Colloid-particle electrostatic thrustors
Specific impulse requirements for electric rocket lunar ferries
SPECIFIC IMPULSE↗
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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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Accelerator grid durability tests of mercury electron-bombardment ion thrustors
Accelerator grid durability tests of mercury electron bombardment ion thrusters
ELECTRON BOMBARDMENT↗
Experimental studies with a mercury bombardment thrustor system
Performance tests of mercury electron bombardment ion engine system
Masek, T. D.↗
ATS-III resistojet thrustor system performance.
ATS 3 spacecraft ammonia-fueled resistojet engine test performance
Lund, W.↗
Feasibility studies on an auxiliary propulsion system using MPD thrustors
Feasibility analysis of auxiliary propulsion system using MPD thrusters
Burkhart, J. A.↗
Chemical nonequilibrium effects in biowaste thrustor performance
Resistojet power and specific impulse performance, investigating biowaste derived propellant chemical nonequilibrium effects
Gaubatz, W. A.↗
Gallium Electromagnetic (GEM) Thrustor Concept and Design
We describe the design of a new type of two-stage pulsed electromagnetic accelerator, the gallium electromagnetic (GEM) thruster. A schematic illustration of the GEM thruster concept is given in Fig. 1. In this concept, liquid gallium propellant is pumped into the first stage through a porous metal electrode using an electromagneticpump[l]. At a designated time, a pulsed discharge (approx.10-50 J) is initiated in the first stage, ablating the liquid gallium from the porous electrode surface and ejecting a dense thermal gallium plasma into the second state. The presence of the gallium plasma in the second stage serves to trigger the high-energy (approx.500 I), send-stage puke which provides the primary electromagnetic (j x B) acceleration.
Polzin, Kurt A.↗
Ion thruster thermal characteristics and performance.
Experimental and analytical results for a typical 20-cm-diam, hollow-cathode ion thrustor are reported. The foundation of the investigation was the application of thermal model correction techniques. Pertinent thermal properties and plasma heating characteristics of the thrustor were determined through correlation and integration of temperature measurement data with a single-state Wiener-Kalman filter. The thrustor self-heating levels on various parts were realistically estimated. Analytically predicted temperatures were forced to agree with the measured values for the purpose of constructing a corrected thermal model, which could then be used to evaluate more realistic thrustor circumstances and environments. The expected accuracy of the resultant analytical network model was demonstrated to be plus or minus 10 K. Thrustor thermal performance data for a typical five-thrustor array are presented as functions of environmental solar intensities. The thermal analyses are also extended to a 30-cm thrustor system.
Wen, L.↗
Optical radiation from regions downstream of mercury bombardment thrusters.
Use of a 0.5-m focal length, plane grating monochromator to measure the radiance of spectral radiation emanating from regions downstream of a mercury bombardment thrustor in the wavelength range investigated was 2800 to 6000 A. This radiation was due primarily to the radiative decay of excited mercury atoms exhausted from the thrustor. Radiance values ranged from 10 to the minus 11th to 10 to the min scm/ster varying with wavelength. For resonant radiation, the spectral radiance may exceed 10 to the minus 8th W/sq cm/ster. From such radiance measurements, it was concluded that the thrustor background radiation should not interfere with the control functions of a star tracker viewing through the thrustor exhaust, provided that the tracker is designed to operate with a sufficiently small field of view. Problems may be encountered, however, during the spacecraft acquisition phase where a larger field of view may be required. Here the thrustor exhaust radiation may be comparable to the star light flux. This problem may be circumvented by locating the tracker view axis so as not to view downstream of the thrustor.
Milder, N. L.↗
An 8-cm electron bombardment thruster for auxiliary propulsion
Thrustor size, beam current level, and specific impulse tradeoffs are considered for mercury electron bombardment ion thrustors to be used for north-south station keeping of geosynchronous spacecraft. An 8-cm-diam thrustor operating at 2750 sec specific impulse at thrust levels of 4.4 mN (1 mlb) to 8.9 mN (2 mlb) with a design life of 20,000 hours and 10,000 cycles is being developed. The thrustor will have a dished two-grid system capable of thrust vectoring of plus or minus 10 deg in two orthogonal directions. A preliminary thrustor has been fabricated and tested; thrustor performance characteristics have been determined at 4.45, 6.68, and 8.90 millinewtons.
Hudson, W. R.↗
Investigation of biowaste resistojets for space station application
The feasibility of using electrically conducting ceramics to heat biowaste propellants to 2000 K in resistojet thrustors was demonstrated. These thrustors are being developed for use on the space station. Among the candidate ceramic heater materials, zirconia and thoria are chemically resistant to the biopropellants, and they are also sufficiently conductive at high temperatures to make them suitable for the heater elements in these thrustors. A proof of concept thrustor design is presented, incorporating a multiple passage cylindrical heater made of zirconia ceramic which is capable of operating at 2000 K wall temperature with CO2 and H2O biopropellants. For the 25 mlb size thrustor, specific impulses of 200 seconds for CO2 and 275 seconds for H2O biopropellants are predicted.
