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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 667 records · Page 37
All sourcesNASA NTRSDOE OSTI
NASA NTRS1984

The DE-AC accelerator system for high beam current density

A radically different ion accelerator system design has been developed to increase the extracted beam current density of ion thrusters. By independently controlling plasma ion extraction and acceleration, the Decoupled Extraction and Amplified Current (DE-AC) accelerator system gives higher beam currents at lower electric field stress levels. Results are presented for DE-AC grid system operation with the Test Bed Ion Engine. Using argon propellant, feasibility of a 60 mA/sq cm beam current density is shown. A discussion of power supply requirements, specific impulse operation and grid system geometrical design is presented also.

Aston, G.↗
NASA NTRS1985

Progress report - Advanced cryogenic OTV engine technology

New technologies for space-based, reusable, throttleable, cryogenic orbit transfer propulsion are being evaluated. A variable-thrust (200 to 3000 lbF), 2000 psi chamber pressure, LO2/LH2 engine has been selected to demonstrate the 20-hour, 500-restart life goal, and a specific impulse in excess of 480 lbF-sec/lbM. The results of recent vehicle-engine integration analyses and the progress in design, fabrication, and testing are provided. Emphasis is placed on the following technology areas being investigated in support of the advanced engine design: LOX hydrostatic bearings; burn-resistant materials for high-pressure GOX turbines and valves; high surface-low flux annular combustion chambers for the dual propellant expander cycle; improved cooling approaches for high-pressure combustion chambers, new concepts in integrated controls; and engine health diagnostics.

Schoenman, L.↗
NASA NTRS1985

Laser propulsion for the orbital transfer mission

America's space activities in the 1990s and beyond will partly consist of missions involving the transportation of cargo from low earth orbit (LEO) to higher orbits or to an escape trajectory. Such missions are to be performed with the aid of an orbit transfer vehicle (OTV). The operation of the OTV can be based on different propulsion concepts. A chemical OTV is characterized by a high thrust and low specific impulse. The result is a short trip time at the cost of large quantities of propellant. On the other hand, low-thrust systems such as electric propulsion units, consume very little propellant, but would have a long trip time. The present paper is concerned with a compromise between these two extremes. The employed propulsion system utilizes laser thermal propulsion, in which a ground or space-based laser is used to beam energy to a thruster on the OTV. The laser light is absorbed by a propellant. The resulting heating of the propellant causes an expansion of the propellant through a nozzle to produce thrust. Details regarding this propulsion concept are discussed, taking into account operational questions and missions.

Frisbee, R. H.↗
NASA NTRS1985

An experimental data base for material selection and design of high-speed, high-pressure, oxygen turbomachinery

New technologies for space-based, reusable, throttleable, cryogenic, orbit transfer propulsion are being evaluated. A variable-thrust (200 to 3000 lbF), 2000 psi chamber pressure, LO2/LH2 engine has been selected to demonstrate the 20-hour, 500-restart life goal, and a specific impulse of 480 lbF-sec/lbM. The advanced design uses warm oxygen to power a fully integrated turbopump that delivers 4500 psi LO2. The selected engine cycle provides the following advantages over conventional expander cycle designs which use H2 to drive the LH2 and LO2 pumps; higher engine operating pressures and performance in a smaller envelope, lower turbine operating temperatures, elimination of interpropellant seals and purges, and an extended throttling range. The design approach and results of testing to characterize materials for use in high-pressure gaseous oxygen are discussed. Test methods include particle impact testing in a sonic-flow, hot GO2 streams and forced, high-speed friction rubbing testing. Materials are exposed to oxygen and nitrogen atmospheres at pressures of 100, 1000, and 3000 psi to identify separately the gas cooling effects from the metal oxidation effects at the rubbing surface. The selection of candidate design materials is based on an analytical parameter defined as the burn factor. Typical materials tested include carbon steel, stainless steel, nickel, copper, and monel alloys.

