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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 757 records · Page 42
All sourcesNASA NTRSDOE OSTI
NASA NTRS1989

Lunar transfer vehicle design issues with electric propulsion systems

This paper describes parametric design studies of electric propulsion lunar transfer vehicles. In designing a lunar transfer vehicle, selecting the 'best' operating points for the design parameters allows significant reductions in the mass in low earth orbit (LEO) for the mission. These parameters include the specific impulse, the power level, and the propulsion technology. Many of the decisions regarding the operating points are controlled by the propulsion and power system technologies that are available for the spacecraft. The relationship between these technologies is discussed and analyzed here. It is found that both ion and MPD propulsion offer significant LEO mass reductions over O2/H2 for lunar transfer vehicle missions. The recommended operating points for the lunar transfer vehicle are an I(sp) of 5000 lb(f)-s/lb(m) and a 1 MW power level. For large lunar missions, krypton may be the best choice for ion propulsion.

Palaszewski, Bryan↗
NASA NTRS1989

Baseline spacecraft and mission design for the SP-100 flight experiment

The design and performance of a spacecraft employing arcjet nuclear electric propulsion, suitable for use in the SP-100 Space Reactor Power System (SRPS) Flight Experiment, are outlined. The vehicle design is based on a 93 kWe ammonia arcjet system operating at an experimentally-measured specific impulse of 1030 s and an efficiency of 42 percent. The arcjet/gimbal assemblies, power conditioning subsystem, propellant feed system, propulsion system thermal control, spacecraft diagnostic instrumentation, and the telemetry requirements are described. A 100 kWe SRPS is assumed. The total spacecraft mass is baselined at 5675 kg excluding the propellant and propellant feed system. Four mission scenarios are described which are capable of demonstrating the full capability of the SRPS. The missions considered include spacecraft deployment to possible surveillance platform orbits, a spacecraft storage mission and an orbit raising round trip corresponding to possible orbit transfer vehicle missions. Launches from Kennedy Space Center using the Titan IV expendable launch vehicle are assumed.

Deininger, William D.↗
NASA NTRS1989

Performance of a 100 kW class applied field MPD thruster

The performance of a 100 kW class, applied field MPD thruster was evaluated and sensitivities of discharge characteristics to arc current, mass flow rate, and applied magnetic field were investigated. Thermal efficiencies as high as 60 percent , thrust efficiencies up to 21 percent, and specific impulses of up to 1150 s were attained with argon propellant. Thrust levels up to 2.5 N were directly measured with an inverted pendulum thrust stand at discharge input powers up to 57 kW. It was observed that thrust increased monotonically with the product of arc current and magnet current.

Mantenieks, M. A.↗
NASA NTRS1989

Ion optics for high power 50-cm-dia ion thrusters

The process used at the NASA Lewis Research Center to fabricate 30 and 50-cm-diameter ion optics is described. The ion extraction capabilities of the 30 and 50-cm diameter ion optics were evaluated on divergent field and ring-cusp discharge chambers and compared. Perveance was found to be sensitive to the effects of the type and power of the discharge chamber and to the accelerator electrode hole diameter. Levels of up to 0.64 N and 20 kW for thrust and input power, respectively, were demonstrated with the divergent-field discharge chamber. Thruster efficiencies and specific-impulse values up to 79 percent and 5000 seconds, respectively, were achieved with the ring-cusp discharge chamber.

Rawlin, Vincent K.↗
NASA NTRS1989

Ignition and combustion of metallized propellants

An experimental program was conducted to investigate the ignition and combustion properties of a gelled, metallized fuel, A1/JP20, burned with gaseous oxygen, and to compare the results to the ignition and combustion properties of neat RP1. Both a Cstar efficiency and a specific impulse efficiency were calculated by comparing experimental values with those from a one-dimensional chemical equilibrium computer code. For most operating conditions, the efficiencies of the A1/JP10 were only 2 to 6 percent less than those of the neat RP1.

