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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 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

Synthesis of ceramic powders and surface films from laser heated gases

Two new processes have been developed that are based on laser-heated gases. Both permit unusually precise levels of process control and, thereby, materials having superior properties. The power process yields Si, Si3N4 and SiC powders that are uniform in size, nonagglomerated, small diameter, spherically shaped and high purity. Manufacturing cost analyses show that submicron powders can be made with an energy cost of approximately 2 kWhr/kg and a dollar cost of 2-3.30 $/kg, exclusive of the costs of feed materials. The laser-induced chemical vapor deposition process (LICVD) causes reactant gases to be heated by absorbing IR light from a laser beam that passes parallel to the substrate surface. Laser heating permits independent control of gas and substrate temperatures while operating in a conventional, thermally activated chemical vapor deposition mode. Spin density, hydrogen content, electrical conductivity and mobility gap properties show the LICVD process capable of producing very high quality films.

Haggerty, J. S.↗
NASA NTRS1987

A prototype 4800 bps voice terminal for the mobile satellite experiment

The authors report the final stage of a research project for the NASA mobile satellite experiment (MSAT-X). The work has resulted in two alternative coding techniques which have achieved good communications quality at 4.8 kb/s while tolerating vehicle noise, fading, and a fairly high-error-rate, bursty channel. The authors have constructed a circuit board based on three AT&T 32-bit floating-point DSP32 digital signal processors (DSP) and providing both excess processing power and considerable flexibility for handling different variations of the algorithms. A Motorola 68HC11 microcomputer chip serves as the board controller and interface handler. Mounted on a wirewrapped card, the prototype's circuit footprint amounts to only 225 cm2 and consumes about 13 W of power. A description is given of the architecture, signal processing and communication functions, circuitry, and firmware of the prototype. Issues such as synchronization, framing, error control, and modem interface are also addressed.

Chan, Wai-Yip↗
NASA NTRS1994

Performance evaluation of the Russian SPT-100 thruster at NASA LeRC

Performance measurements of a Russian flight-model SPT-100 thruster were obtained as part of a comprehensive program to evaluate engineering issues pertinent to integration with Western spacecraft. Power processing was provided by a US Government developed laboratory power conditioner. When received the thruster had been subjected to only a few hours of acceptance testing by the manufacturer. Accumulated operating time during this study totalled 148 h and included operation of both cathodes. Cathode flow fraction was controlled both manually and using the flow splitter contained within the supplied xenon flow controller. Data were obtained at current levels ranging from 3 A to 5 A and thruster voltages ranging from 200 V to 300 V. Testing centered on the design power of 1.35 kW with a discharge current of 4.5 A. The effects of facility pressure on thruster operation were examined by varying the pressure via injection of xenon into the vacuum chamber. The facility pressure had a significant effect on thruster performance and stability at the conditions tested. Periods of current instabilities were noted throughout the testing period and became more frequent as testing progressed. Performance during periods of stability agreed with previous data obtained in Russian laboratories.

