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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 379 records · Page 21
All sourcesNASA NTRSDOE OSTI
NASA NTRS2002

Laboratory Model 50 kW Hall Thruster

A 0.46 meter diameter Hall thruster was fabricated and performance tested at powers up to 72 kilowatts. Thrusts up to 2.9 Newtons were measured. Discharge specific impulses ranged from 1750 to 3250 seconds with discharge efficiencies between 46 and 65 percent. Overall specific impulses ranged from 1550 to 3050 seconds with overall efficiencies between 40 and 57 percent. Performance data indicated significant fraction of multiple-charged ions during operation at elevated power levels. Cathode mass flow rate was shown to be a significant parameter with regard to thruster efficiency.

Manzella, David↗
NASA NTRS2015

Magnesium Hall Thruster

This Phase II project is developing a magnesium (Mg) Hall effect thruster system that would open the door for in situ resource utilization (ISRU)-based solar system exploration. Magnesium is light and easy to ionize. For a Mars- Earth transfer, the propellant mass savings with respect to a xenon Hall effect thruster (HET) system are enormous. Magnesium also can be combusted in a rocket with carbon dioxide (CO2) or water (H2O), enabling a multimode propulsion system with propellant sharing and ISRU. In the near term, CO2 and H2O would be collected in situ on Mars or the moon. In the far term, Mg itself would be collected from Martian and lunar regolith. In Phase I, an integrated, medium-power (1- to 3-kW) Mg HET system was developed and tested. Controlled, steady operation at constant voltage and power was demonstrated. Preliminary measurements indicate a specific impulse (Isp) greater than 4,000 s was achieved at a discharge potential of 400 V. The feasibility of delivering fluidized Mg powder to a medium- or high-power thruster also was demonstrated. Phase II of the project evaluated the performance of an integrated, highpower Mg Hall thruster system in a relevant space environment. Researchers improved the medium power thruster system and characterized it in detail. Researchers also designed and built a high-power (8- to 20-kW) Mg HET. A fluidized powder feed system supporting the high-power thruster was built and delivered to Busek Company, Inc.

Szabo, James J.↗
NASA NTRS2003

High-Energy Two-Stage Pulsed Plasma Thruster

A high-energy (28 kJ per pulse) two-stage pulsed plasma thruster (MSFC PPT-1) has been constructed and tested. The motivation of this project is to develop a high power (approximately 500 kW), high specific impulse (approximately 10000 s), highly efficient (greater than 50%) thruster for use as primary propulsion in a high power nuclear electric propulsion system. PPT-1 was designed to overcome four negative characteristics which have detracted from the utility of pulsed plasma thrusters: poor electrical efficiency, poor propellant utilization efficiency, electrode erosion, and reliability issues associated with the use of high speed gas valves and high current switches. Traditional PPTs have been plagued with poor efficiency because they have not been operated in a plasma regime that fully exploits the potential benefits of pulsed plasma acceleration by electromagnetic forces. PPTs have generally been used to accelerate low-density plasmas with long current pulses. Operation of thrusters in this plasma regime allows for the development of certain undesirable particle-kinetic effects, such as Hall effect-induced current sheet canting. PPT-1 was designed to propel a highly collisional, dense plasma that has more fluid-like properties and, hence, is more effectively pushed by a magnetic field. The high-density plasma loading into the second stage of the accelerator is achieved through the use of a dense plasma injector (first stage). The injector produces a thermal plasma, derived from a molten lithium propellant feed system, which is subsequently accelerated by the second stage using mega-amp level currents, which eject the plasma at a speed on the order of 100 kilometers per second. Traditional PPTs also suffer from dynamic efficiency losses associated with snowplow loading of distributed neutral propellant. The twostage scheme used in PPT-I allows the propellant to be loaded in a manner which more closely approximates the optimal slug loading. Lithium propellant was chosen to test whether or not the reduced electrode erosion found in the Lithium Lorentz Force Accelerator (LiLFA) could also be realized in a pulsed plasma thruster. The use of the molten lithium dense plasma injector also eliminates the need for a gas valve and electrical switch; the injector design fulfills both roles, and uses no moving parts to provide, in principle, a highly reliable propellant feed and electrical switching system. Experimental results reported in this paper include: second-stage current traces, high-speed photographic and holographic imaging of the thruster exit plume, and internal mapping of the discharge chamber magnetic field from B-dot probe data. The magnetic field data is used to create a two-dimensional description of the evolution of the current sheet inside the thruster.

