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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 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Hydrogen arcjet technology

During the 1960's, a substantial research effort was centered on the development of arcjets for space propulsion applications. The majority of the work was at the 30 kW power level with some work at 1-2 kW. At the end of the research effort, the hydrogen arcjet had demonstrated over 700 hours of life in a continuous endurance test at 30 kW, at a specific impulse over 1000 s, and at an efficiency of 0.41. Another high power design demonstrated 500 h life with an efficiency of over 0.50 at the same specific impulse and power levels. At lower power levels, a life of 150 hours was demonstrated at 2 kW with an efficiency of 0.31 and a specific impulse of 935 s. Lack of a space power source hindered arcjet acceptance and research ceased. Over three decades after the first research began, renewed interest exists for hydrogen arcjets. The new approach includes concurrent development of the power processing technology with the arcjet thruster. Performance data were recently obtained over a power range of 0.3-30 kW. The 2 kW performance has been repeated; however, the present high power performance is lower than that obtained in the 1960's at 30 kW, and lifetimes of present thrusters have not yet been demonstrated. Laboratory power processing units have been developed and operated with hydrogen arcjets for the 0.1 kW to 5 kW power range. A 10 kW power processing unit is under development and has been operated at design power into a resistive load.

Sankovic, John M.↗
NASA NTRS1992

Experimental uncertainty survey and assessment

An uncertainty analysis and assessment of the specific impulse determination during Space Shuttle Main Engine testing is reported. It is concluded that in planning and designing tests and in interpreting the results of tests, the bias and precision components of experimental uncertainty should be considered separately. Recommendations for future research efforts are presented.

Coleman, Hugh W.↗
NASA NTRS1992

The production of He-3 and heavy ion enrichment in He-3-rich flares by electromagnetic hydrogen cyclotron waves

A new model is presented for the production of He-3 and heavy ion enrichments in He-3-rich flares using a direct single-stage mechanism. In analogy with the production of electromagnetic hydrogen cyclotron waves in earth's aurora by electron beams, it is suggested that such waves should exist in the electron acceleration region of impulsive solar flares. Both analytic and test-particle models of the effect of such waves in a nonuniform magnetic field show that these waves can selectively accelerate He-3 and heavy ions to MeV energies in a single-stage process, in contrast to other models which require a two-stage mechanism.

Temerin, M.↗
NASA NTRS1990

Using input command pre-shaping to suppress multiple mode vibration

Spacecraft, space-borne robotic systems, and manufacturing equipment often utilize lightweight materials and configurations that give rise to vibration problems. Prior research has led to the development of input command pre-shapers that can significantly reduce residual vibration. These shapers exhibit marked insensitivity to errors in natural frequency estimates and can be combined to minimize vibration at more than one frequency. This paper presents a method for the development of multiple mode input shapers which are simpler to implement than previous designs and produce smaller system response delays. The new technique involves the solution of a group of simultaneous non-linear impulse constraint equations. The resulting shapers were tested on a model of MACE, an MIT/NASA experimental flexible structure.

Hyde, James M.↗
NASA NTRS2013

Induction Heating Model of Cermet Fuel Element Environmental Test (CFEET)

Deep space missions with large payloads require high specific impulse and relatively high thrust to achieve mission goals in reasonable time frames. Nuclear Thermal Rockets (NTR) are capable of producing a high specific impulse by employing heat produced by a fission reactor to heat and therefore accelerate hydrogen through a rocket nozzle providing thrust. Fuel element temperatures are very high (up to 3000 K) and hydrogen is highly reactive with most materials at high temperatures. Data covering the effects of high‐temperature hydrogen exposure on fuel elements are limited. The primary concern is the mechanical failure of fuel elements due to large thermal gradients; therefore, high‐melting‐point ceramics‐metallic matrix composites (cermets) are one of the fuels under consideration as part of the Nuclear Cryogenic Propulsion Stage (NCPS) Advance Exploration System (AES) technology project at the Marshall Space Flight Center. The purpose of testing and analytical modeling is to determine their ability to survive and maintain thermal performance in a prototypical NTR reactor environment of exposure to hydrogen at very high temperatures and obtain data to assess the properties of the non‐nuclear support materials. The fission process and the resulting heating performance are well known and do not require that active fissile material to be integrated in this testing. A small‐scale test bed; Compact Fuel Element Environmental Tester (CFEET), designed to heat fuel element samples via induction heating and expose samples to hydrogen is being developed at MSFC to assist in optimal material and manufacturing process selection without utilizing fissile material. This paper details the analytical approach to help design and optimize the test bed using COMSOL Multiphysics for predicting thermal gradients induced by electromagnetic heating (Induction heating) and Thermal Desktop for radiation calculations.

