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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 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS1982

Low concentration ratio solar array for low Earth orbit multi-100 kW application

An ongoing preliminary design effort directed toward a low-concentration-ratio photovoltaic array system based on 1984 technology and capable of delivering multi-hundred kilowatts (300 kW to 1000 kW range) in low earth orbit is described. The array system consists of two or more array modules each capable of delivering between 80 kW to 172 kW using silicon solar cells or gallium arsenide solar cells respectively. The array module deployed area is 1320 square meters and consists of 4356 pryamidal concentrator elements. The module, when stowed in the Space Shuttle's payload bay, has a stowage volume of a cube with 3.24 meters on a side. The concentrator elements are sized for a geometric concentration ratio (GCR) of six with an aperture area of 0.5 meters x 0.5 meters. The structural analysis and design trades leading to the baseline design are discussed. The configuration, as well as optical, thermal and electrical performance analyses that support the design and overall performance estimates for the array are described.

Nalbandian, S. J.↗
NASA NTRS1983

Low concentration ratio solar array for low Earth orbit multi-100 kW application. Volume 1: Design, analysis and development tests

A preliminary design effort directed toward a low concentration ratio photovoltaic array system capable of delivering multihundred kilowatts (300 kW to 1000 kW range) in low earth orbit is described. The array system consists of two or more array modules each capable of delivering between 113 kW to 175 kW using silicon solar cells or gallium arsenide solar cells, respectively. The array module deployed area is 1320 square meters and consists of 4356 pyramidal concentrator elements. The module, when stowed in the Space Shuttle's payload bay, has a stowage volume of a cube with 3.24 meters on a side. The concentrator elements are sized for a geometric concentration ratio (GCR) of six with an aperture area of .25 sq. m. The structural analysis and design trades leading to the baseline design are discussed. It describes the configuration, as well as optical, thermal and electrical performance analyses that support the design and overall performance estimates for the array are described.

Source record↗
NASA NTRS1992

The 10 kW power electronics for hydrogen arcjets

A combination of emerging mission considerations such as 'launch on schedule', resource limitations, and the development of higher power spacecraft busses has resulted in renewed interest in high power hydrogen arcjet systems with specific impulses greater than 1000 s for Earth-space orbit transfer and maneuver applications. Solar electric propulsion systems with about 10 kW of power appear to offer payload benefits at acceptable trip times. This work outlines the design and development of 10 kW hydrogen arcjet power electronics and results of arcjet integration testing. The power electronics incorporated a full bridge switching topology similar to that employed in state of the art 5 kW power electronics, and the output filter included an output current averaging inductor with an integral pulse generation winding for arcjet ignition. Phase shifted, pulse width modulation with current mode control was used to regulate the current delivered to arcjet, and a low inductance power stage minimized switching transients. Hybrid power Metal Oxide Semiconductor Field Effect Transistors were used to minimize conduction losses. Switching losses were minimized using a fast response, optically isolated, totem-pole gate drive circuit. The input bus voltage for the unit was 150 V, with a maximum output voltage of 225 V. The switching frequency of 20 kHz was a compromise between mass savings and higher efficiency. Power conversion efficiencies in excess of 0.94 were demonstrated, along with steady state load current regulation of 1 percent. The power electronics were successfully integrated with a 10 kW laboratory hydrogen arcjet, and reliable, nondestructive starts and transitions to steady state operation were demonstrated. The estimated specific mass for a flight packaged unit was 2 kg/kW.