Schoenman, L.↗
NASA NTRS1985

Solar thermal propulsion for planetary spacecraft

Previous studies have shown that many desirable planetary exploration missions require large injection delta-V. Solar Thermal Rocket (STR) propulsion, under study for orbit-raising applications may enhance or enable such high-energy missions. The required technology of thermal control for liquid hydrogen propellant is available for the required storage duration. Self-deploying, inflatable solar concentrators are under study. The mass penalty for passive cryogenic thermal control, liquid hydrogen tanks and solar concentrators does not compromise the specific impulse advantage afforded by the STR as compared to chemical propulsion systems. An STR injection module is characterized and performance is evaluated by comparison to electric propulsion options for the Saturn Orbiter Titan Probe (SOTP) and Uranus Flyby Uranus Probe (UFUP) missions.

Sercel, J. C.↗
NASA NTRS1986

NASA electrothermal auxiliary propulsion technology

Electrothermal auxiliary propulsion systems provide high performance options which can have major mission benefits. There are several electrothermal concepts which offer a range of characteristics and benefits. Resistojets are the highest thrust to power option and are currently operational at mission average values of specific impulse, I sub sp approximately 295 sec. Long life, multipropellant resistojets are being developed for the space station, and resistojet technology advancements are being pursued to improve the I sub sp by more than 20 percent for resistojets used in satellite applications. Direct current arcjets have the potential of I sub sp over 400 sec with storable propellants and should provide over 1000 sec with hydrogen. Advanced concepts are being investigated to provide high power density options and possible growth to primary propulsion applications. Broad based experimental and analytical research and technology programs of NASA are summarized and recent significant advances are reviewed.

Stone, J. R.↗
NASA NTRS1985

Two-dimensional kinetics computer program for orbit transfer vehicles

Future Orbit Transfer Vehicles (OTV) presently under consideration require rocket engines delivering a high specific impulse. This high performance can be obtained with large area ratio thrust chambers using oxygen with either hydrogen or hydrocarbon fuels. In the projected nozzles, the combustion products are expanded to low pressure and temperature levels at high Mach number, a domain which has not been experienced with existing rocket engines. In order to assist in the design and evaluation of OTV engines, modifications have recently been incorporated in the Two-Dimensional-Kinetics (TDK) computer program. These modifications include the treatment of shock waves induced by the wall curvature, and shocks which are attached to the wall. The reflection of a shock wave at the flow axis is also treated in an appropriate manner and multiple shock reflections are allowed. A Boundary Layer Module (BLM) has also been incorporated into the TDK computer program in order to provide an automated treatment of the rigorous JANNAF rocket engine performance prediction procedure.

Nickerson, G. R.↗
NASA NTRS1985

Investigation of arcjet nozzle performance

An investigation is performed to examine the feasibility of improving arcjet performance through the use of contoured nozzle designs. The results of preliminary experiments performed on two different nozzle configurations, a 19 degree half-angle cone and a 'bell' shaped nozzle, are described. A unique experimental arcjet apparatus, which effectively tests only the change in nozzle contour on arcjet performance is used in this investigation. The preliminary experimental results indicate approximately an 8 percent improvement in thrust and specific impulse for operation at 15 kW with an ammonia flow rate of 0.20 g/s for a 'bell' shaped nozzle compared to a conical nozzle.

Brophy, J. R.↗
NASA NTRS1985

Resistojet operation with various propellants

A flight-qualified electrothermal thruster demonstrated its adaptability to a variety of propellants. Originally qualified for operation with hydrazine propellant, it was operated with nitrogen, hydrogen, and ammonia propellants, demonstrating 73, 61, and 52 percent overall efficiency with these propellants, respectively, when tested over a wide range of operating conditions. By introducing a preheater to admit hot, rather than cold, propellant inlet gases to the thruster's augmentation heat exchanger, delivered specific impulse closer to theoretical performance limits should be achieved.