Sankovic, John M.↗
NASA NTRS2005

Series-Connected Buck Boost Regulators

A series-connected buck boost regulator (SCBBR) is an electronic circuit that bucks a power-supply voltage to a lower regulated value or boosts it to a higher regulated value. The concept of the SCBBR is a generalization of the concept of the SCBR, which was reported in "Series-Connected Boost Regulators" (LEW-15918), NASA Tech Briefs, Vol. 23, No. 7 (July 1997), page 42. Relative to prior DC-voltage-regulator concepts, the SCBBR concept can yield significant reductions in weight and increases in power-conversion efficiency in many applications in which input/output voltage ratios are relatively small and isolation is not required, as solar-array regulation or battery charging with DC-bus regulation. Usually, a DC voltage regulator is designed to include a DC-to-DC converter to reduce its power loss, size, and weight. Advances in components, increases in operating frequencies, and improved circuit topologies have led to continual increases in efficiency and/or decreases in the sizes and weights of DC voltage regulators. The primary source of inefficiency in the DC-to-DC converter portion of a voltage regulator is the conduction loss and, especially at high frequencies, the switching loss. Although improved components and topology can reduce the switching loss, the reduction is limited by the fact that the converter generally switches all the power being regulated. Like the SCBR concept, the SCBBR concept involves a circuit configuration in which only a fraction of the power is switched, so that the switching loss is reduced by an amount that is largely independent of the specific components and circuit topology used. In an SCBBR, the amount of power switched by the DC-to-DC converter is only the amount needed to make up the difference between the input and output bus voltage. The remaining majority of the power passes through the converter without being switched. The weight and power loss of a DC-to-DC converter are determined primarily by the amount of power processed. In the SCBBR, the unswitched majority of the power is passed through with very little power loss, and little if any increase in the sizes of the converter components is needed to enable the components to handle the unswitched power. As a result, the power-conversion efficiency of the regulator can be very high, as shown in the example of Figure 1. A basic SCBBR includes a DC-to-DC converter (see Figure 2). The switches and primary winding of a transformer in the converter is connected across the input bus, while the secondary winding and switches are connected in series with the output bus, so that the output voltage is the sum of the input voltage and the secondary voltage of the converter. In the breadboard SCBBR, the input voltage applied to the primary winding is switched by use of metal oxide/semiconductor field-effect transistors (MOSFETs) in a full bridge circuit; the secondary winding is center-tapped, with two MOSFET switches and diode rectifiers connected in opposed series in each leg. The sets of opposed switches and rectifiers are what enable operation in either a boost or a buck mode. In the boost mode, input voltage and current, and the output voltage and current are all positive; that is, the secondary voltage is added to the input voltage and the net output voltage can be regulated at a value equal or greater than the input voltage. In the buck mode, input voltage is still positive and the current still flows in the same direction in the secondary, but the switches are controlled such that some power flows from the secondary to the primary. The voltage across the secondary and the current into the primary are reversed. The result is that the output voltage is lower than the input voltage, and some power is recirculated from the converter secondary back to the input. Quantitatively, the advantage of an SCBBR is a direct function of the regulation range required. If, for example, a regulation range of 20 percent is required for a 500-W supply, th it suffices to design the DC-to-DC converter in the SCBBR for a power rating of only 100 W. The switching loss and size are much smaller than those of a conventional regulator that must be rated for switching of all 500 W. The reduction in size and the increase in efficiency are not directly proportional to switched-power ratio of 5:1 because the additional switches contribute some conduction loss and the input and output filters must be larger than those typically required for a 100-W converter. Nevertheless, the power loss and the size can be much smaller than those of a 500-W converter.

Birchenough, Arthur G.↗
NASA NTRS2015

Direct Drive Solar-Powered Arcjet Thruster

Electric thrusters typically require a power processing unit (PPU) to convert the spacecraft-provided power to the voltage and current that a thruster needs for operation. NASA Marshall Space Flight Center has initiated fundamental studies on whether an arcjet thruster can be operated directly with the power produced by solar arrays without any additional conversion. Elimination of the PPU significantly reduces system-level complexity of the propulsion system, and lowers developmental cost and risk. The proposed work will aim to refine the proof-of-concept presently being assembled and begin to identify and address technical questions related to power conditioning and noise suppression in the system, and heating of the thruster in long-duration operation. The apparatus proposed for investigation has a target power level of 400 to 1,000 W. The proposed direct-drive arcjet is potentially a highly scalable concept, applicable to spacecraft with up to hundreds of kilowatts and beyond. The design of the arcjet built for this effort was based on previous low power (1 kW class) arcjets.1-3 It has a precision machined 99.95% pure tungsten anode that also serves as the nozzle with a 0.040-in- (1-mm-) diameter, 0.040-in-long constrictor region. An additional anode with a 0.020-in- (0.5-mm-) diameter, 0.020-inlong constrictor region was purchased, but has not yet been used. The cathode is a 0.125-in-diameter tungsten welding electrode doped with lanthum-oxygen; its tip was precision ground to a 308deg angle and terminates in a blunt end. The two electrodes are separated by a boron-nitride insulator that also serves as the propellant manifold; it ends in six small holes which introduce the propellant gas in the diverging section of the nozzle, directly adjacent to the cathode. The electrodes and insulator are housed in a stainless-steel outer body, with a Macor insulator at the mid-plane to provide thermal isolation between the front and back halves of the device. The gas seals were made using Grafoil gaskets. Figure 1(a) shows the assembled thruster; figure 1(b) shows the thruster in the vacuum chamber with electrical and propellant connections.