Galecki, Diane L.↗
NASA NTRS1988

Spacecraft and mission design for the SP-100 flight experiment

The design and performance of a spacecraft employing arcjet nuclear-electric propulsion, suitable for use in the SP-100 space-reactor power-system (SRPS) flight experiment, are outlined. The vehicle design is based on a 93-kWe ammonia arcjet system operating at an experimentally measured specific impulse of 1031 sec and an efficiency of 42.3 percent. The arcjet/gimbal assemblies, power conditioning subsystem, propellant feed system, propulsion system thermal control, spacecraft diagnostic instrumentation, and the telemetry requirements are described. A 100-kWe SRPS is assumed. The spacecraft mass is baselined at 5675 kg, excluding the propellant and propellant feed system. Four mission scenarios are described which are capable of demonstrating the full capability of the SRPS.

Deininger, William D.↗
NASA NTRS1988

Experiments and analysis of a compact electrothermal thruster

The description and experimental performance of a compact microwave electrothermal thruster are presented. Experimental tests using N and He propellants with input power levels of 200 W - 1.5 kW are presented. Experimental results produced energy efficiencies of 20-60 percent and specific impulse of 250-450 sec.

Asmussen, Jes↗
NASA NTRS1989

Space Shuttle propulsion performance reconstruction from flight data

The aplication of extended Kalman filtering to estimating Space Shuttle Solid Rocket Booster (SRB) performance, specific impulse, from flight data in a post-flight processing computer program. The flight data used includes inertial platform acceleration, SRB head pressure, and ground based radar tracking data. The key feature in this application is the model used for the SRBs, which represents a reference quasi-static internal ballistics model normalized to the propellant burn depth. Dynamic states of mass overboard and propellant burn depth are included in the filter model to account for real-time deviations from the reference model used. Aerodynamic, plume, wind and main engine uncertainties are included.

Rogers, Robert M.↗
NASA NTRS1989

Trajectory optimization and guidance for an aerospace plane

The first step in the approach to developing guidance laws for a horizontal take-off, air breathing single-stage-to-orbit vehicle is to characterize the minimum-fuel ascent trajectories. The capability to generate constrained, minimum fuel ascent trajectories for a single-stage-to-orbit vehicle was developed. A key component of this capability is the general purpose trajectory optimization program OTIS. The pre-production version, OTIS 0.96 was installed and run on a Convex C-1. A propulsion model was developed covering the entire flight envelope of a single-stage-to-orbit vehicle. Three separate propulsion modes, corresponding to an after burning turbojet, a ramjet and a scramjet, are used in the air breathing propulsion phase. The Generic Hypersonic Aerodynamic Model Example aerodynamic model of a hypersonic air breathing single-stage-to-orbit vehicle was obtained and implemented. Preliminary results pertaining to the effects of variations in acceleration constraints, available thrust level and fuel specific impulse on the shape of the minimum-fuel ascent trajectories were obtained. The results show that, if the air breathing engines are sized for acceleration to orbital velocity, it is the acceleration constraint rather than the dynamic pressure constraint that is active during ascent.

Mease, Kenneth D.↗
NASA NTRS1987

Arcjet thruster research and technology. Phase 1: Executive summary

The principal objective of this two phase program is to conduct the development research required to make the low power arcjet a flight ready technology. Many important results were obtained during Phase 1 to move closer to this objective. Fundamental analyses were performed of the arcjet nozzle, the gas kinetic reaction effects, the thermal environment, and the arc stabilizing vortex. These aided the conceptual understanding of the arcjet and guided design work. A hydrazine (N2H4) arcjet was designed that combined a flight qualified catalyst bed with a modular arcjet. Extensive testing was performed which demonstrated the feasibility of using this propellant in an arcjet for the first time. Startup techniques were developed, stability maintained, material compatibility tests conducted, and performance mapping tests performed. Specific impulse values from 400 to 730 seconds were produced with a non-optimized design. These levels are higher than were originally thought possible and proved that extremely high enthalpy values can be obtained with constricted arc technology. Erosion rate data are promising for lifetime extensions to meet flight application requirements. Power control unit (PCU) development was started with the design and fabrication of a laboratory high switching frequency supply. Valuable data were obtained on PCU operation and on the interaction with the dynamic arc. Phase 2 efforts presently underway are resolving key issues for multi-hundred hour lifetimes, are continuing to investigate arcjet/PCU interactions, and will demonstrate duty cycle N2H4 arcjet/PCU operation in a simulated flight mode for lifetimes consistent with initial applications.