Zafran, S.↗
NASA NTRS1986

A xenon ion propulsion module for enhanced spacecraft capability

Design, development and preliminary testing of a two-engine xenon ion propulsion module for early flight evaluation is described. Extensive use is made of flight spare propellant system components and also engineering model 30-cm ion engines and assemblies originally developed for the Solar Electric Propulsion System program. Significant design features include a redundant propellant feed system that incorporates a novel gas pulse assembly for rapid and completely reliable engine startup, a central neutralizer subsystem with dual neutralizers for redundancy, and major ion engine performance improvements resulting in a nearly doubling of the 30-cm engine thrust in addition to operating on xenon rather than mercury propellant. At a module input power of 10.0 kw, maximum thrust and specific impulse are projected to be 0.4 N and 3,500 sec. respectively, for a total module efficiency of 67 percent. Total mass of the xenon ion module, including the propellant tank, is only 70.2 kg. The technical approaches taken towards developing and integrating the subsystems comprising this ion propulsion module are presented and discussed in detail.

Aston, G.↗
NASA NTRS1986

The design and operating characteristics of a 30-kW thermal arcjet engine for space propulsion

Attention is given to the design features of a radiation-cooled, 30-kW thermal arcjet thruster, whose laboratory tests have yielded specific impulses of up to 935 sec at 36-44 percent thrust efficiency, together with a cumulative lifetime of over 400 hours. All materials used, including seals, can sustain operation at temperatures sufficiently elevated to require the radiation of all waste heat. This electric propulsion system is ideally suited to missions such as the Space-Based Radar. A detailed consideration is conducted for the thruster's seals, which are the most critical element of the design.

Pivirotto, T.↗
NASA NTRS1986

SSME thrust chamber modeling with Navier Stokes equations

The capability of predicting two-dimensional, compressible and reacting flow in the combustion chamber and nozzle of the Space Shuttle Main Engine (SSME) is demonstrated. A nonorthogonal body fitted coordinate system has been used to represent the combustor and nozzle geometry. The Navier-Stokes equations are solved for the entire thrust chamber with the k-epsilon turbulence model accounting for compressibility and large pressure gradients effects. Results of the computational test cases reveal all expected features of the transonic nozzle flows including location of sonic line, internal shock and boundary layer build-up. Calculated performance parameters such as thrust, flow rate, and specific impulse are also in reasonble agreement with available data. The results show promising potential of solving full Navier-Stokes equations with heat transfer and two-phase combustion in truly comprehensive modeling of rocket engines.

Przekwas, A. J.↗
NASA NTRS1986

Advanced propulsion for polar-orbiting and coorbiting free flyers

Advanced propulsion systems for polar-orbiting and coorbiting free flyers were investigated. Resistojet, arcjet, ion, magnetoplasmadynamic and chemical-bipropellant nitrogen tetroxide/monomethyl hydrazine (NTO/MMH) propulsion systems were compared to the baseline free-flyer hydrazine (N2H4)-propulsion-system performance. Advanced resistojet, arcjet, ion, and NTO/MMH propulsion systems enable significant propellant-mass savings over the baseline N2H4-propulsion system. Using free-flyer mission requirements from the Langley Research Center Mission-Data Base, detailed propulsion requirements for over thirty free-flyer missions were analyzed. The Polar-Platform trip-time constraints may preclude using a low-thrust electric-propulsion system. Electric propulsion will, however, allow a significant coorbiter propellant-mass reduction. Frequent servicing and nodal-regression effects on the coorbiting free-flyer's orbit increase the required mission velocity change and propellant mass. For many coorbiter missions high-specific-impulse resistojet-, arcjet- and ion-propulsion systems allow substantial life-cycle propellant-mass savings.

Palaszewski, B. A.↗
NASA NTRS1986

PEGASUS - A multi-megawatt nuclear electric propulsion system

A propulsion system (The PEGASUS Drive) consisting of a magnetoplasmadynamic (MPD) thruster driven by a multimegawatt nuclear power system is proposed as the propulsion system for a manned Mars mission. The propulsion system described is based on a mission profile containing a 510-day burn time (for a mission time of approximately 1000 days). Electric propulsion systems have significant advantages over chemical systems, because of high specific impulse, lower propellant requirements, and lower system mass. The thermal power for the PEGASUS Drive is supplied by a boiling liquid-metal fast reactor. The system consists of the reactor, reactor shielding, power conditioning, heat rejection, and MPD thruster subsystems. It is capable of providing a maximum of 8.5 megawatts of electrical power of which 6 megawatts is needed for the thruster system, 1.5 megawatts is available for spacecraft system operations and inflight mission applications, leaving the balance for power system operation.

