Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “Thrusters”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 685 records · Page 38
All sourcesNASA NTRSDOE OSTI
NASA NTRS1979

Liquid oxygen/liquid hydrogen auxiliary power system thruster investigation

The design, fabrication, and demonstration of a 111 newton (25 lb) thrust, integrated auxiliary propulsion system (IAPS) thruster for use with LH2/LO2 propellants is described. Hydrogen was supplied at a temperature range of 22 to 33 K (40 to 60 R), and oxygen from 89 to 122 K (160 to 220 R). The thruster was designed to operate in both pulse mode and steady-state modes for vehicle attitude control, space maneuvering, and as an abort backup in the event of failure of the main propulsion system. A dual-sleeve, tri-axial injection system was designed that utilizes a primary injector/combustor where 100 percent of the oxygen and 8 percent of the hydrogen is introduced; a secondary injector/combustor where 45 percent of the hydrogen is introduced to mix with the primary combustor gases; and a boundary layer injector that uses the remaining 45 percent of the hydrogen to cool the thrust throat/nozzle design. Hot-fire evaluation of this thruster with a BLC injection distance of 2.79 cm (1.10 in.) indicated that a specific impulse value of 390 sec can be attained using a coated molybdenum thrust chamber. Pulse mode tests indicated that a chamber pressure buildup to 90 percent thrust can be achieved in a time on the order of 48 msec. Some problems were encountered in achieving ignition of each pulse during pulse trains. This was interpreted to indicate that a higher delivered spark energy level ( 100 mJ) would be required to maintain ignition reliability of the plasma torch ignition system under the extra 'cold' conditions resulting during pulsing.

Eberle, E. E.↗
NASA NTRS1979

Plasma property and performance prediction for mercury ion thrusters

The discharge chambers of mercury ion thrusters are modelled so the principal effects and processes which govern discharge plasma properties and thruster performance are described. The conservation relations for mass, charge and energy when applied to the Maxwellian electron population in the ion production region yield equations which may be made one-dimensional by the proper choice of coordinates. Solutions to these equations with the appropriate boundary conditions give electron density and temperature profiles which agree reasonably well with measurements. It is then possible to estimate plasma properties from thruster design data and those operating parameters which are directly controllable. By varying the operating parameter inputs to the computer code written to solve these equations, perfromance curves are obtained which agree quite well with measurements.

Longhurst, G. R.↗
NASA NTRS1979

Plasma processes in inert gas thrusters

Inert gas thrusters, particularly with large diameters, have continued to be of interest for space propulsion applications. Two plasma processes are treated in this study: electron diffusion across magnetic fields and double ion production in inert-gas thrusters. A model is developed to describe electron diffusion across a magnetic field that is driven by both density and potential gradients, with Bohm diffusion used to predict the diffusion rate. This model has applications to conduction across magnetic fields inside a discharge chamber, as well as through a magnetic baffle region used to isolate a hollow cathode from the main chamber. A theory for double ion production is presented, which is not as complete as the electron diffusion theory described, but it should be a useful tool for predicting double ion sputter erosion. Correlations are developed that may be used, without experimental data, to predict double ion densities for the design of new and especially larger ion thrusters.

Kaufman, H. R.↗
NASA NTRS1979

Electron diffusion through the baffle aperture of a hollow cathode thruster

The use of a hollow cathode in place of an oxide cathode to increase thruster operating lifetimes requires, among other things, the addition of a baffle to restrict the flow of electrons from the hollow cathode. A theoretical model is developed which relates the baffle aperture area of a hollow-cathode thruster to the magnetic flux density and plasma properties in the aperture region, with the result that this model could be used as an aid in thruster design. Extensive Langmuir probing is undertaken to verify the validity of the model and demonstrate its capability. It is shown that the model can be used to calculate the aperture area required to effect discharge operation at a specified discharge voltage and arc current.

Brophy, J. R.↗
NASA NTRS1979

Propellant injection for MPD thrusters

The propellant gas for a pulsed self-field MPD thruster must be supplied by a valve which will necessarily be separated from the thruster injectors by some ullage volume. A model and experimental data are presented showing the effect of the ullage volume on the propellant pulse shape, along with a discussion of the proper location for the propellant flow choke point(s). Prospective MPD pulse valves are surveyed and the performance of two solenoid valves, one with a pulse length less than 2 milliseconds, is included. Pulse shape data is presented for a thruster specially modified to reduce the ullage volume.