Halbach, C. R.↗
Status of SERT II thrusters and spacecraft - 1976
The historical record of the SERT II ion thrustors and spacecraft performance for 6-1/2 years since the February 1970 launch is reviewed. The most recent ion thrustor operation test shows no changes since 1974. Thrustor 2 is fully operational with no performance degradation. Thrustor 1 has a high voltage grid short, but continues to demonstrate cathode and discharge relight capability. Spacecraft orbit and dynamic analysis indicates a stable, sun-synchronous spacecraft orientation by 1979. An attitude adjustment maneuver was performed in August 1976 to achieve this orientation and provide sufficient continuous solar power for thrustor operation in 1979.
Kerslake, W. R.↗
(abstract) An Assessment of Electric Propulsion Research, Development, and Application in the United States
This paper will discuss the development of Electric Propulsion technology in the U.S. from the 1960's to the present. It will summarize the various activities related to arcjets, resistojets, pulsed plasma thrustors, magneto-plasma-dynamic thrustors, ion engines, and more recently the evaluation of Hall effect thrustors of the SPT or Anode Layer type developed in Russia. Also, demonstration test flight and actual mission applications will be summarized. Finally, the future application of electric propulsion to near-term commercial communications satellites and planetary missions will be projected. This history is rich in diversity, and has involved a succession of types of thrustors, propellants, and electric power sources. With the recent use of arcjets on commercial communication satellites and the flight tests of ion engines for this application, it appears that electric propulsion is finally on the verge of wide spread application.
ion engines electric propulsion arcjets resistojet↗
ATS-6 - Cesium bombardment engine north-south stationkeeping experiment
Two 0.004 N thrust cesium bombardment ion thrustors have been developed and used for north-south stationkeeping in the geostationary Applications Technology Satellite-6 (ATS-6). The thrustor subsystems are mounted on the north and south faces of the earth viewing module such that 0.0026 N of thrust is applied normal to the orbit plane and 0.0036 N is applied radially upward. The change in the orbit inclination of the satellite is maintained at zero by operating the two thrustors alternately so that their thrust components, normal to the orbital plane, are symmetrically applied about the nodal crossings. Initial operation of the thrustors was successful. There was no interference with the satellite communications systems and the predicted spacecraft operating potential was verified. Subsequent trials failed due to a defect in the operation of the propellant reservoirs in zero g. A feed line valve is under development to correct this difficulty.
Worlock, R. M.↗
Electrothermal hydrazine thruster development.
Results of testing of several electrically-heated, thermal-decomposition hydrazine thrustors in the 5- to 70-mlb thrust range. Propellant supply pressures are compatible with those typical of larger catalytic thrustors. With a 5-W power input, pulsed specific impulse varies from about 165 to 215 sec, depending on duty cycle. The thrustors yield highly reproducible 1-mlb-sec impulse bits with 50-msec command pulse widths. The steady-state specific impulse exceeds 230 sec for thrust levels above 10 mlb.
Murch, C. K.↗
Optical properties of thermal control coatings contaminated by MMH/N2O4 5-pound thruster in a vacuum environment with solar simulation.
Cat-a-lac Black and S13G thermal control coatings were exposed to the exhaust of a thrustor in a simulated space environment. Vacuum was maintained at less than 10 microtorr during thrustor firing in the liquid helium cooled facility. The thrustor was fired in a 50-millisecond pulse mode, and the accumulated firing time was 224 seconds. Solar absorptance and thermal emittance of the coatings were measured in-situ at intervals of 300 pulses, using a calorimetric technique. The Cat-a-lac Black coatings showed no change in solar absorptance or thermal emittance. The S13G showed up to 25% increase in solar absorptance but no change in thermal emittance.
Sommers, R. D.↗
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.↗
Hydrogen-oxygen Space Shuttle ACPS thruster technology review.
A comprehensive program has provided the technology groundwork for the use of hydrogen-oxygen propellants in the Space Shuttle Attitude Control Propulsion System (ACPS) thrustors. This work has concentrated on generation of technology for injectors, cooled thrust chambers, valves, and ignition systems. The thrustors are designed to meet a unique and stringent set of requirements, including: long life for 100 mission reuses, high performance, light weight, ability to provide long duration firings as well as small impulse bits, ability to operate over wide ranges of propellant inlet conditions and to withstand reentry heating. The program has included evaluation of thrustors designed for ambient temperature and cold gaseous propellants at the vehicle interface.
Gregory, J. W.↗