Markusic, Tom↗
NASA NTRS2014

Cryogenic Boil-Off Reduction System Testing

Cryogenic propellants such as liquid hydrogen (LH2) and liquid oxygen (LO2) are a part of NASA's future space exploration due to the high specific impulse that can be achieved using engines suitable for moving 10's to 100's of metric tons of payload mass to destinations outside of low earth orbit. However, the low storage temperatures of LH2 and LO2 cause substantial boil-off losses for missions with durations greater than several days. The losses can be greatly reduced by incorporating high performance cryocooler technology to intercept heat load to the propellant tanks and by the integration of self-supporting multi-layer insulation. The active thermal control technology under development is the integration of the reverse turbo- Brayton cycle cryocooler to the propellant tank through a distributed cooling network of tubes coupled to a shield in the tank insulation and to the tank wall itself. Also, the self-supporting insulation technology was utilized under the shield to obtain needed tank applied LH2 performance. These elements were recently tested at NASA Glenn Research Center in a series of three tests, two that reduced LH2 boil-off and one to eliminate LO2 boil-off. This test series was conducted in a vacuum chamber that replicated the vacuum of space and the temperatures of low Earth orbit. The test results show that LH2 boil-off was reduced 60% by the cryocooler system operating at 90K and that robust LO2 zero boil-off storage, including full tank pressure control was achieved.

Propellant Storage↗
NASA NTRS1972

Dual deflectable beam strip engine development.

This paper describes a dual beam thruster that has been designed, constructed, and tested. The system is suitable for two-axes attitude control and is comprised of two orthogonal strips, each capable of producing 0.30 mlb thrust and beam deflections of more than plus or minus 20 deg. The nominal specific impulse for the thruster is 5000 sec, and the thrust level from each strip can be varied from 0 to 100%. Neutralizer filaments that were developed and life tested over 2000 hours producing more than 40 mA of electron emission per watt of input power are also discussed. The system power required for clean ionizers is approximately 200 W.

Dulgeroff, C. R.↗
NASA NTRS2003

Solar Thermal Propulsion Improvements at Marshall Space Flight Center

Solar Thermal Propulsion (STP) is a concept which operates by transferring solar energy to a propellant, which thermally expands through a nozzle. The specific impulse performance is about twice that of chemical combustions engines, since there is no need for an oxidizer. In orbit, an inflatable concentrator mirror captures sunlight and focuses it inside an engine absorber cavity/heat exchanger, which then heats the propellant. The primary application of STP is with upperstages taking payloads from low earth orbit to geosynchronous earth orbit or earth escape velocities. STP engines are made of high temperature materials since heat exchanger operation requires temperatures greater than 2500K. Refractory metals such as tungsten and rhenium have been examined. The materials must also be compatible with hot hydrogen propellant. MSFC has three different engine designs, made of different refractory metal materials ready to test. Future engines will be made of high temperature carbide materials, which can withstand temperatures greater than 3000K, hot hydrogen, and provide higher performance. A specific impulse greater than 1000 seconds greatly reduces the amount of required propellant. A special 1 OkW solar ground test facility was made at MSFC to test various STP engine designs. The heliostat mirror, with dual-axis gear drive, tracks and reflects sunlight to the 18 ft. diameter concentrator mirror. The concentrator then focuses sunlight through a vacuum chamber window to a small focal point inside the STP engine. The facility closely simulates how the STP engine would function in orbit. The flux intensity at the focal point is equivalent to the intensity at a distance of 7 solar radii from the sun.

Gerrish, Harold P.↗
NASA NTRS1973

Testing and evaluation of the LES-6 pulsed plasma thruster by means of a torsion pendulum system

Performance characteristics of the LES-6 pulsed plasma thruster over a range of input conditions were investigated by means of a torsion pendulum system. Parameters of particular interest included the impulse bit and time average thrust (and their repeatability), specific impulse, mass ablated per discharge, specific thrust, energy per unit area, efficiency, and variation of performance with ignition command rate. Intermittency of the thruster as affected by input energy and igniter resistance were also investigated. Comparative experimental data correlation with the data presented. The results of these tests indicate that the LES-6 thruster, with some identifiable design improvements, represents an attractive reaction control thruster for attitude contol applications on long-life spacecraft requiring small metered impulse bits for precise pointing control of science instruments.

