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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 883 records · Page 49
All sourcesNASA NTRSDOE OSTI
NASA NTRS1988

Space-based laser-powered orbital transfer vehicle (Project SLICK)

The project SLICK (Space Laser Interorbital Cargo Kite) involves conceptual designs of reusable space-based laser-powered orbital transfer vehicle (LOTV) for ferrying 16,000 kg cargo primarily between low Earth orbit (LEO) and geosynchronous earth orbit (GEO). The power of LOTV is beamed by a single 32-MW solar-pumped iodide laser orbiting the Earth at an altitude of one Earth radius. The laser engine selected for the LOTV is based on a continuous-wave, steady-state propulsion scheme and uses an array of seven discrete plasmas in a flow of hydrogen propellant. Both all-propulsive and aerobraked LOTV configurations were analyzed and developed. The all-propulsive vehicle uses a rigid 11.5-m aperture primary mirror and its engine produces a thrust of 2000 N at a specific impulse of 1500 sec. For the LEO-to-GEO trip, the payload ratio, m(sub payload/m(sub propellant)+m(sub dry vehicle) = 1.19 and the trip time is about 6 days. The aerobraked version uses a lightweight, retractable wrapped-rib primary mirror which is folded for aerobraking and a 20-m-diameter inflatable-ballute aeroshield which is jettisoned after aeromaneuver. Lifecycle cost analysis shows that the aerobraked configuration may have an economic advantage over the all-propulsive configuration as long as the cost of launching the propellant to LEO is higher than about $500/kg in current dollars.

Source record↗
NASA NTRS1951

Investigation of the Liquid Fluorine-liquid Diborane Propellant Combination in a 100-pound-thrust Rocket Engine

The experimental performance of liquid fluorine and liquid diborane was investigated in a 100-pound-thrust engine at a combustion pressure of 300 pounds per square inch absolute. Methods of handling and transporting liquid fluorine were developed. It was extremely difficult to obtain satisfactory operation because of the high flame speed and high combustion chamber temperatures. The maximum performance obtained was 280 pound seconds per pound, 88 percent of the theoretical maximum. The theoretical performance was recalculated with revised thermodynamic data, indicating a maximum specific impulse of 311 pound seconds per pound as compared with the previously reported value of 323.

COMBUSTION - ROCKET ENGINES↗
NASA NTRS1955

Flight and Preflight Tests of a Ram Jet Burning Magnesium Slurry Fuel and Utilizing a Solid-propellant Gas Generator for Fuel Expulsion

Data obtained from the first flight test of a ram jet utilizing a magnesium slurry fuel are presented. The ram jet accelerated from a Mach number of 1.75 to a Mach number of 3.48 in 15.5 seconds. During this period a maximum values of air specific impulse and gross thrust coefficient were calculated to be 151 seconds and 0.658, respectively. The rocket gas generator used as a fuel-pumping system operated successfully.

Bartlett, Walter, A , jr↗
NASA NTRS1958

Theoretical Combustion Performance of Several High-Energy Fuels for Ramjet Engines

An analytical evaluation of the air and fuel specific-impulse characteristics of magnesium, magnesium octene-1 slurries, aluminum, aluminum octene-1 slurries, boron, boron octene-1 slurries, carbon, hydrogen, alpha-methylnaphthalene, diborane, pentaborane, and octene-1 is presented. While chemical equilibrium was assumed in the combustion process, the expansion was assumed to occur at fixed composition.

Tower, Leonard K↗
NASA NTRS1994

Program For Optimization Of Nuclear Rocket Engines

NOP is versatile digital-computer program devoloped for parametric analysis of beryllium-reflected, graphite-moderated nuclear rocket engines. Facilitates analysis of performance of engine with respect to such considerations as specific impulse, engine power, type of engine cycle, and engine-design constraints arising from complications of fuel loading and internal gradients of temperature. Predicts minimum weight for specified performance.

Plebuch, R. K.↗
NASA NTRS1994

Study Of Fuels For Hybrid Rockets

Report describes experimental study of combustion and rates of regression of selected fuels for hybrid rocket engines. Part of continuing effort to develop fuels with greater rates of regression and lesser dependence on shapes of fuel grains and to maximize potential specific impulse at low cost.