Jones, R. M.↗
NASA NTRS1979

Reduced power processor requirements for the 30-cm diameter HG ion thruster

The characteristics of power processors strongly impact the overall performance and cost of electric propulsion systems. A program was initiated to evaluate simplifications of the thruster-power processor interface requirements. The power processor requirements are mission dependent with major differences arising for those missions which require a nearly constant thruster operating point (typical of geocentric and some inbound planetary missions) and those requiring operation over a large range of input power (such as outbound planetary missions). This paper describes the results of tests which have indicated that as many as seven of the twelve power supplies may be eliminated from the present Functional Model Power Processor used with 30-cm diameter Hg ion thrusters.

Rawlin, V. K.↗
NASA NTRS1979

MPD thruster operation with inductively stored energy

Magnetoplasmadynamic (MPD) thrusters with 10-cm and 2.5-cm long thoriated tungsten cathodes have been operated with both inductive and capacitive energy sources at currents up to 8 kA. Voltage-current characteristics and magnetic field measurements show that over most of the operating range, the inductively driven thruster acts as a continuously varying quasi-steady accelerator at the same current. However, at values of a specified parameter near the onset condition for voltage fluctuations and electrode and insulator erosion, the voltage for the inductively driven thruster is higher than its quasi-steady counterpart. This higher voltage occurs because the equilibration time constant of the MPD discharge, is too large to allow the terminal voltage to attain a quasi-steady value as the inductively driven current decreases.

Clark, K. E.↗
NASA NTRS1979

A model for predicting the wearout lifetime of the LeRC/Hughes 30-cm mercury ion thruster

An investigation of parameters that affect the erosion rates of 30-cm-diameter mercury-ion-thruster components is described. A sputter-erosion model is formulated in terms of the design, operational, and material characteristics of the thruster. The erosion model is applied to the screen electrode, which is assumed to be the life-limiting component of the 30-cm thruster, resulting in a model of wearout lifetime. Results of short-term erosion-rate tests are presented that illustrate the dependence of component wear rates on variables such as discharge voltage, accelerator-grid open-area fraction, ion energy, electrode material, and the partial pressure of facility residual gases such as nitrogen. Test results are compared with wearout rates predicted by the sputter-erosion model.

Beattie, J. R.↗
NASA NTRS1980

Ion thruster plume effects on spacecraft surfaces

A charge-exchange plasma, generated by an ion thruster, is capable of flowing upstream from the ion thruster and therefore represents a source of contamination to a spacecraft. An analytical model of the charge-exchange plasma density around a spacecraft is used to estimate the contamination which various spacecraft materials may be exposed to. Measurements of plasma density around an ion thruster are compared to this model. Results of experimental studies regarding the effects on various spacecraft materials' properties due to exposure to expected mercury contamination levels are presented.

Carruth, M. R., Jr.↗
NASA NTRS1981

Ion thruster plume effects on spacecraft surfaces

A charge-exchange plasma, generated by an ion thruster, is capable of flowing upstream from the ion thruster and therefore represents a source of contamination to a spacecraft. An analytical model of the charge-exchange plasma density around a spacecraft was used to estimate the contamination which various spacecraft materials may be exposed to. Measurements of plasma density around an ion thruster were compared to this model. Results of experimental studied regarding the effects on various spacecraft materials' properties due to exposure to expected mercury contamination levels are presented.

Carruth, M. R., Jr.↗
NASA NTRS1981

Performance of a magnetic multipole line-cusp argon ion thruster

A 17 cm diameter line cusp ion thruster was evaluated with inert gases which are candidate propellants for on orbit and orbit transfer propulsion functions for Large Space Systems. A semiempirical relationship was generated to predict thruster beam current in terms of plasma parameters which would allow initial thruster optimization without ion extraction and the associated large vacuum facilities. The sensitivity of performance to changes in discharge electrode configurations and magnetic circuit was evaluated and is presented. After final optimization a propellant utilization efficiency of 0.9 at a discharge chamber power expenditure of about 260 w per beam ampere was obtained. These performance parameters are the highest yet achieved with argon propellant.

Sovey, J. S.↗
NASA NTRS1981

SERT II 1980 extended flight thruster experiments

The flight results obtained from mid 1979 through December 1980 are presented. Near continuous solar power in 1979 and 1980 has enabled long periods of thruster endurance testing. Three of four propellant tanks were exhausted with no significant change in thruster system operation before being empty. A new plasma mode thrust was characterized and direct thrust measurements obtained. Other tests, including beam neutralization by various neutralizer sources, give insight to electron conduction across plasmas in space and provide a basis to model neutralization of thruster arrays.