Gomez, C. F.↗
NASA NTRS2013

Induction Heating Model of Cermet Fuel Element Environmental Test (CFEET)

Deep space missions with large payloads require high specific impulse and relatively high thrust to achieve mission goals in reasonable time frames. Nuclear Thermal Rockets (NTR) are capable of producing a high specific impulse by employing heat produced by a fission reactor to heat and therefore accelerate hydrogen through a rocket nozzle providing thrust. Fuel element temperatures are very high (up to 3000 K) and hydrogen is highly reactive with most materials at high temperatures. Data covering the effects of high‐temperature hydrogen exposure on fuel elements are limited. The primary concern is the mechanical failure of fuel elements due to large thermal gradients; therefore, high‐melting‐point ceramics‐metallic matrix composites (cermets) are one of the fuels under consideration as part of the Nuclear Cryogenic Propulsion Stage (NCPS) Advance Exploration System (AES) technology project at the Marshall Space Flight Center. The purpose of testing and analytical modeling is to determine their ability to survive and maintain thermal performance in a prototypical NTR reactor environment of exposure to hydrogen at very high temperatures and obtain data to assess the properties of the non‐nuclear support materials. The fission process and the resulting heating performance are well known and do not require that active fissile material to be integrated in this testing. A small‐scale test bed; Compact Fuel Element Environmental Tester (CFEET), designed to heat fuel element samples via induction heating and expose samples to hydrogen is being developed at MSFC to assist in optimal material and manufacturing process selection without utilizing fissile material. This paper details the analytical approach to help design and optimize the test bed using COMSOL Multiphysics for predicting thermal gradients induced by electromagnetic heating (Induction heating) and Thermal Desktop for radiation calculations.

Gomez, Carlos F.↗
NASA NTRS1991

Experimental and analytical comparison of flowfields in a 110 N (25 lbf) H2/O2 rocket

A gaseous hydrogen/gaseous oxygen 110 N (25 lbf) rocket was examined through the RPLUS code using the full Navier-Stokes equations with finite rate chemistry. Performance tests were conducted on the rocket in an altitude test facility. Preliminary parametric analyses were performed for a range of mixture ratios and fuel film cooling pcts. It is shown that the computed values of specific impulse and characteristic exhaust velocity follow the trend of the experimental data. Specific impulse computed by the code is lower than the comparable test values by about two to three percent. The computed characteristic exhaust velocity values are lower than the comparable test values by three to four pct. Thrust coefficients computed by the code are found to be within two pct. of the measured values. It is concluded that the discrepancy between computed and experimental performance values could not be attributed to experimental uncertainty.

Reed, Brian D.↗
NASA NTRS1993

Design and Development of a 3 to 10 kW Ammonia Arcjet

An ammonia arcjet capable of throttling between 3 and 10 kW and producing a specific impulse of 600 s is required for the SSTAR flight experiment. Testing was performed to evaluate the performance of two nozzle configurations on ammonia arcjet performance over this power range. One of the objectives of these tests was to quantify the effect small nozzle changes have on performance. The smaller constrictor engine (2.54 mm diameter) produced a specific impulse of about 650 s over the range of 3 to 10 kW at a specific power of 60 kJ/g exceeding the 500-600 s requirement for the SSTAR flight experiment.

Arcjets STAR Space Track and Autonomous Reposition↗
NASA NTRS1991

Experimental and analytical comparison of flowfields in a 110 N (25 Lbf) H2/O2 rocket

A gaseous hydrogen/gaseous oxygen 110 N (25 lbf) rocket has been examined through the RPLUS code using the full Navier-Stokes equations with finite-rate chemistry. Performance tests were conducted on the rocket in an altitude test facility. Preliminary parametric analyses have been performed for a range of mixture ratios and fuel film cooling percentages. It is shown that the computed values of specific impulse and characteristic exhaust velocity follow the trend of the experimental data. Specific impulse computed by the code is lower than the comparable test values by about two to three percent. The computed characteristic exhaust velocity values are lower than the comparable test values by three to four percent. Thrust coefficients computed by the code are found to be within two percent of the measured values. It is concluded that the discrepancy between computed and experimental performance values could not be attributed to experimental uncertainty.