Strand, Leon D.↗
NASA NTRS1993

Arcjet space thrusters

Electrically powered arc jets which produce thrust at high specific impulse could provide a substantial cost reduction for orbital transfer and station keeping missions. There is currently a limited understanding of the complex, nonlinear interactions in the plasma propellant which has hindered the development of high efficiency arc jet thrusters by making it difficult to predict the effect of design changes and to interpret experimental results. A computational model developed at the University of Tennessee Space Institute (UTSI) to study laser powered thrusters and radio frequency gas heaters has been adapted to provide a tool to help understand the physical processes in arc jet thrusters. The approach is to include in the model those physical and chemical processes which appear to be important, and then to evaluate our judgement by the comparison of numerical simulations with experimental data. The results of this study have been presented at four technical conferences. The details of the work accomplished in this project are covered in the individual papers included in the appendix of this report. We present a brief description of the model covering its most important features followed by a summary of the effort.

Keefer, Dennis↗
NASA NTRS1993

A Soft-Start Circuit for Arcjet Ignition

The reduced propellant flow rates associated with high performance arcjets have placed new emphasis on electrode erosion, especially at startup. A soft-start current profile was defined which limited current overshoot during the initial 30 to 50 ms of operation, and maintained significantly lower than the nominal arc current for the first eight seconds of operation. A 2-5 kW arcjet PPU was modified to provide this current profile, and a 500 cycle test using simulated fully decomposed hydrazine was conducted to determine the electrode erosion during startup. Electrode geometry and mass flow rates were selected based on requirements for a 600 second specific impulse mission average arcjet system. The flow rate was varied throughout the test to simulate the blow down of a flight propellant system. Electrode damage was negligible at flow rates above 33 mg/s, and minor chamfering of the constrictor occurred at flow rates of 33 to 30 mg/s, corresponding to flow rates expected in the last 40 percent of the mission. Constrictor diameter remained unchanged and the thruster remained operable at the completion of the test. The soft-start current profile significantly reduced electrode damage when compared to state of the art starting techniques.

Hamley, John A.↗
NASA NTRS1993

Development of a high specific 1.5 to 5 kW thermal arcjet

A research and development project on the experimental study of a 1.5-5 kW thermal arcjet thruster was started in 1992 at the IRS. Two radiation cooled thrusters were designed, constructed, and adapted to the test facilities, one at each end of the intended power range. These thrusters are currently subjected to an intensive test program with main emphasis on the exploration of thruster performance and thruster behavior at high specific enthalpy and thus high specific impulse. Propelled by simulated hydrazine and ammonia, the thruster's electrode configuration such as constrictor diameter and cathode gap was varied in order to investigate their influence and to optimize these parameters. In addition, test runs with pure hydrogen were performed for both thrusters.

Riehle, M.↗
NASA NTRS1993

Segmented ion thruster

Apparatus and methods for large-area, high-power ion engines comprise dividing a single engine into a combination of smaller discharge chambers (or segments) configured to operate as a single large-area engine. This segmented ion thruster (SIT) approach enables the development of 100-kW class argon ion engines for operation at a specific impulse of 10,000 s. A combination of six 30-cm diameter ion chambers operating as a single engine can process over 100 kW. Such a segmented ion engine can be operated from a single power processor unit.

Brophy, John R.↗
NASA NTRS1993

Powdered aluminum and oxygen rocket propellants: Subscale combustion experiments

Aluminum combined with oxygen has been proposed as a potential lunar in situ propellant for ascent/descent and return missions for future lunar exploration. Engine concepts proposed to use this propellant have not previously been demonstrated, and the impact on performance from combustion and two-phase flow losses could only be estimated. Therefore, combustion tests were performed for aluminum and aluminum/magnesium alloy powders with oxygen in subscale heat-sink rocket engine hardware. The metal powder was pneumatically injected, with a small amount of nitrogen, through the center orifice of a single element O-F-O triplet injector. Gaseous oxygen impinged on the fuel stream. Hot-fire tests of aluminum/oxygen were performed over a mixture ratio range of 0.5 to 3.0, and at a chamber pressure of approximately 480 kPa (70 psia). The theoretical performance of the propellants was analyzed over a mixture ratio range of 0.5 to 5.0. In the theoretical predictions the ideal one-dimensional equilibrium rocket performance was reduced by loss mechanisms including finite rate kinetics, two-dimensional divergence losses, and boundary layer losses. Lower than predicted characteristic velocity and specific impulse performance efficiencies were achieved in the hot-fire tests, and this was attributed to poor mixing of the propellants and two-phase flow effects. Several tests with aluminum/9.8 percent magnesium alloy powder did not indicate any advantage over the pure aluminum fuel.