Hamidian, J. P.↗
NASA NTRS2014

Investigation of the Effects of Facility Background Pressure on the Performance and Voltage-Current Characteristics of the High Voltage Hall Accelerator

The National Aeronautics and Space Administration (NASA) Science Mission Directorate In-Space Propulsion Technology office is sponsoring NASA Glenn Research Center to develop a 4 kW-class Hall thruster propulsion system for implementation in NASA science missions. A study was conducted to assess the impact of varying the facility background pressure on the High Voltage Hall Accelerator (HiVHAc) thruster performance and voltage-current characteristics. This present study evaluated the HiVHAc thruster performance in the lowest attainable background pressure condition at NASA GRC Vacuum Facility 5 to best simulate space-like conditions. Additional tests were performed at selected thruster operating conditions to investigate and elucidate the underlying physics that change during thruster operation at elevated facility background pressure. Tests were performed at background pressure conditions that are three and ten times higher than the lowest realized background pressure. Results indicated that the thruster discharge specific impulse and efficiency increased with elevated facility background pressure. The voltage-current profiles indicated a narrower stable operating region with increased background pressure. Experimental observations of the thruster operation indicated that increasing the facility background pressure shifted the ionization and acceleration zones upstream towards the thruster's anode. Future tests of the HiVHAc thruster are planned at background pressure conditions that are expected to be two to three times lower than what was achieved during this test campaign. These tests will not only assess the impact of reduced facility background pressure on thruster performance, voltage-current characteristics, and plume properties; but will also attempt to quantify the magnitude of the ionization and acceleration zones upstream shifting as a function of increased background pressure.

Electric Propulsion↗
NASA NTRS1998

Performance of an Axisymmetric Rocket Based Combined Cycle Engine During Rocket Only Operation Using Linear Regression Analysis

The all rocket mode of operation is shown to be a critical factor in the overall performance of a rocket based combined cycle (RBCC) vehicle. An axisymmetric RBCC engine was used to determine specific impulse efficiency values based upon both full flow and gas generator configurations. Design of experiments methodology was used to construct a test matrix and multiple linear regression analysis was used to build parametric models. The main parameters investigated in this study were: rocket chamber pressure, rocket exit area ratio, injected secondary flow, mixer-ejector inlet area, mixer-ejector area ratio, and mixer-ejector length-to-inlet diameter ratio. A perfect gas computational fluid dynamics analysis, using both the Spalart-Allmaras and k-omega turbulence models, was performed with the NPARC code to obtain values of vacuum specific impulse. Results from the multiple linear regression analysis showed that for both the full flow and gas generator configurations increasing mixer-ejector area ratio and rocket area ratio increase performance, while increasing mixer-ejector inlet area ratio and mixer-ejector length-to-diameter ratio decrease performance. Increasing injected secondary flow increased performance for the gas generator analysis, but was not statistically significant for the full flow analysis. Chamber pressure was found to be not statistically significant.

Smith, Timothy D.↗
NASA NTRS2009

NEXT Long-Duration Test Neutralizer Performance and Erosion Characteristics

The NASA's Evolutionary Xenon Thruster (NEXT) program is developing the next-generation ion propulsion system with significant enhancements beyond the state-of-the-art to provide future NASA science missions with enhanced capabilities at a low total development cost. A Long-Duration Test (LDT) was initiated in June 2005, to verify the NEXT propellant throughput capability to a qualification-level of 450 kg, 1.5 times the anticipated throughput requirement of 300 kg per thruster based on mission analyses. As of September 2, 2009, the thruster has accumulated 24,400 hr of operation with extensive durations at the following input powers: 6.9, 4.7, 1.1, and 0.5 kW. The thruster has processed 434 kg of xenon, surpassing the NASA Solar Technology Application Readiness (NSTAR) program thruster propellant throughput demonstrated during the extended life testing of the Deep Space 1 flight spare ion thruster and approaching the NEXT development qualification throughput goal of 450 kg. The NEXT LDT has demonstrated a total impulse of 16.1 10(exp 6zzz0 N s; the highest total impulse ever demonstrated by an ion thruster. A reduction in neutralizer flow margin has been the only appreciable source of thruster performance degradation. The behavior of the neutralizer is not easily predicted due to both erosion and deposition observed in previous wear tests. Spot-to-plume mode transition flow data and in-situ erosion results for the LDT neutralizer are discussed. This loss of flow margin has been addressed through a combination of a design change in the prototype-model neutralizer to increase flow margin at low emission current and to update the NEXT throttle table to ensure adequate flow margin as a function of propellant throughput processed. The new throttle table will be used for future LDT operations. The performance of the NEXT LDT neutralizer is consistent with that observed for long-life hollow cathodes. The neutralizer life-limiting failure modes are progressing as expected and the neutralizer data indicate none of the neutralizer failures are imminent.