Hamley, John A.↗
NASA NTRS1997

A 50-kW Module Power Station of Directly Solar-Pumped Iodine Laser

The conceptual design of a 50 kW Directly Solar-Pumped Iodine Laser (DSPIL) module was developed for a space-based power station which transmits its coherent-beam power to users such as the moon, Martian rovers, or other satellites with large (greater than 25 kW) electric power requirements. Integration of multiple modules would provide an amount of power that exceeds the power of a single module by combining and directing the coherent beams to the user's receiver. The model developed for the DSPIL system conservatively predicts the laser output power (50 kW) that appears much less than the laser output (93 kW) obtained from the gain volume ratio extrapolation of experimental data. The difference in laser outputs may be attributed to reflector configurations adopted in both design and experiment. Even though the photon absorption by multiple reflections in experimental cavity setup was more efficient, the maximum secondary absorption amounts to be only 24.7 percent of the primary. However, the gain volume ratio shows 86 percent more power output than theoretical estimation that is roughly 60 percent more than the contribution by the secondary absorption. Such a difference indicates that the theoretical model adopted in the study underestimates the overall performance of the DSPIL. This fact may tolerate more flexible and radical selection of design parameters than used in this design study. The design achieves an overall specific power of approximately 5 W/kg and total mass of 10 metric tons.

Choi, S. H.↗
NASA NTRS2010

The Deep Space Network's X/X/Ka Feed: Modifications for 100 kW CW Uplink Operation

The Deep Space Network, which provides communication services for NASA's robotic missions, consists of a number of 34m beam waveguide antennas and conventional 70m dual-reflector antennas located around the globe, [1]. The 34m beam waveguide antennas employ a three-band feed covering the deep space uplink band near 7.2 GHz, and downlink bands at 8.45 and 32 GHz. Simultaneous uplink commanding at 25 kW CW and ultra low noise reception in both bands is supported along with monopulse tracking at 32 GHz, [2]. An existing uplink capability of 25 kW is also available on the 70m antennas using a more conventional X/X diplexing feed. In order to provide an equivalent uplink capability with the 34m antennas the X/X/Ka feed is currently being modified for 100 kW CW operation, [3]. Here we will discuss both the existing feed and the 100 kW modifications which are underway.

Feed Uplink↗
NASA NTRS2015

Performance and Facility Background Pressure Characterization Tests of NASAs 12.5-kW Hall Effect Rocket with Magnetic Shielding Thruster

NASA's Space Technology Mission Directorate (STMD) Solar Electric Propulsion Technology Demonstration Mission (SEP/TDM) project is funding the development of a 12.5-kW Hall thruster system to support future NASA missions. The thruster designated Hall Effect Rocket with Magnetic Shielding (HERMeS) is a 12.5-kW Hall thruster with magnetic shielding incorporating a centrally mounted cathode. HERMeS was designed and modeled by a NASA GRC and JPL team and was fabricated and tested in vacuum facility 5 (VF5) at NASA GRC. Tests at NASA GRC were performed with the Technology Development Unit 1 (TDU1) thruster. TDU1's magnetic shielding topology was confirmed by measurement of anode potential and low electron temperature along the discharge chamber walls. Thermal characterization tests indicated that during full power thruster operation at peak magnetic field strength, the various thruster component temperatures were below prescribed maximum allowable limits. Performance characterization tests demonstrated the thruster's wide throttling range and found that the thruster can achieve a peak thruster efficiency of 63% at 12.5 kW 500 V and can attain a specific impulse of 3,000 s at 12.5 kW and a discharge voltage of 800 V. Facility background pressure variation tests revealed that the performance, operational characteristics, and magnetic shielding effectiveness of the TDU1 design were mostly insensitive to increases in background pressure.

Performance↗
NASA NTRS2017

150 kW Class Solar Electric Propulsion Spacecraft Power Architecture Model

The National Aeronautics and Space Administration (NASA) Solar Electric Propulsion Technology Demonstration Mission (SEP TDM), in conjunction with PC Krause and Associates, has created a Simulink-based power architecture model for a 50 kilo-Watt (kW) solar electric propulsion system. NASA has extended this model to investigate 150 kW solar electric propulsion systems. Increasing the power capability to 150 kW is an intermediate step to the anticipated power requirements for Mars and other deep space applications. The high-power solar electric propulsion capability has been identified as a critical part of NASA’s future beyond-low-Earth-orbit for human-crewed exploration missions. This paper presents four versions of a 150 kW architecture, simulation results, and a discussion of their merits.