Meyer, Mike L.↗
NASA NTRS1993

Performance assessment of low pressure nuclear thermal propulsion

A low pressure nuclear thermal propulsion (LPNTP) system, which takes advantage of hydrogen dissociation/recombination, was proposed as a means of increasing engine specific impulse (Isp). The effect of hydrogen dissociation/recombination on LPNTP Isp is examined. A two-dimensional computer model was used to show that the optimum chamber pressure is approximately 100 psia (at a chamber temperature of 3,000 K), with an Isp approximately 15 s higher than at 1,000 psia. At high chamber temperatures and low chamber pressures, the increase in Isp is due to both lower average molecular weights caused by dissociation and added kinetic energy from monatomic hydrogen recombination. Monatomic hydrogen recombination increases the Isp more then hydrogen dissociation. Variations in the mole fraction of monatomic hydrogen are similar to variations in static pressure along the axial nozzle position. Most recombination occurs close to the nozzle throat. Practical variations in nozzle geometry have minimal impact on recombination. Other models which can simulate a wider range of nozzle designs should be used in the future. The uncertainty of the hydrogen kinetic reaction rates at high temperatures (approximately 3,000 K) affects the accuracy of the analysis and should be verified with simple bench tests.

Gerrish, H. P., Jr.↗
NASA NTRS1993

CFD modeling of microwave electrothermal thrusters

Microwave-heated plasmas in convergent nozzles are analyzed using a coupled Maxwell and Navier-Stokes solver to examine relevant issues associated with microwave thermal propulsion. Parametric studies are conducted to understand the effect of power, pressure, and plasma location with respect to the nozzle throat. For nozzles in the 0.5 to 3 N range with helium flow, results show that specific impulses up to 550-650 seconds are possible, with further increases being limited by severe wall-heating. Coupling efficiencies of over 90 percent are consistently obtained, with overall efficiencies ranging from 40 percent to 80 percent. Size scale-up studies-done by scaling the frequency from 2.45 GHz to 0.91 GHz-indicate that plasma migration toward the walls occurs more frequently for the lower frequency. Increasing the cavity aspect ratio and detuning the cavity are found to be effective ways of keeping the plasma on axis.

Schwer, Douglas A.↗
NASA NTRS1993

Extended duration lunar lander

Selenium Technologies has been conducting preliminary design work on a manned lunar lander for use in NASA's First Lunar Outpost (FLO) program. The resulting lander is designed to carry a crew of four astronauts to a prepositioned habitat on the lunar surface, remain on the lunar surface for up to 45 days while the crew is living in the habitat, then return the crew to earth via direct reentry and land recovery. Should the need arise, the crew can manually guide the lander to a safe lunar landing site, and live in the lander for up to ten days on the surface. Also, an abort to earth is available during any segment of the mission. The main propulsion system consists of a cluster of four modified Pratt and Whitney RL10 rocket engines that use liquid methane (LCH4) and liquid oxygen (LOX). Four engines are used to provide redundancy and a satisfactory engine out capability. Differences between the new propulsion system and the original system include slightly smaller engine size and lower thrust per engine, although specific impulse remains the same despite the smaller size. Concerns over nozzle ground clearance and engine reliability, as well as more information from Pratt and Whitney, brought about this change. The power system consists of a combination of regenerative fuel cells and solar arrays. While the lander is in flight to or from the moon, or during the lunar night, fuel cells provide all electrical power. During the lunar day, solar arrays are deployed to provide electrical power for the lander as well as electrolyzers, which separate some water back into hydrogen and oxygen for later use by the fuel cells. Total storage requirements for oxygen, hydrogen, and water are 61 kg, 551 kg, and 360 kg, respectively. The lander is a stage-and-a-half design with descent propellant, cargo, and landing gear contained in the descent stage, and the main propulsion system, ascent propellant, and crew module contained in the ascent stage. The primary structure for both stages is a truss, to which all tanks and components are attached. The crew module is a conical shape similar to that of the Apollo Command Module, but significantly larger with a height and maximum diameter of six meters.

Babic, Nikola↗
NASA NTRS1994

Power Electronics Development for the SPT-100 Thruster

Russian electric propulsion technologies have recently become available on the world market. Of significant interest is the Stationary Plasma Thruster (SPT) which has a significant flight heritage in the former Soviet space program. The SPT has performance levels of up to 1600 seconds of specific impulse at a thrust efficiency of 0.50. Studies have shown that this level of performance is well suited for stationkeeping applications, and the SPT-100, with a 1.35 kW input power level, is presently being evaluated for use on Western commercial satellites. Under a program sponsored by the Innovative Science and Technology Division of the Ballistic Missile Defense Organization, a team of U.S. electric propulsion specialists observed the operation of the SPT-100 in Russia. Under this same program, power electronics were developed to operate the SPT-100 to characterize thruster performance and operation in the U.S. The power electronics consisted of a discharge, cathode heater, and pulse igniter power supplies to operate the thruster with manual flow control. A Russian designed matching network was incorporated in the discharge supply to ensure proper operation with the thruster. The cathode heater power supply and igniter were derived from ongoing development projects. No attempts were made to augment thruster electromagnet current in this effort. The power electronics successfully started and operated the SPT-100 thruster in performance tests at NASA Lewis, with minimal oscillations in the discharge current. The efficiency of the main discharge supply was measured at 0.92, and straightforward modifications were identified which could increase the efficiency to 0.94.