Herman, Daniel A.↗
NASA NTRS1972

Continuing development of the short-pulsed ablative space propulsion system.

Major advancements have been made in solid propellant pulsed plasma propulsion technology. A LES-6 microthruster has operated for 8900 hours in orbit. Millipound thrustors have been developed and tested at performance levels of as low as 112 watts/mlb. Efficiencies of up to 50% and specific impulse values up to 5000 sec have been measured. An improved microthrustor has operated at 207 watts/mlb continuously for 1440 hours. Correlations of design and performance data exist to allow thrustors to be built to meet a particular level of performance.

Palumbo, D. J.↗
NASA NTRS1977

Pseudo noise test set operation manual

Pseudonoise (PN) test set was designed to provide a reliable, operationally simple unit which allows the previously mentioned analog rms end-to-end error measurement of most communication systems to be easily performed. It also provides a band-limited pseudorandom noise as input, thereby escaping the disadvantages of most deterministic signals while retaining the advantages of using true-band-limited noise. In addition, the PN test set has the capability to be used as a means with which the autocorrelation function and impulse response may be determined for certain systems.

Weathers, G.↗
NASA NTRS1981

The 30 cm J Series Mercury Bombardment Thruster

The 30 cm J Series Thruster has been base-lined for the Solar Electric Propulsion Vehicle. This paper discusses the evolution of the thruster and details its performance, lifetime, and operational characteristics. An operational envelope covering nominal input powers of 2650 to 700 W is defined. Total thruster efficiencies over this envelope range from 0.711 to 0.472 with a corresponding impulse and thrust ranging from 2980 to 1850 sec and 129.1 to 36.4 mN, respectively. Completed life test segments and the ability of the thruster to meet the design goal of 15,000 hr of life as well as the life limitations are discussed. A description and rationale of the present set of thruster operational control algorithms is also presented.

Bechtel, R. T.↗
NASA NTRS1981

Properties of solar flare electrons, deduced from hard X-ray and spatially resolved microwave observations

An important question concerning an understanding of impulsive solar flares is related to the energetic electrons responsible for the microwave and the hard X-ray emission. A description is presented of an investigation in which spatially resolved microwave observations of an impulsive flare and hard X-ray data from the Solar Maximum Mission (SMM) are used to test the hypothesis that the two types of emission come from the same basic electron population. The considered observations are found to imply that the microwaves and hard X-rays were not produced by a common population of electrons with either a Maxwellian or single power-law energy distribution. It is suggested that the calculations should be repeated when observations of stronger events become available, for which a better determination of the X-ray spectrum is possible. The possibility is considered that microwaves and moderately hard X-rays come from spatially different regions.

Marsh, K. A.↗
NASA NTRS1991

Nuclear rocket propulsion. NASA plans and progress, FY 1991

NASA has initiated planning for a technology development project for nuclear rocket propulsion systems for space explorer initiative (SEI) human and robotic missions to the moon and Mars. An interagency project is underway that includes the Department of Energy National Laboratories for nuclear technology development. The activities of the project planning team in FY 1990 and 1991 are summarized. The progress to date is discussed, and the project plan is reviewed. Critical technology issues were identified and include: (1) nuclear fuel temperature, life, and reliability; (2) nuclear system ground test; (3) safety; (4) autonomous system operation and health monitoring; and (5) minimum mass and high specific impulse.

Clark, John S.↗
NASA NTRS1992

Recursive Inversion Of Externally Defined Linear Systems

Technical memorandum discusses mathematical technique described in "Recursive Inversion by Finite-Impulse-Response Filters" (ARC-12247). Technique is recursive algorithm yielding finite-impulse-response approximation of unknown single-input/single-output, causal, time-invariant, linear, real system, response of which is sequence of impulses. Useful in such diverse applications as medical diagnoses, identification of military targets, geophysical exploration, and nondestructive testing.

Bach, Ralph E., Jr.↗
NASA NTRS1991

Nuclear rocket propulsion: NASA plans and progress - FY 1991

NASA has initiated planning for a technology development project for nuclear rocket propulsion systems for space exploration initiative (SEI) human and robotic missions to the Moon and to Mars. An interagency project is underway that includes the Department of Energy National Laboratories for nuclear technology development. The activities of the project planning team in FY 1990 and 1991 are summarized. The progress to date is discussed, and the project plan is reviewed. Critical technology issues were identified and include: (1) nuclear fuel temperature, life, and reliability; (2) nuclear system ground test; (3) safety; (4) autonomous system operation and health monitoring; and (5) minimum mass and high specific impulse.

Clark, John S.↗
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.↗