Source record↗
NASA NTRS1990

Atomic hydrogen as a launch vehicle propellant

An analysis of several atomic hydrogen launch vehicles was conducted. A discussion of the facilities and the technologies that would be needed for these vehicles is also presented. The Gross Liftoff Weights (GLOW) for two systems were estimated; their specific impulses (I sub sp) were 750 and 1500 lb(sub f)/s/lb(sub m). The atomic hydrogen launch vehicles were also compared to the currently planned Advanced Launch System design concepts. Very significant GLOW reductions of 52 to 58 percent are possible over the Advanced Launch System designs. Applying atomic hydrogen propellants to upper stages was also considered. Very high I(sub sp) (greater than 750 lb(sub f)/s/lb(sub m)) is needed to enable a mass savings over advanced oxygen/hydrogen propulsion. Associated with the potential benefits of high I(sub sp) atomic hydrogen are several challenging problems. Very high magnetic fields are required to maintain the atomic hydrogen in a solid hydrogen matrix. The magnetic field strength was estimated to be 30 kilogauss (3 Tesla). Also the storage temperature of the propellant is 4 K. This very low temperature will require a large refrigeration facility for the launch vehicle. The design considerations for a very high recombination rate for the propellant are also discussed. A recombination rate of 210 cm/s is predicted for atomic hydrogen. This high recombination rate can produce very high acceleration for the launch vehicle. Unique insulation or segmentation to inhibit the propellant may be needed to reduce its recombination rate.

Palaszewski, Bryan A.↗
NASA NTRS1990

On the use of external burning to reduce aerospace vehicle transonic drag

The external combustion of hydrogen to reduce the transonic drag of aerospace vehicles is currently being investigated. A preliminary analysis based on a constant pressure control volume is discussed. Results indicate that the specific impulse of the external burning process rivals that of a turbojet and depends on the severity of the initial base drag as well as on the Mach flight number and the equivalence ratio. A test program was conducted to investigate hydrogen-air flame stability at the conditions of interest and to demonstrate drag reduction on a simple expansion ramp. Initial test results are presented and compared with the control-volume analysis. The expansion ramp surface pressure coefficient showed little variation with fuel pressure and altitude, in disagreement with the analysis. Flame stability results were encouraging and indicate that stable combustion is possible over an adequate range of conditions. Facility interference and chemical kinetics phenomena that make interpretation of subscale ground test data difficult are discussed.

Trefny, Charles J.↗
NASA NTRS1990

Arcjet load characteristics

Experiments were conducted to define the interface characteristics and constraints of 1 kW class arcjets run on simulated decomposition products of hydrazine and power processors. The impacts of power supply output current ripple on arcjet performance were assessed by variation of the ripple frequency from 100 Hz to 100 kHz with 10 percent peak-to-peak ripple amplitude at 1.2 kW. Ripple had no significant effects on thrust, specific impulse or efficiency. The impact of output ripple on thruster lifetime was not assessed. The static and dynamic impedances of the arcjet were quantified with two thrusters of nearly identical configuration. Superposition of an AC component on the DC arc current was used to characterize the dynamic impedance as a function of flow rate and DC current level. A mathematical model was formulated from these data. Both the static and dynamic impedance magnitude were found to be dependent on mass flow rate. The amplitude of the AC component was found to have little effect on the dynamic impedance. Reducing the DC level from 10 to 8 amps led to a large change in the magnitude of the dynamic impedance with no observable phase change. The impedance data compared favorably between the two thrusters.