Polzin, Kurt↗
NASA NTRS2001

Application of Magnetized Target Fusion to High-Energy Space Propulsion

Most fusion propulsion concepts that have been investigated in the past employ some form of inertial or magnetic confinement. Although the prospective performance of these concepts is excellent, the fusion processes on which these concepts are based still require considerable development before they can be seriously considered for actual applications. Furthermore, these processes are encumbered by the need for sophisticated plasma and power handling systems that are generally quite inefficient and have historically resulted in large, massive spacecraft designs. Here we present a comparatively new approach, Magnetized Target Fusion (MTF), which offers a nearer-term avenue for realizing the tremendous performance benefits of fusion propulsion'. The key advantage of MTF is its less demanding requirements for driver energy and power processing. Additional features include: 1) very low system masses and volumes, 2) high gain and relatively low waste heat, 3) substantial utilization of energy from product neutrons, 4) efficient, low peak-power drivers based on existing pulsed power technology, and 5) very high Isp, specific power and thrust. MTF overcomes many of the problems associated with traditional fusion techniques, thus making it particularly attractive for space applications. Isp greater than 50,000 seconds and specific powers greater than 50 kilowatts/kilogram appear feasible using relatively near-term pulse power and plasma gun technology.

Thio, Y. C. F.↗
NASA NTRS1977

New process produces high-power Schottky diodes

Processing procedure using low-temperature platinum silicide, results in successful high-yield fabrication of large-area mesa-geometry Schottky diodes, with reverse breakdown voltages as high as 150 volts and leak currents less than 5 milliamps at 212 F.

Cordes, L. F.↗
NASA NTRSDate not supplied

Strategy for Developing Technologies for Megawatt-class Nuclear Electric Propulsion Systems

