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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 991 records · Page 55
All sourcesNASA NTRSDOE OSTI
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.↗
NASA NTRS1989

Structural and thermal response of 30 cm diameter ion thruster optics

Tabular and graphical data are presented which are intended for use in calibrating and validating structural and thermal models of ion thruster optics. A 30 cm diameter, two electrode, mercury ion thruster was operated using two different electrode assembly designs. With no beam extraction, the transient and steady state temperature profiles and center electrode gaps were measured for three discharge powers. The data showed that the electrode mount design had little effect on the temperatures, but significantly impacted the motion of the electrode center. Equilibrium electrode gaps increased with one design and decreased with the other. Equilibrium displacements in excess of 0.5 mm and gap changes of 0.08 mm were measured at 450 W discharge power. Variations in equilibrium gaps were also found among assemblies of the same design. The presented data illustrate the necessity for high fidelity ion optics models and development of experimental techniques to allow their validation.

Macrae, G. S.↗
NASA NTRS1988

An approach to the parametric design of ion thrusters

A methodology that can be used to determine which of several physical constraints can limit ion thruster power and thrust, under various design and operating conditions, is presented. The methodology is exercised to demonstrate typical limitations imposed by grid system span-to-gap ratio, intragrid electric field, discharge chamber power per unit beam area, screen grid lifetime, and accelerator grid lifetime constraints. Limitations on power and thrust for a thruster defined by typical discharge chamber and grid system parameters when it is operated at maximum thrust-to-power are discussed. It is pointed out that other operational objectives such as optimization of payload fraction or mission duration can be substituted for the thrust-to-power objective and that the methodology can be used as a tool for mission analysis.

Wilbur, Paul J.↗
NASA NTRS1989

Plume characteristics of MPD thrusters: A preliminary examination

A diagnostics facility for MPD thruster plume measurements was built and is currently undergoing testing. The facility includes electrostatic probes for electron temperature and density measurements, Hall probes for magnetic field and current distribution mapping, and an imaging system to establish the global distribution of plasma species. Preliminary results for MPD thrusters operated at power levels between 30 and 60 kW with solenoidal applied magnetic fields show that the electron density decreases exponentially from 1x10(2) to 2x10(18)/cu m over the first 30 cm of the expansion, while the electron temperature distribution is relatively uniform, decreasing from approximately 2.5 eV to 1.5 eV over the same distance. The radiant intensity of the ArII 4879 A line emission also decays exponentially. Current distribution measurements indicate that a significant fraction of the discharge current is blown into the plume region, and that its distribution depends on the magnitudes of both the discharge current and the applied magnetic field.

Myers, Roger M.↗
NASA NTRS1989

ECR plasma thruster research - Preliminary theory and experiments

A preliminary theory of the operation of the electron-cyclotron-resonance (ECR) plasma thruster is described along with an outline of recent experiments. This work is presented to communicate the status of an ongoing research effort directed at developing a unified theory to quantitatively describe the operation of the ECR plasma thruster. The theory is presented as a set of nonlinear ordinary differential equations and boundary conditions which describe the plasma density, velocity, and electron temperature. Diagnostic tools developed to measure plasma conditions in the existing research device are described.

Sercel, Joel C.↗
NASA NTRS1989

Structural and thermal response of 30 cm diameter ion thruster optics

This paper presents tabular and graphical data intended for use in calibrating and validating structural and thermal models of ion thruster optics. A 30-cm-diameter, two electrode, mercury ion thruster was operated using two different electrode assembly designs. With no beam extraction, the transient and steady state temperature profiles and center electrode gaps were measured for three discharge powers. The data showed that the electrode mount design had little effect on the temperatures, but significantly impacted the motion of the electrode center. Equilibrium electrode gaps increased with one design and decreased with the other. Equilibrium displacements in excess of 0.5 mm and gap changes of 0.08 mm were measured at 450 W discharge power. Variations in equilibrium gaps were also found among assemblies of the same design. The presented data illustrate the necessity for high fidelity ion optics models and development of experimental techniques to allow their validation.

