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At least 217 records · Page 12

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NASA NTRS1984

Ion thruster performance model

A model of ion thruster performance is developed for high flux density, cusped magnetic field thruster designs. This model is formulated in terms of the average energy required to produce an ion in the discharge chamber plasma and the fraction of these ions that are extracted to form the beam. The direct loss of high energy (primary) electrons from the plasma to the anode is shown to have a major effect on thruster performance. The model provides simple algebraic equations enabling one to calculate the beam ion energy cost, the average discharge chamber plasma ion energy cost, the primary electron density, the primary-to-Maxwellian electron density ratio and the Maxwellian electron temperature. Experiments indicate that the model correctly predicts the variation in plasma ion energy cost for changes in propellant gas (Ar, Kr and Xe), grid transparency to neutral atoms, beam extraction area, discharge voltage, and discharge chamber wall temperature. The model and experiments indicate that thruster performance may be described in terms of only four thruster configuration dependent parameters and two operating parameters. The model also suggests that improved performance should be exhibited by thruster designs which extract a large fraction of the ions produced in the discharge chamber, which have good primary electron and neutral atom containment and which operate at high propellant flow rates.

Brophy, J. R.↗

NASA NTRS1987

Low power arcjet thruster pulse ignition

An investigation of the pulse ignition characteristics of a 1 kW class arcjet using an inductive energy storage pulse generator with a pulse width modulated power converter identified several thruster and pulse generator parameters that influence breakdown voltage including pulse generator rate of voltage rise. This work was conducted with an arcjet tested on hydrogen-nitrogen gas mixtures to simulate fully decomposed hydrazine. Over all ranges of thruster and pulser parameters investigated, the mean breakdown voltages varied from 1.4 to 2.7 kV. Ignition tests at elevated thruster temperatures under certain conditions revealed occasional breakdowns to thruster voltages higher than the power converter output voltage. These post breakdown discharges sometimes failed to transition to the lower voltage arc discharge mode and the thruster would not ignite. Under the same conditions, a transition to the arc mode would occur for a subsequent pulse and the thruster would ignite. An automated 11 600 cycle starting and transition to steady state test demonstrated ignition on the first pulse and required application of a second pulse only two times to initiate breakdown.

Sarmiento, Charles J.↗

NASA NTRS1986

Investigation of beamed-energy ERH thruster performance

The objective of this study was to determine the performance of an External Radiation Heated (ERH) thruster. In this thruster, high intensity laser energy is focused to ignite either a Laser Supported Combustion (LSC) wave or a Laser Supported Detonation (LSD) wave. Thrust is generated as the LSC or LSD wave propagates over the thruster's surface, or in the proposed thruster configuration, the vehicle afterbody. Thrust models for the LSC and LSD waves were developed and simulated on a computer. Performance parameters investigated include the effect of laser intensity, flight Mach number, and altitude on mean-thrust and coupling coefficient of the ERH thruster. Results from these models suggest that the ERH thruster using LSC/LSD wave ignition could provide propulsion performance considerably greater than any propulsion system currently available.

Myrabo, Leik N.↗

NASA NTRS1987

A theoretical evaluation of laser-sustained plasma thruster performance

An extensive numerical experiment has been conducted to evaluate rocket thruster performance using a laser-sustained hydrogen plasma as the propellant. The plasma was sustained using a 30 kW CO2 laser beam operated at 10.6 microns focused inside the thruster. The steady-state Navier-Stokes equations coupled with the laser power absorption process have been solved numerically. A pressure based Navier-Stokes solver using body-fitted coordinate was used to calculate the laser-supported rocket flow which included both recirculating and transonic flow regions. The local thermodynamic equilibrium (LTE) assumption was used for the plasma thermophysical and optical properties. Geometric ray tracing was adopted to describe the laser beam. Several different throat size thrusters operated at 150 and 300 kPa chamber stagnation pressure were studied. It was found that the thruster performance (vacuum specific impulse) was highly dependent on the operating conditions, and a properly designed laser supported thruster can attain a specific impulse around 1500 secs. The heat loading on the thruster wall was also estimated and was in the range of that for a conventional chemical rocket.

Jeng, San-Mou↗

NASA NTRS1987

Low power arcjet thruster pulse ignition

Sarmiento, Charles J.↗

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

Sovey, James S.↗

NASA NTRS1990

High-power xenon ion thrusters

Steady state performance characteristics for a 25-cm diameter laboratory-model xenon ion thruster operating over the power range from 1.8 to 6.8 kW are described. Performance data are also presented for a similar 30-cm diameter thruster operated over the 1.4 to 11.1 kW power range. For each of the two thrusters, thermal characteristics, discharge-chamber performance data, overall thruster performance data, and beam-current density profiles are presented. A high level of performance and stability over a wide range of input power is demonstrated for the two thrusters. It is concluded that there appears to be no technological obstacles in the way of developing flight-design thrusters that operate in the 5- to 10-kW power range for near-term applications to orbit raising.