Power Control↗
NASA NTRS1979

The 25 kW power module evolution study. Part 3: Conceptual design for power module evolution. Volume 6: WBS and dictionary

Program elements of the power module (PM) system, are identified, structured, and defined according to the planned work breakdown structure. Efforts required to design, develop, manufacture, test, checkout, launch and operate a protoflight assembled 25 kW, 50 kW and 100 kW PM include the preparation and delivery of related software, government furnished equipment, space support equipment, ground support equipment, launch site verification software, orbital verification software, and all related data items.

Source record↗
NASA NTRS1981

Preliminary design development of 100 KW rotary power transfer device

Contactless power transfer devices for transferring electrical power across a rotating spacecraft interface were studied. A power level of 100 KW was of primary interest and the study was limited to alternating current devices. Rotary transformers and rotary capacitors together with the required dc to ac power conditioning electronics were examined. Microwave devices were addressed. The rotary transformer with resonant circuit power conditioning was selected as the most feasible approach. The rotary capacitor would be larger while microwave devices would be less efficient. A design analysis was made of a 100 KW, 20 kHz power transfer device consisting of a rotary transformer, power conditioning electronics, drive mechanism and heat rejection system. The size, weight and efficiency of the device were determined. The characteristics of a baseline slip ring were presented. Aspects of testing the 100 KW power transfer device were examined. The power transfer device is a feasible concept which can be implemented using presently available technologies.

Weinberger, S. M.↗
NASA NTRS1984

Proposed system design for a 20 kW pulsed electrothermal thruster

A conceptual design is presented for a Pulsed Electrothermal (PET) propulsion system for the Air Force Space Based Radar satellite, which has a mass of 7000 kg. The proposed system boosts the SBR satellite from 150 n.m. to 600 n.m. with a 4 deg plane change, for a total mission Delta v of 1 km/sec. Satellite power available is 50 kW, and 45 kW are used to drive two water-injected 20 kW PET thrusters, delivering 5.6 N thrust to the SBR at 1000 seconds specific impulse. The predicted mission trip time is 15 days. The proposed system consumes 850 kg of water propellant, stored in a central tank and injected with pressurized helium. Component mass estimates based on space-qualified hardware are presented for the propellant handling, power conditioning and thruster subsystems. The estimated total mass is 400 kg and the propulsion system specific mass is alpha = 10 kg/kW. The proposed system efficiency of 0.62 at 1000 seconds specific impulse is supported by experimental performance measurements.

Burton, R. L.↗
NASA NTRS1987

Operating characteristics of a 10 kW xenon ion propulsion module

Performance testing of a two-engine functional model xenon ion propulsion module is described. Use of highly modified J-series 30 cm ion engines reconfigured for xenon propellant, a computer controlled operating system, and precise flow control system are shown to result in very reliable ion module operation at high input power levels. Ion engine operation at a nominal 4.0 ampere beam current and 30.0 volt discharge gives a specific impulse of 3310 sec, a total engine efficiency of 64.3 percent, and a thrust-to-power ratio of 39.5 mN/kW at an input power level of 5.10 kW. These modified J series ion engines are shown to be capable of throttling over an 8:1 range from a power level of 5.48 kW at 3285 sec to a power level of 0.70 kW at 1856 sec. In addition, complete ion engine performance mapping of important system level parameters such as thrust, specific impulse, efficiency and thrust-to-power ratio are presented.