Reed, Brian D.↗
NASA NTRS1994

Operating characteristics of the Russian D-55 thruster with anode layer

Performance measurements of a Russian engineering-model Thruster with Anode Layer (TAL) were obtained as part of a program to evaluate the operating characteristics of Russian Hall-thruster technology. The TAL model D-55 was designed to operate in the 1-2 kW power range on xenon. When received, the thruster had undergone only a few hours of acceptance testing by the manufacturer. Direct thrust measurements were obtained at a background pressure of 0.0003 Pa (2 x 10(exp -6) torr) at power levels ranging from 0.3 kW to 2.1 kW. At the nominal power level of 1.3 kW, a specific impulse level of 1600 s with a corresponding efficiency of 0.48 was attained. At all flow rates tested, the efficiency increased linearly with specific impulse until a maximum was reached, and then the efficiency leveled off. Increasing the anode flow rate shifted the efficiency upward, reaching 0.50 at 1850 s specific impulse. The thruster was equipped with inner and outer electromagnets which were isolated from the discharge and from each other. Variation of the magnetic field, obtained by changing the currents through the magnets, had little effect on performance, except at current levels below 70 percent of nominal. For a given operating condition, the performance was slightly affected by facility pressure. As the pressure was increased by a factor of thirty to 0.008 Pa (6 x 10(exp -5) torr), the current steadily increased by 4 percent, and the thrust increased by 2 percent. Performance comparisons were made with the Stationary Plasma Thruster, and the efficiency and specific impulse values were similar at power levels ranging from 0.9 kW to 1.5 kW. Endurance testing was not performed, and comparisons of lifetime were not made.

Sankovic, John M.↗
NASA NTRS2002

Effect of 670-nm Light-Emitting Diode Light On Neuronal Cultures

Light close to and within the near infrared range has documented benefits for promoting wound healing in human and animal studies. Our preliminary results using light-emitting diodes (LEDs) in this range have also demonstrated two-to five-fold increases in growth-phase-specific DNA synthesis in normal fibroblasts, muscle cells, osteoblasts, and mucosal epithelial cells in tissue cultures. However, the mechanisms of action of such light on cells are poorly understood. We hypothesized that the therapeutic effects of such light result from the stimulation of cellular events associated with increases in cytochrome oxidase activity. As a first step in testing our hypothesis, we subjected primary neuronal cultures to impulse blockade by tetrodotoxin (TTX), a voltage-dependent sodium channel blocker, and applied LED light at 670 nm to determine if it could partially or fully reverse the reduction of cytochrome oxidase activity by TTX. The wavelength and parameters were previously tested to be beneficial for wound healing.

Wong-Riley, Margaret T. T.↗
NASA NTRS1988

Performance and lifetime assessment of MPD arc thruster technology

A summary of performance and lifetime characteristics of pulsed and steady-state magnetoplasmadynamic (MPD) thrusters is presented. The technical focus is on cargo vehicle propulsion for exploration-class missions to the Moon and Mars. Relatively high MPD thruster efficiencies of 0.43 and 0.69 have been reported at about 5000 s specific impulse using hydrogen and lithium, respectively. Efficiencies of 0.10 to 0.35 in the 1000 to 4500 s specific impulse range have been obtained with other propellants (e.g., Ar, NH3, N2). Thermal efficiency data in excess of 0.80 at MW power levels using pulsed thrusters indicate the potential of high MPD thruster performance. Extended tests of pulsed and steady-state MPD thrusters yield total impulses at least two to three orders of magnitude below that necessary for cargo vehicle propulsion. Performance tests and diagnostics for life-limiting mechanisms of megawatt-class thrusters will require high fidelity test stands which handle in excess of 10 kA and a vacuum facility whose operational pressure is less than 3 x 10 to the -4 torr.