Hamley, John A.↗
NASA NTRS1990

Laboratory Simulation of the Effect of Rocket Thrust on a Precessing Space Vehicle

Ground tests of solid propellant rocket motors have shown that metal-containing propellants produce various amounts of slag (primarily aluminum oxide) which is trapped in the motor case, causing a loss of specific impulse. Although not yet definitely established, the presence of a liquid pool of slag also may contribute to nutational instabilities that have been observed with certain spin-stabilized, upper-stage vehicles. Because of the rocket's axial acceleration, absent in the ground tests, estimates of in-flight slag mass have been very uncertain. Yet such estimates are needed to determine the magnitude of the control authority of the systems required for eliminating the instability. A test rig with an eccentrically mounted hemispherical bowl was designed and built which incorporates a follower force that properly aligns the thrust vector along the axis of spin. A program that computes the motion of a point mass in the spinning and precessing bowl was written. Using various RPMs, friction factors, and initial starting conditions, plots were generated showing the trace of the point mass around the inside of the fuel tank. The apparatus will incorporate future design features such as a variable nutation angle and a film height measuring instrument. Data obtained on the nutational instability characteristics will be used to determine order of magnitude estimates of control authority needed to minimize the sloshing effect.

Alvarez, Oscar↗
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 NTRS1990

Flight motor set 360L001 (STS-26R). Volume 1: System overview, revision A

The NASA space shuttle flight STS-26R, launched at 11:37.00.009 am, EDT on 29 Sep. 1988, used the redesigned solid rocket motors (RSRM) 360L001A and 360L001B. Evaluation of the ground environment instrumentation (GEI) data recorded prior to flight showed no launch commit criteria violations; that the field joint heater and aft skirt thermal conditioning systems performed adequately; and that the GEI data showed good agreement with thermal model predictions. Evaluation of the developmental flight instrumentation (DFI) revealed excellent agreement with both the predicted and required ballistic specifications. All parameters were well within the GEI specification requirements including propellant burn rates, specific impulse values, and thrust imbalance. Recorded strain values also indicated satisfactory radial growth and stress levels, as well as verification of adequate safety factors. Postflight inspection of the insulation, seals, case, and nozzles showed overall excellent performance. Some thermal DFI protective cork was missing, and inoperative field joint vent valves on the thermal protection cork allowed water entry into the field joints upon splashdown. Evaluation of these anomalies, as well as complete evaluation of all Redesigned Solid Rocket Motor components, is contained.

Garecht, Diane M.↗
NASA NTRS1988

High/variable mixture ratio oxygen/hydrogen engines

A LOX/LH2 high/variable mixture ratio booster upper stage is described. The engine has high thrust-weight ratio as a booster and high specific impulse as an upper stage engine. Operation at high mixture ratio utilizes the propellants at high bulk density. The engine may use multiple turbopump-preburners for higher thrust ratings. The engine uses the full flow cycle to obtain minimum turbine inlet temperatures for a given chamber pressure and to avoid interpropellant shaft seals and other single point failure modes. A portion of the liquid hydrogen is used to regeneratively cool the thrust chamber assembly. The warmed hydrogen coolant is then used to drive the fuel boost turbopump. All propellants arrive at the gas-gas injector ready to burn. Shear mixing of the parallel flowing high velocity, low density fuel-rich gases with the high density, low velocity oxidizer-rich gases provides complete combustion with a modest chamber volume. Combustion stability is assured by the injection of the heated fuel-rich gases and the comparatively low volume ratio of the propellants before and after combustion. The high area ratio nozzle skirt is fitted with a low area ratio nozzle skirt insert for optimum low altitude performance. The overall engine characteristics make it a candidate for ALS, Shuttle-C, LRB, and SSTO applications.

Knuth, William H.↗
NASA NTRS1990

Low power arcjet performance

An experimental investigation was performed to evaluate arc jet operation at low power. A standard, 1 kW, constricted arc jet was run using nozzles with three different constrictor diameters. Each nozzle was run over a range of current and mass flow rates to explore stability and performance in the low power engine. A standard pulse-width modulated power processor was modified to accommodate the high operating voltages required under certain conditions. Stable, reliable operation at power levels below 0.5 kW was obtained at efficiencies between 30 and 40 percent. The operating range was found to be somewhat dependent on constrictor geometry at low mass flow rates. Quasi-periodic voltage fluctuations were observed at the low power end of the operating envelope, The nozzle insert geometry was found to have little effect on the performance of the device. The observed performance levels show that specific impulse levels above 350 seconds can be obtained at the 0.5 kW power level.

Curran, Francis M.↗