In late fiscal year 2020, the Space Nuclear Propulsion (SNP) project began the process of formulating an investment strategy to support development of the technologies required for a high-power (megawatt-class) nuclear electric propulsion (NEP) system capable of performing human-scale missions. This activity was initiated concurrent with several high-level studies and assessments were either under way or had just concluded. Studies of human-scale Mars missions have been performed several times over the past two decades. One of the most recent studies examined opposition-class human Mars missions to occur in the late 2030s timeframe [1,2]. The mission architecture assumed a hybrid NEP/chem-propelled vehicle that used a high specific impulse (Isp) NEP-system and a liquid oxygen (LOx)-liquid methane high thrust chemical stage (two 110 kN (25 klbf) thrust, 365 s Isp engines) for maneuvers performed to enter and exit gravity wells. Trajectory analyses performed in this study showed that such a mission could be performed with 2-4 MWe directed into the electric propulsion system (operating for 20,000+ hours), with the large range representing different opposition-class Mars mission opportunities and permutations on the trajectory design, concept of operations, and technology choices. In 2020, the NASA Engineering and Safety Center (NESC) performed a study to evaluate the maturity of the different technologies required for nuclear propulsion systems [3]. The executive summary of this report provided the following top-level conclusions: • “The majority of critical technologies for… NEP/Chem… systems are relatively immature” • “TRLs [technology readiness levels] in the literature are often overestimated” • “The majority of critical technologies… for NEP/Chem… systems are at a relatively high level of advancement degree of difficulty (AD2 > 4) for maturation, requiring a dual development approach” • “The proper assessment of baseline TRL and AD2 values and the estimation of requirements and resources required for advancement have been consistent issues for NEP,” • “Non-advocate reviews should occur at the start of a technology program and at all key milestones.” In 2021, the National Academies of Science, Engineering, and Medicine (NASEM) issued a separate report [4] identifying the “primary technical and programmatic challenges, merits, and risks for maturing space nuclear propulsion technologies of interest to a future human Mars exploration mission.” That work contained several important findings, including: • “Developing a MWe-class NEP system for the baseline mission would require increasing power by orders of magnitude relative to NEP system flight- or ground-based technology demonstrations completed to date.” • “Subscale in-space flight testing of NEP systems cannot address many of the risks and potential failure modes associated with the baseline mission NEP system. With sufficient M&S [modeling & simulation] and ground testing, including modular subsystem tests at full scale and power, flight qualification requirements can be met by the cargo missions that will precede the first crewed mission to Mars. Fully integrated ground testing may not be required.” • “As a result of low and intermittent investment over the past several decades, it is unclear if even an aggressive program would be able to develop an NEP system capable of executing the baseline mission in 2039.” These efforts motivated the SNP project to investigate the technologies available for a megawatt-class high power nuclear electric propulsion system. That system is illustrated schematically in Figure 1 and is comprised of five separate top-level critical technology elements (CTEs). 1. Nuclear Reactor – Thermal power source for the system, utilizing high-assay low enriched uranium (HALEU) as the nuclear fuel. Reactor radiation shielding is also included in this CTE. 2. Power Conversion – Operates as a thermodynamic cycle, accepting nuclear reactor thermal power as an input and converting it to mechanical power. 3. Power Management and Distribution (PMAD) – Accepts as an input mechanical power from the power conversion system, which is used to generate electrical power. The PMAD system also distributes the generated electrical power to all other parts of the spacecraft, including the high-power EP system. The PMAD system may also perform duties such as isolation, fault detection, and power transformation/rectification for different spacecraft systems, including the thrusters. 4. Electric Propulsion (EP) – Accepts as an input electrical power, which is used to accelerate a propellant to high speeds to produce thrust. This system includes the power processing unit (PPU), which converts the power it receives to the correct current and voltage required by the thrusters, and the propellant storage and feed systems, which contain and meter the flow of propellant to the thrusters. 5. Thermal Management (Radiators/Heat Rejection) – The cold side of the thermodynamic power conversion cycle, accepts thermal power from the power conversion system and radiatively rejects that heat to space. In this paper, we describe the SNP project formulation and investment strategy that aims to accomplish the research and development required to advance the technology readiness for each CTE. The strategy relies heavily upon experimental testing supported by modeling and simulation to yield realistic assessments of the technologies, which in turn will be used to inform future NEP system-level design decisions and any potential technology downselects.

Kurt A Polzin↗
NASA NTRS2015

Low-Power Operation and Plasma Characterization of a Qualification Model SPT-140 Hall Thruster for NASA Science Missions

Hall thruster systems based on commercial product lines can potentially lead to lower cost electric propulsion (EP) systems for deep space science missions. A 4.5-kW SPT-140 Hall thruster presently under qualification testing by SSL leverages the substantial heritage of the SPT-100 being flown on Russian and US commercial satellites. The Jet Propulsion Laboratory is exploring the use of commercial EP systems, including the SPT-140, for deep space science missions, and initiated a program to evaluate the SPT-140 in the areas of low power operation and thruster operating life. A qualification model SPT-140 designated QM002 was evaluated for operation and plasma properties along channel centerline, from 4.5 kW to 0.8 kW. Additional testing was performed on a development model SPT-140 designated DM4 to evaluate operation with a Moog proportional flow control valve (PFCV). The PFCV was commanded by an SSL engineering model PPU-140 Power Processing Unit (PPU). Performance measurements on QM002 at 0.8 kW discharge power were 50 mN of thrust at a total specific impulse of 1250 s, a total thruster efficiency of 0.38, and discharge current oscillations of under 3% of the mean current. Steady-state operation at 0.8 kW was demonstrated during a 27 h firing. The SPT-140 DM4 was operated in closed-loop control of the discharge current with the PFCV and PPU over discharge power levels of 0.8-4.5 kW. QM002 and DM4 test data indicate that the SPT-140 design is a viable candidate for NASA missions requiring power throttling down to low thruster input power.

