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At least 991 records · Page 55

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NASA NTRS1984

Demonstration of a new electrothermal thruster concept

The design and test of a microwave electrothermal thruster are described. The device, which employs a coaxial microwave discharge, was tested in nitrogen gas with 200-600 W of 2.45-GHz input power. Experimental measurements of thrust, specific impulse, and energy efficiency are presented for different flow and discharge pressures. Measured energy efficiencies varied between 30-60 percent and the performance compared favorably with other electrothermal thrusters operating in nitrogen gas. The experimental performance demonstrated the feasibility of the concept.

Whitehair, S.↗

NASA NTRS1984

Proposed system design for a 20 kW pulsed electrothermal thruster

A conceptual design is presented for a Pulsed Electrothermal (PET) propulsion system for the Air Force Space Based Radar satellite, which has a mass of 7000 kg. The proposed system boosts the SBR satellite from 150 n.m. to 600 n.m. with a 4 deg plane change, for a total mission Delta v of 1 km/sec. Satellite power available is 50 kW, and 45 kW are used to drive two water-injected 20 kW PET thrusters, delivering 5.6 N thrust to the SBR at 1000 seconds specific impulse. The predicted mission trip time is 15 days. The proposed system consumes 850 kg of water propellant, stored in a central tank and injected with pressurized helium. Component mass estimates based on space-qualified hardware are presented for the propellant handling, power conditioning and thruster subsystems. The estimated total mass is 400 kg and the propulsion system specific mass is alpha = 10 kg/kW. The proposed system efficiency of 0.62 at 1000 seconds specific impulse is supported by experimental performance measurements.

Burton, R. L.↗

NASA NTRS1983

Extended performance technology study 30-cm thruster

The extended performance technology study was an investigation of advanced discharge chambers and thruster components that were designed to operate under conditions which result in an increase in the thrust and thrust to power ratio of the state of the art J-series thruster. The high level of performance was achieved by a discharge chamber that employs a ring cusp magnetic confinement arrangement and a three grid ion extraction assembly. It is shown that the ring cusp magnetic field geometry confines the plasma to the volume immediately adjacent to the ion extraction assembly. A high emission current hollow cathode that demonstrated operation at an emission current as high as J sub E = 40 A, and measurements which show the breakdown voltage of individual sections of the J-series propellant flow electrical isolator is about 340 V per section are investigated.

Beattie, J. R.↗

NASA NTRS1984

A high performance magnetoplasmadynamic thruster

A flared-anode MPD thruster has been modified to permit injection of propellant through the backplate near the anode wall. At 6 g/sec argon, this thruster displays an onset current of 41.4 kA, almost double the value observed for propellant injection at the cathode and intermediate radial positions. A magnetic field survey of the interelectrode region shows current density is highest at the upstream and downstream ends of the chamber. The operating efficiency at onset current inferred from magnetic field data exceeds 50 percent, but swinging-gate thrust stand measurements reveal a progressive divergence between inferred and actual thrust with increasing power. Near onset, the measured thrust is approximately 20 percent lower than that inferred from magnetic probing. Explanations for this behavior have been explored with viscous drag emerging as the most probable cause of performance degradation.

Wolff, M.↗

NASA NTRS1984

MPD thruster operation with stepped current input

Experimental study of the response of a quasi-steady MPD thruster to an incremental increase or decrease in current illuminates the processes by which the discharge adjusts to changes in its operating level. For changes well below the onset condition, the thruster maintains quasi-steady operation throughout. Changes imposed near onset produce transition periods of non-quasi-steady operation. The duration of the transient appears to be determined by magnetic field diffusion in the portion of the discharge downstream of the thrust chamber.

Jahn, R. G.↗

NASA NTRS1984

Electrothermal thruster diagnostics

A flight-qualified electrothermal thruster demonstrated its adaptability to a variety of propellants. Originally qualified for operation with hydrazine propellant, it was operated with nitrogen, hydrogen, and ammonia propellants, demonstrating 73, 61, and 52 percent overall efficiency with these propellants, respectively, when tested over a wide range of operating conditions. By introducing a preheater to admit hot, rather than cold, propellant inlet gases to the thruster's augmentation heat exchanger, delivered specific impulse closer to theoretical performance limits should be achieved.

Zafran, S.↗

NASA NTRS1984

Experiments with a microwave electrothermal thruster concept

The design and test of two microwave electrothermal thrusters is described. The coaxial and cylindrical cavity concepts are tested in nitrogen gas with 200-2,000 W of 2.45 GHz input power. Experimental measurements and calculations of thrust, specific impulse and energy efficiency are presented for different gas flow rates and discharge pressures. Measured energy efficiencies varied between 10-60 percent and the performance compared favorably with other electrothermal thrusters operating in nitrogen gas. The experimental performance demonstrated the feasibility of the concept.

