specific impulse at 1 kW with hydrogen/nitrogen mixtures. It was determined that the amount of ammonia present in the gas stream had a significant effect on the arcjet volt-ampere characteristics. Also, hydrogen/nitrogen mixtures simulating ammonia gave arc characteristics approximately the same as those of pure ammonia. Finally, no differences in arc volt-ampere characteristics were seen between low and high ambient pressure operation in the same facility. A 3 to 5 V difference was seen when different facilities were compared, but this difference was probably due to differences in the voltage drops across the current connections, and not due to arcjet operational differences in the two facilities.

Hardy, Terry L.↗
NASA NTRS1987

Nuclear powered Mars cargo transport mission utilizing advanced ion propulsion

Nuclear-powered ion propulsion technology was combined with detailed trajectory analysis to determine propulsion system and trajectory options for an unmanned cargo mission to Mars in support of manned Mars missions. A total of 96 mission scenarios were identified by combining two power levels, two propellants, four values of specific impulse per propellant, three starting altitudes, and two starting velocities. Sixty of these scenarios were selected for a detailed trajectory analysis; a complete propulsion system study was then conducted for 20 of these trajectories. Trip times ranged from 344 days for a xenon propulsion system operating at 300 kW total power and starting from lunar orbit with escape velocity, to 770 days for an argon propulsion system operating at 300 kW total power and starting from nuclear start orbit with circualr velocity. Trip times for the 3 MW cases studied ranged from 356 kW to 413 days. Payload masses ranged from 5700 to 12,300 kg for the 300 kW power level, and from 72,200 to 81, 500 kg for the 3 MW power level.

Galecki, Diane L.↗
NASA NTRS1987

The NASA/USAF arcjet research and technology program

Direct current arcjets have the potential to provide specific impulses greater than 500 sec with storable propellants, and greater than 1000 sec with hydrogen. This level of performance can provide significant benefits for such applications as orbit transfer, station keeping, orbit change, and maneuvering. The simplicity of the arcjet system and its elements of commonality with state-of-the-art resistojet systems offer a relatively low risk transition to these enhanced levels of performance for low power (0.5 to 1.5 kW) station keeping applications. Arcjets at power levels of 10 to 30 kW are potentially applicable to orbit transfer missions. Furthermore, with the anticipated development of space nuclear power systems, arcjets at greater than 100 kW may become attractive. This paper describes the ongoing NASA/USAF program and describes major recent accomplishments.

Stone, James R.↗
NASA NTRS1987

Preliminary performance characterizations of an engineering model multipropellant resistojet for space station application

Presented are the results of a program to describe the operational characteristics of an engineering model multipropellant resistojet for application as an auxiliary propulsion system for the space station. Performance was measured on hydrogen, helium, methane, water (steam), nitrogen, air, argon, and carbon dioxide. Thrust levels ranged from 109 to 355 mN, power levels ranged from 167 to 506 W, and specific impulse values ranged from 93 to 385 sec, depending on the propellant, chamber pressure, and heater current level selected. Detailed thermal maps of the heater and heat exchanger were also obtained for operation with carbon dioxide.

Morren, W. Earl↗
NASA NTRS1987

Analysis of quasi-hybrid solid rocket booster concepts for advanced earth-to-orbit vehicles

A study was conducted to assess the feasibility of quasi-hybrid solid rocket boosters for advanced earth-to-orbit vehicles. Thermochemical calculations were conducted to determine the effect of liquid hydrogen addition, solids composition change plus liquid hydrogen addition, and the addition of an aluminum/liquid hydrogen slurry on the theoretical performance of a PBAN solid propellant rocket. The Space Shuttle solid rocket booster was used as a reference point. All three quasi-hybrid systems theoretically offer higher specific impulse when compared with the Space Shuttle solid rocket boosters. However, based on operational and safety considerations, the quasi-hybrid rocket is not a practical choice for near-term earth-to-orbit booster applications. Safety and technology issues pertinent to quasi-hybrid rocket systems are discussed.

Zurawski, Robert L.↗
NASA NTRS1987

An 8-cm ion thruster characterization

The performance of the Ion Auxiliary Propulsion System (IAPS) thruster was increased to thrust T = 32 mN, specific impulse I sub sp = 4062 s, and thrust-to-power ratio T/P = 33 mN/kW. This performance was obtained by increasing the discharge power, accelerating voltage, propellant flow rate, and chamber magnetic field. Adding a plenum and main vaporizer for propellant distribution was the only major change required in the thruster. The modified thruster characterization is presented. A cathode magnet assembly did not improve performance. A simplified power processing unit was designed and evaluated. This unit decreased the parts count of the IAPS power processing unit by a factor of ten.