Garner, Charles E.↗
NASA NTRS1972

Development of a multikilowatt ion thruster power processor

A feasibility study was made of the application of silicon-controlled, rectifier series, resonant inverter, power conditioning technology to electric propulsion power processing operating from a 200 to 400 Vdc solar array bus. A power system block diagram was generated to meet the electrical requirements of a 20 CM hollow cathode, mercury bombardment, ion engine. The SCR series resonant inverter was developed as a primary means of power switching and conversion, and the analog signal-to-discrete-time-interval converter control system was applied to achieve good regulation. A complete breadboard was designed, fabricated, and tested with a resistive load bank, and critical power processor areas relating to efficiency, weight, and part count were identified.

Schoenfeld, A. D.↗
NASA NTRS1981

Goals of thermionic program for space power

The thermionic and Brayton reactor concepts were compared for application to space power. For a turbine inlet temperature of 15000 K the Brayton powerplant weighted 5 to 40% less than the thermionic concept. The out of core concept separates the thermionic converters from their reactor. Technical risks are diminished by: (1) moving the insolator out of the reactor; (2) allowing a higher thermal flux for the thermionic converters than is required of the reactor fuel; and (3) eliminating fuel swelling's threat against lifetime of the thermionic converters. Overall performance can be improved by including power processing in system optimization for design and technology on more efficient, higher temperature power processors. The thermionic reactors will be larger than those for competitive systems with higher conversion efficiency and lower reactor operating temperatures. It is concluded that although the effect of reactor size on shield weight will be modest for unmanned spacecraft, the penalty in shield weight will be large for manned or man-tended spacecraft.

English, R. E.↗
NASA NTRS1984

Space Power Management and Distribution Status and Trends

An overview of space power management and distribution (PMAD) is provided which encompasses historical and current technology trends. The PMAD components discussed include power source control, energy storage control, and load power processing electronic equipment. The status of distribution equipment comprised of rotary joints and power switchgear is evaluated based on power level trends in the public, military, and commercial sectors. Component level technology thrusts, as driven by perceived system level trends, are compared to technology status of piece-parts such as power semiconductors, capacitors, and magnetics to determine critical barriers.

Reppucci, G. M.↗
NASA NTRS2010

NASA's Evolutionary Xenon Thruster (NEXT) Prototype Model 1R (PM1R) Ion Thruster and Propellant Management System Wear Test Results

The results of the NEXT wear test are presented. This test was conducted with a 36-cm ion engine (designated PM1R) and an engineering model propellant management system. The thruster operated with beam extraction for a total of 1680 hr and processed 30.5 kg of xenon during the wear test, which included performance testing and some operation with an engineering model power processing unit. A total of 1312 hr was accumulated at full power, 277 hr at low power, and the remainder was at intermediate throttle levels. Overall ion engine performance, which includes thrust, thruster input power, specific impulse, and thrust efficiency, was steady with no indications of performance degradation. The propellant management system performed without incident during the wear test. The ion engine and propellant management system were also inspected following the test with no indication of anomalous hardware degradation from operation.

VanNoord, Jonathan L.↗
NASA NTRS1977

Mass study for modular approaches to a solar electric propulsion module

The propulsion module comprises six to eight 30-cm thruster and power processing units, a mercury propellant storage and distribution system, a solar array ranging in power from 18 to 25 kW, and the thermal and structure systems required to support the thrust and power subsystems. Launch and on-orbit configurations are presented for both modular approaches. The propulsion module satisfies the thermal design requirements of a multimission set including: Mercury, Saturn, and Jupiter orbiters, a 1-AU solar observatory, and comet and asteroid rendezvous. A detailed mass breakdown and a mass equation relating the total mass to the number of thrusters and solar array power requirement is given for both approaches.