Asmussen, J.↗

NASA NTRS1985

Exact spacecraft maneuvers with jerk-free transition from thrusters to momentum wheels

A combined detumbling maneuver with constant but vectored external thrusters is presently followed, as in Dwyer and Batten (1985) by an exact reorientation that employs internal momentum transfer. Unlike the Dwyer and Batten technique, however, the proposed reorientation maneuver is of such a trajectory that spacecraft angular acceleration is preserved during the transition from thrusters to reaction wheels. This prevents jerk-driven excitation of unmodeled elastic modes.

Dwyer, T. A. W., III↗

NASA NTRS1985

MPD thruster performance - Propellant distribution and species effects

Previous research has shown that MPD thruster performance, as measured by thrust, specific impulse, and efficiency, is limited by a so-called 'onset' current level. Experimental results are presented which show that the onset current level varies with the radial position of mass injection. Specifically, mass injection into either of two zones results in drastically different performance for both argon and xenon propellants. This difference in performance is evidenced by the onset current levels, the onset V-J signatures, and the overall V-J characteristics. The behavior of the thruster with the total mass flow split between the two zones follows a transition between the zonal behavioral limits. The behavior of propellant mixtures also follows a transition between the behavior of the individual propellant species. The flow region between the two zones is characterized by a particle number density discontinuity, a large axial current density, and a large electron Hall parameter, suggesting a possible correlation between observed onset phenomena and a shear-flow discontinuity located in the region between the two zones.

Kelly, A. J.↗

NASA NTRS1985

Onset phenomena in MPD thrusters

An experimental study has clarified some aspects of MPD thruster onset phenomena. The steep increase in terminal voltage that occurs as the onset current is approached may have different causes, depending on the propellant injection geometry. For propellant injection at the cathode radius, terminal voltage increase corresponds to a growing anode fall voltage; for injection at a larger radius, the increase is related to the back emf in the near-cathode plasma. The formation of the onset current pattern within the arc has been mapped experimentally as the thruster responds to an input current step which rises from below onset to the onset value. The appearance of terminal voltage hash at onset correlates with the extension into the exhaust region of a significant fraction of the arc current.

Barnett, J. W.↗

NASA NTRS1986

Dual-throat thruster thermal model

The dual-throat engine is one of the dual nozzle engine concepts studied for advanced space transportation applications. It provides a thrust change and an in-flight area ratio change through the use of two concentric combustors with their throats arranged in series. Test results are presented for a dual throat thruster burning gaseous oxygen and hydrogen at primary (inner) chamber pressures from 380 to 680 psia. Heat flux profiles were obtained from calorimetric cooling channels in the inner nozzle, outer or secondary chamber and the tip of the inner nozzle. Data were obtained for two nozzle spacings over a chamber pressure ratio (secondary/primary) range of 0.45 to 0.83 with both chambers firing (Mode I). Fluxes near the end of the inner nozzle were significantly higher than in Mode II when only the inner chamber was fired, due to the flow separation and recirculation caused by the back pressure imposed by the secondary chamber. As the pressure ratio increased, these heat fluxes increased and the region of high heat flux relative to Mode II extended farther upstream. The use of the gaseous hydrogen bleed flow in the secondary chamber to control heat fluxes in the primary plume attachment region was investigated in Mode II testing. A thermal model of a dual throat thruster was developed and upgraded using the experimental data.

Ewen, R. L.↗

NASA NTRS1987

A theoretical investigation of laser-sustained plasma thruster

A numerical code has been successfully developed for the investigation of thruster performance using a laser-sustained hydrogen plasma as the propellant. The plasma was sustained using a 10.6-micron CO2 laser beam focused at different positions within the thruster. The physical model assumed that plasma is in thermodynamical equilibrium (LTE), and geometric ray tracing was adopted to describe the laser beam. The steady-state, axisymmetric, Navier-Stokes equations coupled with the laser power absorption process have been solved numerically. A pressure based Navier-Stokes numerical solver using body-fitted coordinates was used to calculate the laser-supported rocket flow which includes both subsonic and supersonic flow regions. From the limited parametric study, which did not try to optimize the rocket performance, it was found that better performance was obtained when the laser beam was focused closer to the rocket throat.

Jeng, San-Mou↗

NASA NTRS1987

Microwave electrothermal thruster performance in helium gas

The microwave electrothermal thruster presented uses an internally tuned, single-mode cylindrical cavity applicator to focus and match microwave energy into an electrodeless, high pressure flowing gas discharge that is located within a quartz discharge chamber. Experimental measurements of microwave coupling efficiency, thruster energy efficiency, and specific impulse, are obtained for N and He discharges; the efficiency of microwave energy transfer to the discharge is found to be of the order of 95 percent. Higher temperature nozzle materials and more efficient discharge chambers will further enhance performance.