Coomes, E. P.↗
NASA NTRS1986

NASA electrothermal auxiliary propulsion technology

Electrothermal auxiliary propulsion systems provide high performance options which can have major mission benefits. There are several electrothermal concepts which offer a range of characteristics and benefits. Resistojets are the highest thrust to power option and are currently operational at mission average values of specific impulse, I sub sp approximately 295 sec. Long life, multipropellant resistojets are being developed for the Space Station, and resistojet technology advancements are being pursued to improve the I sub sp by more than 20 percent for resistojets used in satellite applications. Direct current arcjets have the potential of I sub sp over 400 sec with storable propellants and should provide over 1000 sec with hydrogen. Advanced concepts are being investigated to provide high power density options and possible growth to primary propulsion applications. Broad based experimental and analytical research and technology programs of NASA are summarized and recent significant advances are reviewed.

Stone, J. R.↗
NASA NTRS1986

Further shock tunnel studies of scramjet phenomena

Scramjet phenomena were studied using the shock tunnel T3 at the Australian National University. Simple two dimensional models were used with a combination of wall and central injectors. Silane as an additive to hydrogen fuel was studied over a range of temperatures and pressures to evaluate its effect as an ignition aid. The film cooling effect of surface injected hydrogen was measured over a wide range of equivalence. Heat transfer measurements without injection were repeated to confirm previous indications of heating rates lower than simple flat plate predictions for laminar boundary layers in equilibrium flow. The previous results were reproduced and the discrepancies are discussed in terms of the model geometry and departures of the flow from equilibrium. In the thrust producing mode, attempts were made to increase specific impulse with wall injection. Some preliminary tests were also performed on shock induced ignition, to investigate the possibility in flight of injecting fuel upstream of the combustion chamber, where it could mix but not burn.

Morgan, R. G.↗
NASA NTRS1986

A design study of hydrazine and biowaste resistojets

A generalized modeling program was adapted in BASIC on a personal computer to compare the performance of four types of biowaste resistojets and two types of hydrazine augmenters. Analyzed biowaste design types were: (1) an electrically conductive ceramic heater-exchanger of zirconia; (2) a truss heater of platinum in cross flow; (3) an immersed bicoiled tubular heater-exchanger; and (4) a nonexposed, refractory metal, radiant heater in a central cavity within a heat exchanger case. Concepts 2 and 3 are designed to have an efficient, stainless steel outer pressure case. The hydrazine design types are: (5) an immersed bicoil heater exchanger and (6) a nonexposed radiant heater now with a refractory metal case. The ceramic biowaste resistojet has the highest specific impulse growth potential at 2000 K of 192.5 (CO2) and 269 s (H2O). The bicoil produces the highest augmenter temperature of 1994 K for a 2073 K heater giving 317 s at .73 overall efficiency. Detailed temperature profiles of each of the designs are shown. The scaled layout drawings of each are presented with recommended materials and fabrication methods.

Page, R. J.↗
NASA NTRS1986

PEGASUS: A multi-megawatt nuclear electric propulsion system

A propulsion system (PEGASUS) consisting of an electric thruster driven by a multimegawatt nuclear power system is proposed for a manned Mars mission. Magnetoplasmadynamic and mercury-ion thrusters are considered, based on a mission profile containing a 510-day burn time (for a mission time of approximately 1000 days). Both thrusters are capable of meeting the mission parameters. Electric propulsion systems have significant advantages over chemical systems, because of high specific impulse, lower propellant requirements, and lower system mass. The power for the PEGASUS system is supplied by a boiling liquid-metal fast reactor. The power system consists of the reactor, reactor shielding, power conditioning subsystems, and heat rejection subsystems. It is capable of providing a maximum of 8.5 megawatts of electrical power of which 6 megawatts is needed for the thruster system, leaving 1.5 megawatts available for inflight mission applications.

Coomes, Edmund P.↗