Wessel, F. J.↗
NASA NTRS1987

A low-power arcjet cyclic lifetest

A cyclic lifetest of a low power dc arcjet thruster using a hydrogen/nitrogen propellant mixture simulating hydrazine is currently in progress. Over 300 hr of operation have been accumulated to date in 2 hr duty cycles at a power level of about 1.15 kW, approximating that available on commercial communications satellites. A burn-in period was carried out before consistent operation was attained. After this period, the arcjet operated in a very stable fashion from cycle to cycle. At the beginning of each cycle, there was a brief starting transient followed by a rapid rise to a steady-state voltage. The steady-state voltage increased by about 5 V over the first 95 cycles. After this, it increased by only 1 V through the remainder of the test. Thrust measurements taken before the life test and again after the completion of the 144th cycle showed that both thrust, specific impulse, and arc voltage had increased over this period of operation. No life limiting mechanisms were observed during the course of the testing.

Curran, Francis M.↗
NASA NTRS1987

The VISTA spacecraft: Advantages of ICF (Inertial Confinement Fusion) for interplanetary fusions propulsion applications

Inertial Confinement Fusion (ICF) is an attractive engine power source for interplanetary manned spacecraft, especially for near-term missions requiring minimum flight duration, because ICF has inherent high power-to-mass ratios and high specific impulses. We have developed a new vehicle concept called VISTA that uses ICF and is capable of round-trip manned missions to Mars in 100 days using A.D. 2020 technology. We describe VISTA's engine operation, discuss associated plasma issues, and describe the advantages of DT fuel for near-term applications. Although ICF is potentially superior to non-fusion technologies for near-term interplanetary transport, the performance capabilities of VISTA cannot be meaningfully compared with those of magnetic-fusion systems because of the lack of a comparable study of the magnetic-fusion systems. We urge that such a study be conducted.

Orth, Charles D.↗
NASA NTRS1987

Nuclear electric power for multimegawatt orbit transfer vehicles

Multimegawatt nuclear propulsion is an attractive option for orbit transfer vehicles. The masses of these platforms are expected to exceed the capability of a single launch from Earth necessitating assembly in space in a parking orbit. The OTV would transfer the platform from the parking orbit to the operational orbit and then return for the next mission. Electric propulsion is advantageous because of the high specific impulse achieved by the technology, 1000 to 5000 s and beyond, to reduce the propellant required. Nuclear power is attractive as the power system because of the weight savings over solar systems in the multimegawatt regime, and multimegawatts of power are required. A conceptual diagram is shown of an OTV with a command control module using electric thrusters powered from an SP-100 class nuclear reactor power system.

Casagrande, R. D.↗
NASA NTRS1987

Megawatt level electric propulsion perspectives

For long range space missions, deliverable payload fraction is an inverse exponential function of the propellant exhaust velocity or specific impulse of the propulsion system. The exhaust velocity of chemical systems are limited by their combustion chemistry and heat transfer to a few km/s. Nuclear rockets may achieve double this range, but are still heat transfer limited and ponderous to develop. Various electric propulsion systems can achieve exhaust velocities in the 10 km/s range, at considerably lower thrust densities, but require an external electrical power source. A general overview is provided of the currently available electric propulsion systems from the perspective of their characteristics as a terminal load for space nuclear systems. A summary of the available electric propulsion options is shown and generally characterized in the power vs. exhaust velocity plot. There are 3 general classes of electric thruster devices: neutral gas heaters, plasma devices, and space charge limited electrostatic or ion thrusters.

Jahn, Robert G.↗
NASA NTRS1988

An extended life and performance test of a low-power arcjet

An automated, cyclic life test was performed to demonstrate the reliability and endurance of a low power dc cycle arcjet thruster. Over 1000 hr and 500 on-off cycles were accumulated which would represent the requirements for about 15 years of on-orbit lifetime. A hydrogen/nitrogen propellant mixture was used to simulate decomposed hydrazine propellant and the power level was nominally 1.2 kW after the burn-in period. The arcjet operated in a very repeatable fashion from cycle to cycle. The steady state voltage increased by approximately 6 V over the first 300 hr, and then by only 3 V through the remainder of the test. Thrust measurements taken before, during, and after the test verified that the thruster performed in a consistent fashion throughout the tests at a specific impulse of 450 to 460 sec. Post-test component evaluation revealed limited erosion on both the anode and cathode. Other thruster components, including graphite seals, appeared undamaged.

Curran, Francis M.↗
NASA NTRS1988

Development of a liquid-fed water resistojet

A concept for a forced-flow once-through water vaporizer for application to resistojet thrusters was evaluated as an element of a laboratory model thruster and tested to investigate its operating characteristics. The vaporizer design concept employs flow swirling to attach the liquid flow to the boiler chamber wall, providing for separation of the two liquid phases. This vaporizer was modified with a nozzle and a centrally-located heater to facilitate vaporization, superheating, and expansion of the propellant, allowing it to function as a resistojet. Performance was measured at thrust levels ranging from 170 to 360 mN and at power levels ranging from 443 to 192 W. Maximum measured specific impulse was 192 sec.