Aston, Graeme↗
NASA NTRS1990

The 30-kW ammonia arcjet technology

The technical results are summarized of a 30 kW class ammonia propellant arcjet technology program. Evaluation of previous arcjet thruster performance, including materials analysis of used thruster components, led to the design of an arcjet with improved performance and thermal characteristics. Tests of the new engine demonstrated that engine performance is relatively insensitive to cathode tip geometry. Other data suggested a maximum sustainable arc length for a given thruster configuration, beyond which the arc may reconfigure in a destructive manner. A flow controller calibration error was identified. This error caused previously reported values of specific impulse and thrust efficiency to be 20 percent higher than the real values. Corrected arcjet performance data are given. Duration tests of 413 and 252 hours, and several tests 100 hours in duration, were performed. The cathode tip erosion rate increased with increasing arc current. Elimination of power source ripple did not affect cathode tip whisker growth. Results of arcjet modeling, diagnostic development and mission analyses are also discussed. The 30 kW ammonia arcjet may now be considered ready for development for a flight demonstration, but widespread application of 30 kW class arcjet will require improved efficiency and lifetime.

Deininger, W. D.↗
NASA NTRS1991

Preliminary performance and life evaluations of a 2-kW arcjet

The first results of a program to expand the operational envelope of low-power arcjets to higher specific impulse and power levels are presented. The performance of a kW-class laboratory model arcjet thruster was characterized at three mass flow rates of a 2:1 mixture of hydrogen and nitrogen at power levels ranging from 1.0 to 2.0 kW. This same thruster was then operated for a total of 300 h at a specific impulse and power level of 550 s and 2.0 kW, respectively, in three continuous 100-h sessions. Thruster operation during the three test segments was stable, and no measurable performance degradation was observed during the test series. Substantial cathode erosion was observed during an inspection following the second 100-h test segment. Most notable was the migration of material from the center of the cathode tip to a ring around a large crater. The anode sustained no significant damage during the endurance test segments. Some difficulty was encountered during start-up after disassembly and inspection following the second 100-h test segment, which caused constrictor erosion. This resulted in a reduced flow restriction and arc chamber pressure, which in turn caused a reduction in the arc impedance.

Morren, W. Earl↗
NASA NTRS1991

Preliminary performance and life evaluation of a 2-kW arcjet

The first results of a program to expand the operational envelope of low-power arcjets to higher specific impulse and power levels are presented. The performance of a kW-class laboratory model arcjet thruster was characterized at three mass flow rates of a 2:1 mixture of hydrogen and nitrogen at power levels ranging from 1.0 to 2.0 kW. This same thruster was then operated for a total of 300 h at a specific impulse and power level of 550 s and 2.0 kW, respectively, in three continuous 100-h sessions. Thruster operation during the three test segments was stable, and no measurable performance degradation was observed during the test series. Substantial cathode erosion was observed during an inspection following the second 100-h test segment. Most notable was the migration of material from the center of the cathode tip to a ring around a large crater. The anode sustained no significant damage during the endurance test segments. Some difficulty was encountered during start-up after disassembly and inspection following the second 100-h test segment, which caused constrictor erosion. This resulted in a reduced flow restriction and arc chanber pressure, which in turn caused a reduction in the arc impedance.

Morren, W. E.↗
NASA NTRS1997

Development of Power Electronics for a 0.2kW-Class Ion Thruster

Applications that might benefit from low power ion propulsion systems include Earth-orbit magnetospheric mapping satellite constellations, low Earth-orbit satellites, geosynchronous Earth-orbit satellite north-south stationkeeping, and asteroid orbiters. These spacecraft are likely to have masses on the order of 50 to 500 kg with up to 0.5 kW of electrical power available. A power processing unit for a 0.2 kW-class ion thruster is currently under development for these applications. The first step in this effort is the development and testing of a 0.24 kW beam power supply. The design incorporates a 20 kHz full bridge topology with multiple secondaries connected in series to obtain outputs of up to 1200 V(sub DC). A current-mode control pulse width modulation circuit built using discrete components was selected for this application. An input voltage of 28 +/- 4 V(sub DC) was assumed, since the small spacecraft for which this system is targeted are anticipated to have unregulated low voltage busses. Efficiencies in excess of 91 percent were obtained at maximum output power. The total mass of the breadboard was less than 1.0 kg and the component mass was 0.53 kg. It is anticipated that a complete flight power processor could weigh about 2.0 kg.