Sovey, James S.↗
NASA NTRS1988

Performance and lifetime assessment of MPD arc thruster technology

A summary of performance and lifetime characteristics of pulsed and steady-state magnetoplasmadynamic (MPD) thrusters is presented. The technical focus is on cargo vehicle propulsion for exploration-class missions to the Moon and Mars. Relatively high MPD thruster efficiencies of 0.43 and 0.69 have been reported at about 5000 s specific impulse using hydrogen and lithium, respectively. Efficiencies of 0.10 to 0.35 in the 1000 to 4500 s specific impulse range have been obtained with other propellants (e.g., Ar, NH3, N2). Thermal efficiency data in excess of 0.80 at MW power levels using pulsed thrusters indicate the potential of high MPD thruster performance. Extended tests of pulsed and steady-state MPD thrusters yield total impulses at least two to three orders of magnitude below that necessary for cargo vehicle propulsion. Performance tests and diagnostics for life-limiting mechanisms of megawatt-class thrusters will require high fidelity test stands which handle in excess of 10 kA and a vacuum facility whose operational pressure is less than 3 x 10 to the -4 torr.

Sovey, James S.↗
NASA NTRS1976

Transient excitation and data processing techniques employing the fast fourier transform for aeroelastic testing

The development of testing techniques useful in airplane ground resonance testing, wind tunnel aeroelastic model testing, and airplane flight flutter testing is presented. Included is the consideration of impulsive excitation, steady-state sinusoidal excitation, and random and pseudorandom excitation. Reasons for the selection of fast sine sweeps for transient excitation are given. The use of the fast fourier transform dynamic analyzer (HP-5451B) is presented, together with a curve fitting data process in the Laplace domain to experimentally evaluate values of generalized mass, model frequencies, dampings, and mode shapes. The effects of poor signal to noise ratios due to turbulence creating data variance are discussed. Data manipulation techniques used to overcome variance problems are also included. The experience is described that was gained by using these techniques since the early stages of the SST program. Data measured during 747 flight flutter tests, and SST, YC-14, and 727 empennage flutter model tests are included.

Jennings, W. P.↗
NASA NTRS1992

Results of a 1462 hour ammonia arcjet endurance test

A total of 1462 hours of operaton were demonstrated in an endurance test of a 30 kWe-class ammonia arcjet operated at 10 kWe. The propellant flow rate was 0.170 g/s, and the measured performance increased from approximately 650 s specific impulse at 36 percent efficiency at the beginning of the test to 675 s at 39 percent near the end. The voltage increased and the current dropped slightly over the first 400 hours, and then remained approximately constant for the remainder of the test. The test, scheduled for 1500 hours, was terminated after an insulator in the rear of the engine failed. Post-test examination of the electrodes revealed only minimal damage. Although there was evidence of a number of mass transport processes occurring in the discharge chamber, the primary life-limiting wear mechanisms appear to be cathode tip erosion and constrictor melting.

Polk, J. E.↗
NASA NTRS1960

Flight Performance of a Spin-Stabilized 20-Inch-Diameter Solid-Propellant Spherical Rocket Motor

A successful flight test of a spin-stabilized 20-inch-diameter solid-propellant rocket motor having a propellant mass fraction of 0.92 has been made. The motor was fired at altitude after being boosted by a three-stage test vehicle. Analysis of the data indicates that a total impulse of 44,243 pound-second with a propellant specific impulse of approximately 185 was achieved over a total action time of about 12 seconds. These results are shown to be in excellent agreement with data from ground static firing tests of these motors. The spherical rocket motor with an 11-pound payload attained a velocity of 15,620 feet per second (m = 16.7) with an incremental velocity increase for the spherical motor stage of 12,120 feet per second.

Levine, Jack↗
NASA NTRS1988

Electro-impulse de-icing - A status report

The advantages and disadvantages of the Electro-Impulse De-Icing (EIDI) system are examined. The design and operation of the EIDI are described. The effect of repeated impulsive loads on the structure and skin of the aircraft surface is investigated. It is observed that the wing skin directly over the coil receives the greatest stresses. Data from the testing of metal leading edges reveal that after 11,500 impulses cracks appear in the sheet brackets attaching the ribs to the coil beams. The composite leading edge was evaluated for fatigue, and after 20,000 impulses no cracks were detected. The advantages of the band-aid coil mount are discussed.

Zumwalt, G. W.↗
NASA NTRS1972

Development of a radioisotope-fueled thruster for satellite propulsion.

Description of the design and testing of two distinct thrustor concepts which use ammonia propellant and have a specific impulse as high as 310 sec during pulsed operation. Considerations of thrustor design, material selection, and system testing are discussed in the light of performance, lifetime, and nuclear safety requirements. Results of a detailed study of the advantages and disadvantages of the use of such a thrustor are presented.

Rosen, S. G.↗