Beattie, J. R.↗

NASA NTRS1990

Optimal thruster configurations for the GP-B spacecraft

Thruster locations and control schemes which make the most efficient use of the available propellant of a constant flow propulsion system are investigated. A set of three necessary and sufficient conditions are established for a thruster system to be optimal in the sense of maximizing the smallest magnitude force which can be generated by the thrusters. Analytical expressions are derived for the 'the envelope of least authority' which is defined as the smallest force magnitude versus moment magnitude which can be generated by any given thruster system. Several thruster configurations are proposed for the GP-B spacecraft and their performance is compared with respect to their envelopes of least authority. The results are applicable to any system composed of thrusters which saturate at some finite value.

Wiktor, P.↗

NASA NTRS1991

Los Alamos NEP research in advanced plasma thrusters

Research was initiated in advanced plasma thrusters that capitalizes on lab capabilities in plasma science and technology. The goal of the program was to examine the scaling issues of magnetoplasmadynamic (MPD) thruster performance in support of NASA's MPD thruster development program. The objective was to address multi-megawatt, large scale, quasi-steady state MPD thruster performance. Results to date include a new quasi-steady state operating regime which was obtained at space exploration initiative relevant power levels, that enables direct coaxial gun-MPD comparisons of thruster physics and performance. The radiative losses are neglible. Operation with an applied axial magnetic field shows the same operational stability and exhaust plume uniformity benefits seen in MPD thrusters. Observed gun impedance is in close agreement with the magnetic Bernoulli model predictions. Spatial and temporal measurements of magnetic field, electric field, plasma density, electron temperature, and ion/neutral energy distribution are underway. Model applications to advanced mission logistics are also underway.

Schoenberg, Kurt↗

NASA NTRS1992

Low-Isp derated ion thruster operation

The performance and lifetime expectations of 30 cm xenon ion thruster technology at low values of specific impulse were evaluated, with emphasis on 1000-2500 s operation. Power levels of up to 2.0 kW, appropriate for auxiliary and orbit maneuvering propulsion, were processed at thrust-to-power ratios up to 57 mN/kW. These tests were conducted using a derated 30 cm ion thruster with high-perveance design two-grid ion optics with xenon propellent. Lifetime projections were made based on a simple analysis of critical component erosion rates, and it was found that a strong correlation exists with the ratio of the specific impulse-to-input power. Under all operating conditions for which the projected thruster lifetime is less than 10,000 hrs, the life-limiting component of this technology is erosion of the accelerator grid due to charge-exchange ions. The use of alternative grid materials such as carbon is estimated to increase useful thruster lifetimes by as much as an order of magnitude and may enable long-life high thrust-density, sub-2500 s Isp operation. The performance and life of the derated thruster appears similar to that of the Russian SPT-100 thruster in the 1.0-2.0 kW, 1600-2000 s operational envelope.

Patterson, Michael J.↗

NASA NTRS1992

The evolutionary development of high specific impulse electric thruster technology

Electric propulsion flight and technology demonstrations conducted in the USA, Europe, Japan, China, and USSR are reviewed with reference to the major flight qualified electric propulsion systems. These include resistojets, ion thrusters, ablative pulsed plasma thrusters, stationary plasma thrusters, pulsed magnetoplasmic thrusters, and arcjets. Evolutionary mission applications are presented for high specific impulse electric thruster systems. The current status of arcjet, ion, and magnetoplasmadynamic thrusters and their associated power processor technologies are summarized.

Sovey, James S.↗

NASA NTRS1992

Stability test and analysis of the Space Shuttle Primary Reaction Control Subsystem thruster

The results are reported of a test program conducted on the Space Shuttle Primary Reaction Control Subsystem thruster in order to investigate the effects of trapped helium bubbles and saturated propellants on stability, determine if thruster-to-thruster stability variations are significant, and determine stability under STS-representative conditions. It is concluded that the thruster design is highly reliable in flight and that burn-through has not occurred. Significantly unstable thrusters are screened out, and wire wrap is found to protect against chamber burn-throughs and to provide a fail-safe thruster for this situation.

Applewhite, John↗

NASA NTRS1992

Scaling of 100 kW class applied-field MPD thrusters

Three cylindrical applied-field magnetoplasmadynamic thrusters were tested with argon propellant over a broad range of operating conditions to establish empirical scaling laws for thruster performance. Argon flow rates, discharge currents, and applied-field strengths were varied between 0.025 and 0.14 g/s, 750 to 2000 A, and 0.034 to 0.20T, respectively. The results showed that the thrust reached over five times the self-field value, and that thrust increased linearly with the product of discharge current and applied-field strength and quadratically with the anode radius. While increasing the propellant flow rate increased the thrust, it did not affect the rate of thrust increase with applied-field strength, and at low propellant flow rates the self-field thrust approached 30 percent of the measured thrust. The voltage increased linearly with applied-field strength but was insensitive to the discharge current. The rate of voltage increase with applied-field strength dependent on anode radius. Thruster efficiency increased monotonically with applied-field strength and propellant flow rate. Peak thruster efficiencies were insensitive to changes in anode radius. Electrode power loss and thruster efficiency measurements showed that while the electrode losses decreased with increasing anode radius the plasma losses increased. The opposite behaviors of electrode and plasma losses demonstrates the need to identify ways of independently controlling the thruster loss mechanisms.

