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At least 955 records · Page 53
All sourcesNASA NTRSDOE OSTI
NASA NTRS1979

Pulsed plasma thruster backflow characteristics

The exhaust plume of a millipound pulsed plasma thruster has been investigated using a unique vacuum facility with cryogenically cooled low back-scatter chamber walls. The mass flux distribution of the plume was determined, with emphasis on the region upstream of the thruster. Measurements of the plume material scattered from the chamber walls were used with articulated collimator-quartz crystal microbalance measurements of the total flux to estimate the upstream flux originating in the plume. This flux was found to be of order 10 to the minus 10 g/sq cm/ pulse and to vary approximately with the inverse square of the radius. Measurements indicate the directed plume is essentially axisymmetric with a slight elongation in the direction parallel to the interelectrode gap. This downstream flux decreases with radius along a Gaussian profile with 90% of the plume confined to a 40 deg half-angle cone.

Rudolph, L. K.↗
NASA NTRS1979

A model for nitrogen chemisorption in ion thrusters

A theoretical model describing the formation of nitrogen species subject to chemisorption on ion thruster discharge chamber surfaces is presented. Molecules, atoms, atomic ions and molecular ions are identified as the important species in the analysis. Current densities of the atomic and molecular ions predicted by the model are compared to current densities measured in the beam of a SERT II thruster. The predicted and measured values of these two current densities are shown to agree within about + or - 100%. The mechanisms involved in the erosion of a surface subjected to simultaneous nitrogen chemisorption and sputter erosion by high energy ions are also discussed.

Wilbur, P. J.↗
NASA NTRS1979

Pulsed plasma thruster contamination studies

The exhaust plume of the one millipound pulsed plasma thruster has a measurable backflow upstream of the nozzle exit plane which may deposit on and degrade the performance of exposed spacecraft surfaces. High speed photographs and Faraday cup measurements suggest that this backflow is predominantly an electrically neutral, relatively low energy vapor. Articulated collimator quartz crystal microbalance measurements of this backflow were made for a thruster with a radically modified nozzle and a flat plate backflow shield, to determine the backflow sensitivity to nozzle design changes. The results are compared with the original nozzle backflow and show a measurable reduction in the backflow directly upstream of the shield.

Rudolph, L. K.↗
NASA NTRS1979

Interactions between a spacecraft and an ion thruster produced environment

A model of the charge-exchange plasma flow around a spacecraft is used to determine charge-exchange plasma densities and arrival rates at various locations on the spacecraft. A general assessment of particle deposition and effects of such are presented as they relate to the spacecraft and particular science instruments. Various aspects of the interactions between the ion thruster produced environment and a Comet Rendezvous (CR) spacecraft are addressed. No instruments will be affected adversely when the thrusters are off, excluding those that have passive radiators which may collect mercury.

Carruth, M. R., Jr.↗
NASA NTRS1981

Propagation of charge-exchange plasma produced by an ion thruster

Under the proper conditions there is an end-effect of a long, cylindrical Langmuir probe which allows a significant increase in collected ion current when the probe is aligned with a flowing plasma. This effect was used to determine the charge-exchange plasma flow direction at various locations relative to the ion thruster. The ion current collected by the probe as a function of its angle with respect to the plasma flow allows determination of the plasma density and plasma flow velocity at the probe's location upstream of the ion thruster optics. The density values obtained from the ion current agreed to within a factor of two of density values obtained by typical voltage-current Langmuir probe characteristics.

Carruth, M. R., Jr.↗
NASA NTRS1981

Interactions between a spacecraft and an ion thruster produced environment

The electron bombardment ion thruster is a candidate propulsion device for several proposed interplanetary missions. A comet rendezvous mission is expected to be the first use of a Solar Electric Propulsion System (SEPS). Because this is the first use of such a system, it is necessary to understand the interactions between the spacecraft and the environment produced by the SEPS. A preliminary assessment of the interactions between a thruster produced environment and the Comet Rendezvous spacecraft, including its science instruments, is presented which concludes that compatibility between the SEPS and the spacecraft can be obtained.

Carruth, M. R., Jr.↗
NASA NTRS1981

Recent work on an RF ion thruster

An experimental investigation of an rf ion thruster using an immersed coupler in an argon discharge is reported. The conical coil, used to couple rf power into the discharge, is placed inside the discharge vessel. The discharge was self-sustained by 100-150 MHz rf power at low environmental pressures. The ion extraction was accomplished by conventional accelerated grid optics from an unoptimized 8 cm diameter ion thruster.