Hamley, John A.↗
NASA NTRS1993

Inertial electrostatic confinement as a power source for electric propulsion

The potential use of an Inertial Electrostatic Confinement (IEC) power source for space propulsion has previously been suggested by the authors and others. In the past, these discussions have generally followed the charged-particle electric-discharge engine (QED) concept proposed by Bussard, in which the IEC is used to generate an electron beam which vaporizes liquid hydrogen for use as a propellant. However, in the present study, we consider an alternate approach, using the IEC to drive a conventional electric thruster unit. This has the advantage of building on the rapidly developing technology for such thrusters, which operate at higher specific impulse. Key issues related to this approach include the continued successful development of the physics and engineering of the IEC unit, as well as the development of efficient step-down dc voltage transformers. The IEC operates by radial injection of energetic ions into a spherical vessel. A very high ion density is created in a small core region at the center of the vessel, resulting in extremely high fusion power density in the core. Present experiments at the U. of Illinois in small IEC devices (less than 60-cm. dia.) have demonstrated much of the basic physics underlying this concept, e.g. producing approximately 10(exp 6) D-D neutrons/sec steady-state with deuterium gas flow injection. The ultimate goal is to increase the power densities by several orders of magnitude and to convert to D-He-3 injection. If successful, such an experiment would represent a milestone proof-of-principle device for eventual space power use. Further discussion of IEC physics and status will be presented with a description of the overall propulsion system and estimated performance.

Miley, George H.↗
NASA NTRS1994

Nuclear electric propulsion: A better, safer, cheaper transportation system for human exploration of Mars

NASA has completed a preliminary mission and systems study of nuclear electric propulsion (NEP) systems for 'split-sprint' human exploration and related robotic cargo missions to Mars. This paper describes the study, the mission architecture selected, the NEP system and technology development needs, proposed development schedules, and estimated development costs. Since current administration policy makers have delayed funding for key technology development activities that could make Mars exploration missions a reality in the near future, NASA will have time to evaluate various alternate mission options, and it appears prudent to ensure that Mars mission plans focus on astronaut and mission safety, while reducing costs to acceptable levels. The split-sprint nuclear electric propulsion system offers trip times comparable to nuclear thermal propulsion (NTP) systems, while providing mission abort opportunities that are not possible with 'reference' mission architectures. Thus, NEP systems offer short transit times for the astronauts, reducing the exposure of the crew to intergalactic cosmic radiation. The high specific impulse of the NEP system, which leads to very low propellant requirements, results in significantly lower 'initial mass in low earth orbit' (IMLEO). Launch vehicle packaging studies show that the NEP system can be launched, assembled, and deployed, with about one less 240-metric-ton heavy lift launch vehicle (HLLV) per mission opportunity - a very Technology development cost of the nuclear reactor for an NEP system would be shared with the proposed nuclear surface power systems, since nuclear systems will be required to provide substantial electrical power on the surface of Mars. The NEP development project plan proposed includes evolutionary technology development for nuclear electric propulsion systems that expands upon SP-100 (Space Power - 100 kw(e)) technology that has been developed for lunar and Mars surface nuclear power, and small NEP systems for interplanetary probes. System upgrades are expected to evolve that will result in even shorter trip times, improved payload capabilities, and enhanced safety and reliability.

Clark, John S.↗
NASA NTRS1992

Computational study of single-expansion-ramp nozzles with external burning

A computational investigation of the effects of external burning on the performance of single expansion ramp nozzles (SERN) operating at transonic speeds is presented. The study focuses on the effects of external heat addition and introduces a simplified injection and mixing model based on a control volume analysis. This simplified model permits parametric and scaling studies that would have been impossible to conduct with a detailed CFD analysis. The CFD model is validated by comparing the computed pressure distribution and thrust forces, for several nozzle configurations, with experimental data. Specific impulse calculations are also presented which indicate that external burning performance can be superior to other methods of thrust augmentation at transonic speeds. The effects of injection fuel pressure and nozzle pressure ratio on the performance of SERN nozzles with external burning are described. The results show trends similar to those reported in the experimental study, and provide additional information that complements the experimental data, improving our understanding of external burning flowfields. A study of the effect of scale is also presented. The results indicate that combustion kinetics do not make the flowfield sensitive to scale.

Yungster, Shaye↗