Myers, Roger M.↗

NASA NTRS1992

Low-Isp derated ion thruster operation

The performance and lifetime expectations of 30 cm xenon ion thruster technology at low values of specific impulse were evaluated, with emphasis on 1000-2500 s operation. Power levels of up to 2.0 kW, appropriate for auxiliary and orbit maneuvering propulsion, were processed at thrust-to-power ratios up to 57 mN/kW. These tests were conducted using a derated 30 cm ion thruster with high-perveance design two-grid ion optics with xenon propellant. Lifetime projections were made based on a simple analysis of critical component erosion rates, and it was found that a strong correlation exists with the ratio of the specific impulse-to-input power. Under all operating conditions for which the projected thruster lifetime is less than 10,000 hrs, the life-limiting component of this technology is erosion of the accelerator grid due to charge-exchange ions. The use of alternative grid materials such as carbon is estimated to increase useful thruster lifetimes by as much as an order of magnitude and may enable long-life high thrust-density, sub-2500 s Isp operation. The performance and life of the derated thruster appears similar to that of the Russian SPT-100 thruster in the 1.0-2.0 kW, 1600-2000 s operational envelope.

Patterson, Michael J.↗

NASA NTRS1994

Power Electronics Development for the SPT-100 Thruster

Russian electric propulsion technologies have recently become available on the world market. Of significant interest is the Stationary Plasma Thruster (SPT) which has a significant flight heritage in the former Soviet space program. The SPT has performance levels of up to 1600 seconds of specific impulse at a thrust efficiency of 0.50. Studies have shown that this level of performance is well suited for stationkeeping applications, and the SPT-100, with a 1.35 kW input power level, is presently being evaluated for use on Western commercial satellites. Under a program sponsored by the Innovative Science and Technology Division of the Ballistic Missile Defense Organization, a team of U.S. electric propulsion specialists observed the operation of the SPT-100 in Russia. Under this same program, power electronics were developed to operate the SPT-100 to characterize thruster performance and operation in the U.S. The power electronics consisted of a discharge, cathode heater, and pulse igniter power supplies to operate the thruster with manual flow control. A Russian designed matching network was incorporated in the discharge supply to ensure proper operation with the thruster. The cathode heater power supply and igniter were derived from ongoing development projects. No attempts were made to augment thruster electromagnet current in this effort. The power electronics successfully started and operated the SPT-100 thruster in performance tests at NASA Lewis, with minimal oscillations in the discharge current. The efficiency of the main discharge supply was measured at 0.92, and straightforward modifications were identified which could increase the efficiency to 0.94.

Hamley, John A.↗

NASA NTRS1994

An investigation of magnetic field effects on plume density and temperature profiles of an applied-field MPD thruster

Applied-field magnetoplasmadynamic (MPD) thruster performance is below levels required for primary propulsion missions. While MPD thruster performance has been found to increase with the magnitude of the applied-field strength, there is currently little understanding of the impact of applied-field shape on thruster performance. The results of a study in which a single applied-field thruster was operated using three solenoidal magnets with diameters of 12.7, 15.2, and 30.4-cm are presented. Thruster voltage and anode power deposition were measured for each applied field shape over a range of field strengths. Plume electron number density and temperature distributions were measured using a Langmuir probe in an effort to determine the effect of field shape on plume confinement by the diverging magnetic-field for each of the three magnetic field shapes. Results show that the dependence of the measured thruster characteristics on field shape were non-monotonic and that the field shape had a significant effect on the plume density and temperature profiles.

Bullock, S. Ray↗

NASA NTRS1995

NASA 30 Cm Ion Thruster Development Status

A 30 cm diameter xenon ion thruster is under development at NASA to provide an ion propulsion option for missions of national interest and it is an element of the NASA Solar Electric Propulsion Technology Applications Readiness (NSTAR) program established to validate ion propulsion for space flight applications. The thruster has been developed to an engineering model level and it incorporates innovations in design, materials, and fabrication techniques compared to those employed to conventional ion thrusters. The performance of both functional and engineering model thrusters has been assessed including thrust stand measurements, over an input power range of 0.5-2.3 kW. Attributes of the engineering model thruster include an overall mass of 6.4 kg, and an efficiency of 65 percent and thrust of 93 mN at 2.3 kW input power. This paper discusses the design, performance, and lifetime expectations of the functional and engineering model thrusters under development at NASA.

Patterson, Michael J.↗