Sharp, G. R.↗
NASA NTRS1986

Scaling of data communications for an advanced supercomputer network

The goal of NASA's Numerical Aerodynamic Simulation (NAS) Program is to provide a powerful computational environment for advanced research and development in aeronautics and related disciplines. The present NAS system consists of a Cray 2 supercomputer connected by a data network to a large mass storage system, to sophisticated local graphics workstations and by remote communication to researchers throughout the United States. The program plan is to continue acquiring the most powerful supercomputers as they become available. The implications of a projected 20-fold increase in processing power on the data communications requirements are described.

Levin, E.↗
NASA NTRS1987

Spacecraft and mission design for the SpaceNuclear PowerSystem Reference Mission

The design and performance of a spacecraft employing arcjet nuclear electric propulsion, suitable for use in the Space Nuclear Power System Reference Mission, are outlined. The vehicle design is based on a 92 kW ammonia arcjet system operating at an I(sp) of 1050 s and an efficiency of 45 percent. The arcjet/gimbal system, power processing unit, and propellant feed-system are described. A 100 kW(e) space nuclear power system is assumed and the spacecraft mass is baselined at 5250 kg excluding the propellant, propellent feed system, and integrated chemical boost engine. A radiation/arcjet efflux diagnostics package is included in the performance analysis. Three mission scenarios are described and are capable of demonstrating the full capability of the space nuclear power source. The missions considered include power system deployment to possible SDI platform orbits and a spacecraft storage mission to an orbit of three times geosynchronous (GEO) with return to GEO corresponding to Delta V's between 7400 m/s, and 7900 m/s. This spacecraft meets the Reference Mission constraint of low developmental risk and is scaleable to power levels projected for future space platforms.

Deininger, William D.↗
NASA NTRS1990

A feasibility study and mission analysis for the Hybrid Plume Plasma Rocket

The Hybrid Plume Plasma Rocket (HPPR) is a high power electric propulsion concept which is being developed at the MIT Plasma Fusion Center. This paper presents a theoretical overview of the concept as well as the results and conclusions of an independent study which has been conducted to identify and categorize those technologies which require significant development before the HPPR can be considered a viable electric propulsion device. It has been determined that the technologies which require the most development are high power radio-frequency and microwave generation for space applications and the associated power processing units, low mass superconducting magnets, a reliable, long duration, multi-megawatt space nuclear power source, and long term storage of liquid hydrogen propellant. In addition to this, a mission analysis of a one-way transfer from low earth orbit (LEO) to Mars indicates that a constant acceleration thrust profile, which can be obtained using the HPPR, results in faster trip times and greater payload capacities than those afforded by more conventional constant thrust profiles.

Sullivan, Daniel J.↗
NASA NTRS2001

Ion Propulsion Technology Programs at NASA Glenn Research Center

As lead center for the agency in electric and ion propulsion, the NASA Glenn Research Center (GRC) is pursuing technology development in ion propulsion for a range of mission applications. The program goal is to develop key technologies for advanced NSTAR-derivative high-power ion propulsion, lightweight low power high-performance ion propulsion, 'micro' ion propulsion, and engine and component technologies for high-power electric propulsion for very ambitious missions. Products include: (1) a 5 kW, 400 kg throughput ion thruster and power processing technology; (2) extremely-lightweight high-efficiency sub-kilowatt ion thruster and power processor; (3) a 1-25 W high-specific impulse ion engine; and (4) engine and component technologies for high-power (30 kW class) ion and Hall engines. Identified applications include outer planetary science missions such as Europa orbiter/lander, Comet Nucleus Sample Return mission, Titan Explorer, Neptune/Triton, Pluto-Kuiper Belt Objects Mission, various second generation interplanetary Micro spacecraft, and the Interstellar Probe Mission. Additional information is contained in the original extended abstract.

Patterson, M. J.↗