Whitehair, S.↗

NASA NTRS1987

Experiments on a repetitively pulsed electrothermal thruster

This paper presents experimental results from an investigation of a pulsed electrothermal (PET) thruster using water propellant. The PET thruster is operated on a calibrated thrust stand, and produces a thrust to power ratio of T/P = 0.07 + or - 0.01 N/kW. The discharge conditions are inferred from a numerical model which predicts pressure and temperature levels of 300-500 atm and 20,000 K, respectively. These values in turn correctly predict the measured values of impulse bit and discharge resistance. The inferred ideal exhaust velocity from these conditions is 17 km/sec, but the injection of water propellant produces a test tank background pressure of 10-20 Torr, which reduces the exhaust velocity to 14 km/sec. This value corresponds to a thrust efficiency of 54 + or - 7 percent when all experimental errors are taken into account.

Burton, R. L.↗

NASA NTRS1987

MPD thruster plasma instability studies

Results of an ongoing experimental/theoretical effort to understand the influence of plasma waves and instabilities upon coaxial plasma thruster performance and longevity are discussed. An experiment which permits measurement of the plasma dispersion relation plus a suite of plasma parameters for each one millisecond quasi-steady discharge is described. This investigation is guided by a linearized Boltzmann-Poisson kinetic numerical model which can accommodate the effects of collisions, inhomogeneities and the magnetic field. Preliminary tests reveal the presence of the low-frequency branch of an ion cyclotron wave. Although the stability of these waves still has to be established through spatial growth rate measurements, it is well known that they are prone to 'current-driven' instabilities. 'Current-driven' instabilities in the collisional tensor conducting plasma of the MPD thruster are analytically discussed and their relevance to the MPD plasma is illustrated by estimating a critical total current at which their onset occurs.

Kelly, A. J.↗

NASA NTRS1988

The effects of magnetic nozzle configurations on plasma thrusters

Magnetoplasmadynamics (MPD) arc devices have been operated at power levels from 10 KW to 0.1 MW. When these devices have magnetic fields applied to them, they show marked increases in thrust in direct proportion to the magnitide of the applied field. Electrode erosion may be influenced by applied fields. This proposal will study the application of variable magnetic fields over a range of thruster powers, gas densities, and thruster configurations. It is proposed to examine this behavior with numerical codes and limited but relevant experimental tests.

York, Thomas M.↗

NASA NTRS1988

Preliminary design study of hydrogen and ammonia resistojets for prime and auxiliary thrusters

Designs of high performance resistojets for primary and auxiliary propulsion are described.Thruster power for the primary propulsion application was in the 2 to 3 kW range while auxiliary propulsion power per thruster was 0.15 to 0.25 kW. Propellants considered were hydrogen and ammonia. The report described design techniques used to forecast the temperature and energy flux distributions using mathematical modeling by personal microcomputer. BASIC language is used throughout to give the designer rapid interaction and control. Both designs integrate compact first stage coils with concentric tubular heaters. The hybrid heater design allows better thruster power matching with the spacecraft power bus. Projected specific impulse levels were 760 to 830 s for hydrogen and 380 to 410 s for ammonia.

Page, Russell J.↗

NASA NTRS1989

Development of a high power microwave thruster, with a magnetic nozzle, for space applications

This paper describes the current development of a high-power microwave electrothermal thruster (MET) concept at the NASA Lewis Research Center. Such a thruster would be employed in space for applications such as orbit raining, orbit maneuvering, station change, and possibly trans-lunar or trans-planetary propulsion of spacecraft. The MET concept employs low frequency continuous wave (CW) microwave power to create and continuously pump energy into a flowing propellant gas at relative high pressure via a plasma discharge. The propellant is heated to very high bulk temperatures while passing through the plasma discharge region and then is expanded through a throat-nozzle assembly to produce thrust, as in a conventional rocket engine. Apparatus, which is described, is being assembled at NASA Lewis to test the MET concept to CW power levels of 30 kW at a frequency of 915 MHz. The microwave energy is applied in a resonant cavity applicator and is absorbed by a plasma discharge in the flowing propellant. The ignited plasma acts as a lossy load, and with optimal tuning, energy absorption efficiencies over 95 percent (based on the applied microwave power) are expected. Nitrogen, helium, and hydrogen will be tested as propellants in the MET, at discharge chamber pressures to 10 atm.

Power, John L.↗