Lee, R. Q.↗
NASA NTRS1981

MPD thruster application study

Developmental considerations for the magneto-plasma-dynamic (MPD) thruster are defined. General characteristics of an MPD engine are compared to those of chemical propulsion and ion bombardment engines and performance criteria which are mission specific are examined. Requirements for thruster ground testing facilities are discussed and the utilization of the space shuttle for an orbital flight test is addressed.

Source record↗
NASA NTRS1981

Energy storage systems for MPD thrusters

Because of its high thrust density, the magnetoplasmadynamic (MPD) thruster is a promising candidate for many advanced space missions. Its high power requirements lead to operation in a pulsed mode using an intermediate energy storage device. The characteristics of a system consisting of a solar array, energy storage capacitor, and MPD thruster are studied for array powers in the range 25-375 kw. Assuming simple analytic models for the circuit components, the circuit charge and discharge equations are solved numerically, resulting in the system efficiency and capacitance. The system efficiency is inversely proportional to array power and decreases with circuit resistance. Alternative methods of energy storage such as a pulse forming network and a homopolar generator, are presented, and an overall comparison between all of the methods is given.

Gabriel, S. B.↗
NASA NTRS1981

Magnetoelectrostatic thruster physical geometry tests

Inert gas tests are conducted with several magnetoelectrostatic containment discharge chamber geometries. The configurations tested include three discharge chamber lengths; three boundary magnet patterns; two different flux density magnet materials; hemispherical and conical shaped thrusters having different surface-to-volume ratios; and two and three grid ion optics. Argon mass utilizations of 60 to 79% are attained at 210 to 280 eV/ion in different test configurations. Short hemi thruster configurations are found to produce 70 to 92% xenon mass utilization at 185 to 220 eV/ion.

Ramsey, W. D.↗
NASA NTRS1981

High frequency plasma generators for ion thruster applications

Two concepts for high frequency discharge ion thrusters are described. Both sources are designed for use with 30 cm grid sets and argon propellant and utilize multi-cusp permanent magnet geometries for plasma confinement. The RF induction source is a conventional design representing a synthesis of the RIT and multi-cusp concepts. The preliminary data (without system optimization) indicate a discharge efficiency comparable to that obtained in 30 cm hollow cathode multi-cusp argon thrusters. The electron cyclotron heating source is electrodeless and exhibits plasma characteristics which should lead to greatly reduced discharge chamber and screen sputter rates with the optimization of the magnetic fields, microwave frequency, and feed configuration.

Divergilio, W. F.↗
NASA NTRS1981

Diagnostic system design for the Ion Auxiliary Propulsion System /IAPS/ - Flight test of two 8 cm mercury ion thrusters

The experimental design of a Diagnostic Subsystem (DSS) as part of an Ion Auxiliary Propulsion System (IAPS) to be flown on P80-1 spacecraft in May 1983, is discussed. The DSS is composed of several detectors measuring thruster efflux, material deposition and spacecraft potential relative to the local space plasma in the vicinity of two 8 cm mercury ion thrusters. The detectors consist of two QCM units measuring frequency in the range of two to 65 KHz. Nine solar cell arrays have the capability of measuring current and voltage from 0-600 mA and 0-0.9 V. Seven ion collectors can measure ion currents with bias voltages of 0, 25, 55 and 96 V. The potential probe can measure current at 16 different commandible levels varying from one to 5 K microamperes within a voltage range of -25 to 175 V. The analysis of the ground-based data indicates that the hardware is qualified for flight, with the detectors and electronic units having passed all functional and environmental tests. Block diagrams are given and the functional parameters of the different design configurations are described.

Hurst, E. B.↗
NASA NTRS1981

Retrofit and acceptance test of 30-cm ion thrusters

Six 30 cm mercury thrusters were modified to the J-series design and evaluated using standardized test procedures. The thruster performance meets the design objectives (lifetime objective requires verification), and documentation (drawings, etc.) for the design is completed and upgraded. The retrofit modifications are described and the test data for the modifications are presented and discussed.