Pinero, Luis R.↗
NASA NTRS1999

Preliminary Evaluation of a 10 kW Hall Thruster

A 10 kW Hall thruster was characterized over a range of discharge voltages from 300-500 V and a range of discharge currents from 15-23 A. This corresponds to power levels from a low of 4.6 kW to a high of 10.7 kW. Over this range of discharge powers, thrust varied from 278 mN to 524 mN, specific impulse ranged from 1644 to 2392 seconds, and efficiency peaked at approximately 59%. A continuous 40 hour test was also undertaken in an attempt to gain insight with regard to long term operation of the engine. For this portion of the testing the thruster was operated at a discharge voltage of 500 V and a discharge current of 20 A. Steady-state temperatures were achieved after 3-5 hrs and very little variation in performance was detected.

Jankovsky, Robert S.↗
NASA NTRS2001

1000 Hours of Testing Completed on 10-kW Hall Thruster

Between the months of April and August 2000, a 10-kW Hall effect thruster, designated T- 220, was subjected to a 1000-hr life test evaluation. Hall effect thrusters are propulsion devices that electrostatically accelerate xenon ions to produce thrust. Hall effect propulsion has been in development for many years, and low-power devices (1.35 kW) have been used in space for satellite orbit maintenance. The T-220, shown in the photo, produces sufficient thrust to enable efficient orbital transfers, saving hundreds of kilograms in propellant over conventional chemical propulsion systems. This test is the longest operation ever achieved on a high-power Hall thruster (greater than 4.5 kW) and is a key milestone leading to the use of this technology for future NASA, commercial, and military missions.

Mason, Lee S.↗
NASA NTRS2016

Structural Design Considerations for a 50 kW-Class Solar Array for NASA's Asteroid Redirect Mission

NASA is planning an Asteroid Redirect Mission (ARM) to take place in the 2020s. To enable this multi-year mission, a 40 kW class solar electric propulsion (SEP) system powered by an advanced 50 kW class solar array will be required. Powered by the SEP module (SEPM), the ARM vehicle will travel to a large near-Earth asteroid, descend to its surface, capture a multi-metric ton (t) asteroid boulder, ascend from the surface and return to the Earth-moon system to ultimately place the ARM vehicle and its captured asteroid boulder into a stable distant orbit. During the years that follow, astronauts flying in the Orion multipurpose crew vehicle (MPCV) will dock with the ARM vehicle and conduct extra-vehicular activity (EVA) operations to explore and sample the asteroid boulder. This paper will review the top structural design considerations to successfully implement this 50 kW class solar array that must meet unprecedented performance levels. These considerations include beyond state-of-the-art metrics for specific mass, specific volume, deployed area, deployed solar array wing (SAW) keep in zone (KIZ), deployed strength and deployed frequency. Analytical and design results are presented that support definition of stowed KIZ and launch restraint interface definition. An offset boom is defined to meet the deployed SAW KIZ. The resulting parametric impact of the offset boom length on spacecraft moment of inertias and deployed SAW quasistatic and dynamic load cases are also presented. Load cases include ARM spacecraft thruster plume impingement, asteroid surface operations and Orion docking operations which drive the required SAW deployed strength and damping. The authors conclude that to support NASA's ARM power needs, an advanced SAW is required with mass performance better than 125 W/kg, stowed volume better than 40 kW/cu m, a deployed area of 200 sq m (100 sq m for each of two SAWs), a deployed SAW offset distance of nominally 3-4 m, a deployed SAW quasistatic strength of nominally 0.1 g in any direction, a deployed loading displacement under 2 m, a deployed fundamental frequency above 0.1 Hz and deployed damping of at least 1%. These parameters must be met on top of challenging mission environments and ground testing requirements unique to the ARM project.

solar arrays↗