Kerslake, W. R.↗
NASA NTRS1981

Characteristics of 30-centimeter mercury ion thrusters

The technology development of the 30 centimeter J series mercury ion thruster for prime propulsion application in solar electric propulsion systems is described. Thruster design is reviewed. A standardized set of test and data recording procedures formulated to allow for the characterization of the J series thruster is described. Characteristics measured are the magnetic baffle characterization, the neutralizer characterization, perveance, the minimum eV/ion measurement, and the electrical and propellant utilization efficiency measurements. Test results are presented.

Maloy, J. E.↗
NASA NTRS1980

Inert gas thrusters

Some advances in component technology for inert gas thrusters are described. The maximum electron emission of a hollow cathode with Ar was increased 60-70% by the use of an enclosed keeper configuration. Operation with Ar, but without emissive oxide, was also obtained. A 30 cm thruster operated with Ar at moderate discharge voltages give double-ion measurements consistent with a double ion correlation developed previously using 15 cm thruster data. An attempt was made to reduce discharge losses by biasing anodes positive of the discharge plasma. The reason this attempt was unsuccessful is not yet clear. The performance of a single-grid ion-optics configuration was evaluated. The ion impingement on the single grid accelerator was found to approach the value expected from the projected blockage when the sheath thickness next to the accelerator was 2-3 times the aperture diameter.

Kaufman, H. R.↗
NASA NTRS1981

Additional application of the NASCAP code. Volume 2: SEPS, ion thruster neutralization and electrostatic antenna model

The interactions of spacecraft systems with the surrounding plasma environment were studied analytically for three cases of current interest: calculating the impact of spacecraft generated plasmas on the main power system of a baseline solar electric propulsion stage (SEPS), modeling the physics of the neutralization of an ion thruster beam by a plasma bridge, and examining the physical and electrical effects of orbital ambient plasmas on the operation of an electrostatically controlled membrane mirror. In order to perform these studies, the NASA charging analyzer program (NASCAP) was used as well as several other computer models and analytical estimates. The main result of the SEPS study was to show how charge exchange ion expansion can create a conducting channel between the thrusters and the solar arrays. A fluid-like model was able to predict plasma potentials and temperatures measured near the main beam of an ion thruster and in the vicinity of a hollow cathode neutralizer. Power losses due to plasma currents were shown to be substantial for several proposed electrostatic antenna designs.

Katz, I.↗
NASA NTRS1981

SERT II 1980 extended flight thruster experiments

The SERT II spacecraft, launched in 1970, has been maintained in an operational, but intermittent status since 1971. This paper presents the flight results obtained from mid 1979 through December 1980. Near continuous solar power in 1979 and 1980 has enabled long periods of thruster endurance testing. Three of four propellant tanks have been exhausted with no significant change in thruster system operation before being empty. A new plasma mode thrust has been characterized and direct thrust measurements obtained. Other tests, including beam neutralization by various neutralizer sources, give insight to electron conduction across plasmas in space and provide a basis to model neutralization of thruster arrays.

Kerslake, W. R.↗
NASA NTRS1981

Thrust and efficiency of a self-field MPD thruster

Thrust and efficiency of a quasi-steady multi-MW argon MPD thruster are determined for one-millisecond current pulses. Terminal voltage and impulse bit per pulse are measured for a benchmark thruster geometry on a swinging gate thrust stand in a dielectric vacuum tank, at a background pressure of 0.0001 torr, for a range of argon flow rates and arc currents. The quasi-steady thrust data scale quadratically with arc current, and confirm previous estimates of the electromagnetic and electrothermal components of thrust from magnetic and pressure probe measurements. Thruster efficiency is found to increase monotonically with specific impulse, reaching a value of 25% for 4.5 to 6.0 g/sec at 2000 seconds. Results of further experiments show that the inferred specific impulse for voltage fluctuation onset can be increased to 3000 seconds, and the thrust efficiency to above 30%.

Burton, R. L.↗
NASA NTRS1981

Performance of large inert-gas thrusters

The performance of large inert-gas thrusters is predicted based on present knowledge of ion optics performance and discharge chamber operation. Calculated performance data are given for argon and xenon propellants. The effect of varying propellant utilization and thruster diameter is discussed and the optimum choice of beam diameter for very large systems is indicated for low, intermediate, and high specific impulses. Optimum discharge chamber depths are also specified. Although detailed design considerations may modify the predictions, the general trends indicated should still be useful for directing future technology efforts and evaluating mission studies involving large thrusters.

Kaufman, H. R.↗