Morren, W. Earl↗
NASA NTRS1987

A computer program for performance prediction of tripropellant rocket engines with tangential slot injection

For the development of a Heavy Lift Launch Vehicle (HLLV) several engines with different operating cycles and using LOX/Hydrocarbon propellants are presently being examined. Some concepts utilize hydrogen for thrust chamber wall cooling followed by a gas generator turbine drive cycle with subsequent dumping of H2/O2 combustion products into the nozzle downstream of the throat. In the Space Transportation Booster Engine (STBE) selection process the specific impulse will be one of the optimization criteria; however, the current performance prediction programs do not have the capability to include a third propellant in this process, nor to account for the effect of dumping the gas-generator product tangentially inside the nozzle. The purpose is to describe a computer program for accurately predicting the performance of such an engine. The code consists of two modules; one for the inviscid performance, and the other for the viscous loss. For the first module, the two-dimensional kinetics program (TDK) was modified to account for tripropellant chemistry, and for the effect of tangential slot injection. For the viscous loss, the Mass Addition Boundary Layer program (MABL) was modified to include the effects of the boundary layer-shear layer interaction, and tripropellant chemistry. Calculations were made for a real engine and compared with available data.

Dang, Anthony↗
NASA NTRS1987

Liquid Engine Jet Atomization Workshop report

The workshop was oriented to disclose information and unsettled problems to understand the fundamental physical mechanism of the droplet formation process. Based on presentation and discussion of results, recommendations were made which should lead to associated future activities. To accomplish this task, existing observations and experiments, contributing to the basic knowledge, providing data for analytical concept verification, and forming a basis for empirical correlations were solicited. Advanced analytical modeling methods or results from specific studies were requested as well as the experience and advice from injector designers. All effort is directed to advance current analytical techniques, simulating the flow behavior downstream of the injection elements in a liquid rocket combustion chamber. Such a tool can be used to optimize injector designs with respect to short length weight savings, wall material protection, or large heat energy transport to a regenerative cooling fluid, while simultaneously achieving the maximum specific impulse in performance. The liquid atomization process also forms a sound basis for combustion instability analysis.

Gross, Klaus W.↗
NASA NTRS1988

Performance of 10-kW class xenon ion thrusters

Presented are performance data for laboratory and engineering model 30 cm-diameter ion thrusters operated with xenon propellant over a range of input power levels from approximately 2 to 20 kW. Also presented are preliminary performance results obtained from laboratory model 50 cm-diameter cusp- and divergent-field ion thrusters operating with both 30 cm- amd 50 cm-diameter ion optics up to a 20 kW input power. These data include values of discharge chamber propellant and power efficiencies, as well as values of specific impulse, thruster efficiency, thrust and power. The operation of the 30 cm- and 50 cm-diameter ion optics are also discussed.

Patterson, Michael J.↗
NASA NTRS1987

Technology challenges for the National Aero-Space Plane

The National Aerospace Plane (NASP) will require an exceptionally high degree of integration between propulsion and aerodynamic configuration, in order to achieve the requisite specific impulse and low structural weight. This is to be achieved through the use of forebody shock compression and afterbody exhaust expansion. Attention is presently given to the materials and structural concepts required for the realization of these NASP airframe functions, in view of the exceptionally high aerothermodynamic loads that will be experienced at hypersonic speeds. Active cooling will have to be used in certain critical airframe and propulsion components. CFD characterizations of these processes must be carefully developed and fully validated.

Piland, William M.↗
NASA NTRS1988

Detailed operating characteristics of a 30-kW ammonia arcjet engine with a contoured nozzle

Detailed experimental investigations were conducted to define the operating characteristics of a 30-kW type, arcjet engine with a contoured nozzle. The experiments were conducted in a fully instrumented facility with the engine mounted on a thrust stand. Engine performance was mapped over a power range of 10.6 to 23.0 kW with a mass flow range of 0.175 to 0.350 g/s. The data is presented in graphical form with the relationships between arc potential, arc current, power, thrust, specific impulse and thrust efficiency presented. This paper provides the first detailed accounting of arcjet operation with a contoured nozzle. Further theoretical and experimental work needs to be done to fully explain many of the features exhibited by the data.

Deininger, William D.↗
NASA NTRS1988

An extended life and performance test of a low-power arcjet

An automated, cyclic life test was performed to demonstrate the reliability and endurance of a low power dc cycle arcjet thruster. Over 1000 hr and 500 on-off cycles were accumulated which would represent the requirements for about 15 years of on-orbit lifetime. A hydrogen/nitrogen propellant mixture was used to simulate decomposed hydrazine propellant and the power level was nominally 1.2 kW after the burn-in period. The arcjet operated in a very repeatable fashion from cycle to cycle. The steady state voltage increased by aproximately 6 V over the first 300 hr, and then by only 3 V through the remainder of the test. Thrust measurements taken before, during, and after the test verified that the thruster performed in a consistent fashion throughout the tests at a specific impulse of 450 to 460 sec. Post-test component evaluation revealed limited erosion on both the anode and cathode. Other thruster components, including graphite seals, appeared undamaged.

Curran, Francis M.↗