Macrae, G.↗
NASA NTRS1989

Numerical modeling of a vortex stabilized arcjet thruster

A numerical method to solve the equations governing a vortex stabilized arcjet thruster is being developed. The model will allow the effects of swirling flow and geometry on arcjet thruster performance to be determined. The propellant flow equations, i.e., the axisymmetric, thin layer, Navier-Stokes equations, are solved using a Gauss-Siedel line-relaxation procedure. An implicit FTCS method is used to solve the electromagnetic field equations. A grid generation scheme was developed for an arbitrary arcjet geometry. The model will allow the radial and axial components of velocity and current distributions to be determined from a region upstream of the cathode, through the constrictor, to the exit plane of the nozzle.

Keith, Theo G., Jr.↗
NASA NTRS1989

Plume characteristics of MPD thrusters - A preliminary examination

A diagnostics facility for MPD thruster plume measurements has been built and is currently undergoing testing. The facility includes electrostatic probes for electron temperature and density measurements, Hall probes for magnetic field and current distribution mapping, and an imaging system to establish the global distribution of plasma species. Preliminary results for MPD thrusters operated at power levels between 30 and 60 kW with solenoidal applied magnetic fields show that the electron density decreases exponentially from 1 x 10 to the 20th/cu m to 2 x 10 to the 18th/cu m over the first 30 cm of the expansion, while the electron temperature distribution is relatively uniform, decreasing from approximately 2.5 eV to 1.5 eV over the same distance. The radiant intensity of the ArII 4879 A line emission follows similar patterns. Current distribution measurements indicate that significant fraction of the discharge current is blown into the plume region, and that its distribution depends on the magnitudes of both the discharge current and the applied magnetic field.

Myers, R. M.↗
NASA NTRS1988

An approach to the parametric design of ion thrusters

A methodology that can be used to determine which of several physical constraints can limit ion thruster power and thrust, under various design and operating conditions, is presented. The methodology is exercised to demonstrate typical limitations imposed by grid system span-to-gap ratio, intragrid electric field, discharge chamber power per unit beam area, screen grid lifetime and accelerator grid lifetime constraints. Limitations on power and thrust for a thruster defined by typical discharge chamber and grid system parameters when it is operated at maximum thrust-to-power are discussed.

Wilbur, Paul J.↗
NASA NTRS1989

The effects of arcjet thruster operating condition constrictor geometry on the plasma plume

Measurements of plasma number density and electron temperature were obtained in the plumes of lab arcjet thrusters using electrostatic probes of both spherical and cylindrical geometry. The two arcjet thrusters used had different constrictor and/or nozzle geometries and operated on mixtures of nitrogen, hydrogen, and ammonia to simulate the decomposition products of hydrazine and ammonia. An increase in the measured electron density was observed for both geometries with increasing arc power at a constant mass flow rate and with increasing mass flow rate at a constant arc current. For a given operating condition, the electron number density decreased exponentially off centerline and followed an inverse distance squared relationship along the thrust axis. Typical measured electron temperatures ranged from 0.1 to 0.2 eV.

Carney, Lynnette M.↗
NASA NTRS1990

Geometric effects in applied-field MPD thrusters

Three applied-field magnetoplasmadynamic (MPD) thruster geometries were tested with argon propellant to establish the influence of electrode geometry on thruster performance. The thrust increased approximately linearly with anode radius, while the discharge and electrode fall voltages increased quadratically with anode radius. All these parameters increased linearly with applied-field strength. Thrust efficiency, on the other hand, was not significantly influenced by changes in geometry over the operating range studied, though both thrust and thermal efficiencies increased monotonically with applied field strength. The best performance, 1820 sec I(sub sp) at 20 percent efficiency, was obtained with the largest radius anode at the highest discharge current (1500 amps) and applied field strength (0.4 Tesla).

Myers, R. M.↗