Poeschel, R. L.↗
NASA NTRS1982

Fuel optimal maneuvers for spacecraft with fixed thrusters

Several mathematical models, including a minimum integral square criterion problem, were used for the qualitative investigation of fuel optimal maneuvers for spacecraft with fixed thrusters. The solutions consist of intervals of "full thrust" and "coast" indicating that thrusters do not need to be designed as "throttleable" for fuel optimal performance. For the primary model considered, singular solutions occur only if the optimal solution is "pure translation". "Time optimal" singular solutions can be found which consist of intervals of "coast" and "full thrust". The shape of the optimal fuel consumption curve as a function of flight time was found to depend on whether or not the initial state is in the region admitting singular solutions. Comparisons of fuel optimal maneuvers in deep space with those relative to a point in circular orbit indicate that qualitative differences in the solutions can occur. Computation of fuel consumption for certain "pure translation" cases indicates that considerable savings in fuel can result from the fuel optimal maneuvers.

Carter, T. C.↗
NASA NTRS1981

Advanced space propulsion thruster research

Experiments showed that stray magnetic fields can adversely affect the capacity of a hollow cathode neutralizer to couple to an ion beam. Magnetic field strength at the neutralizer cathode orifice is a crucial factor influencing the coupling voltage. The effects of electrostatic accelerator grid aperture diameters on the ion current extraction capabilities were examined experimentally to describe the divergence, deflection, and current extraction capabilities of grids with the screen and accelerator apertures displaced relative to one another. Experiments performed in orificed, mercury hollow cathodes support the model of field enhanced thermionic electron mission from cathode inserts. Tests supported the validity of a thermal model of the cathode insert. A theoretical justification of a Saha equation model relating cathode plasma properties is presented. Experiments suggest that ion loss rates to discharge chamber walls can be controlled. A series of new discharge chamber magnetic field configurations were generated in the flexible magnetic field thruster and their effect on performance was examined. A technique used in the thruster to measure ion currents to discharge chamber walls is described. Using these ion currents the fraction of ions produced that are extracted from the discharge chamber and the energy cost of plasma ions are computed.

Wilbur, P. J.↗
NASA NTRS1981

Electric thruster research

The multipole discharge chamber of an electrostatic ion thruster is discussed. No reductions in discharge losses were obtained, despite repeated demonstration of anode potentials more positive than the bulk of the discharge plasma. The penalty associated with biased anode operation was reduced as the magnetic integral above the biased anodes was increased. The hollow cathode is discussed. The experimental configuration of the Hall current thruster had a uniform field throughout the ion generation and acceleration regions. To obtain reliable ion generation, it was necessary to reduce the magnetic field strength, to the point where excessive electron backflow was required to establish ion acceleration. The theoretical study of ion acceleration with closed electron drift paths resulted in two classes of solutions. One class has the continuous potential variation in the acceleration region that is normally associated with a Hall current accelerator. The other class has an almost discontinuous potential step near the anode end of the acceleration region. This step includes a significant fraction of the total acceleration potential difference.

Kaufman, H. R.↗
NASA NTRS1981

30-cm mercury ion thruster technology

The LeRC/Hughes 30-cm mercury ion thruster has been developed to a state of maturity such that is has become meaningful to formulate models for describing the performance characteristics of the major subassemblies. The thruster hollow cathode and the ion optics subassemblies have been investigated with this objective and conceptual, semiquantitative models have been formulated for relating lifetime and performance capabilities to design and operating parameters. This paper summarizes the investigations, discusses the factors considered for inclusion in the models, and describes the status of the models.

Poeschel, R. L.↗
NASA NTRS1982

Pulsed plasma thruster ignitor plug ignition characteristics

The operating characteristics of the semiconductor spark gap ignitor plug used to initiate the discharge in a pulsed plasma thruster are investigated. Current and voltage waveform measurements for the ignitor plug alone, and for the plug mounted in the thruster cathode indicate an average dynamic impedance of 0.2 ohms for a wide range of operating conditions, and a trigger circuit energy transfer efficiency to the plug on the order of 25%. Two modes of trigger ignition are found which are related to the rise time of the applied voltage pulse. Analysis of plasma characteristics reveals plug erosion to occur primarily in the semiconductor and anode regions, by mechanisms including plug metal embrittlement, plasma sputtering and vaporization caused by the discharge current pulse. Measurements also indicate a plume velocity on the order of 1580,000 cm/sec which increases with trigger circuit stored energy. The results establish levels of performance for future applications of semiconductor spark gap ignitor plugs in plasma devices.

Brady, M. E.↗