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At least 433 records · Page 24
All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

Operating characteristics of a hollow-cathode neutralizer for 5 and 8 centimeter-diameter electron bombardment mercury ion thrusters

Thin-tip 0.3-cm-outside-diameter hollow-cathode neutralizers were used to investigate causes of neutralizer tip erosion experienced in thruster endurance tests. Bell-jar tests indicated that neutralizers with new rolled tantalum foil inserts coated with an emissive mixture eroded very little over the neutral flow rates investigated (3 to 10 mA) for simulated 5- and 8-cm-diameter thruster neutralizer conditions. Tip erosion rates of neutralizers operated with no insert or emissive mixture increased by two orders of magnitude for both configurations as the neutral flow rate decreased. Spectroscopic analysis of the discharge plasma from neutralizers operated with inserts coated with the emissive mixture detected tungsten at all neutral flow rates for both thruster neutralizer conditions. The only source of tungsten was the tip. Therefore, detection of tungsten indicated neutralizer tip erosion. Barium, an element of the emissive mixture, was detected at low neutral flow rates for the 5-cm-diameter thruster neutralizer operating condition only.

Weigand, A. J.↗
NASA NTRS1975

A Thruster Sub-System Module (TSSM) for solar electric propulsion

Solar Electric Propulsion (SEP) is currently being studied for possible use in a number of near-earth and planetary missions. Thruster systems for these missions could be integrated directly into a spacecraft or modularized into a Thruster Sub-System Module (TSSM). A TSSM for electric propulsion missions would consist of a 30-cm ion thruster, thruster gimbal system, propellant storage and feed system, associated Power Processing Unit (PPU), thermal control system and complete supporting structure. The TSSM would be wholly self-contained and be essentially a plug-in or strap-on electric stage with simple mechanical, thermal, electrical and propellant interfaces. The TSSM described in this report is designed for a broad range of missions requiring from two to ten TSSM's mounted in a 2 by x configuration. The thermal control system is designed to accommodate waste heat from the power processor based on realistic efficiencies when the TSSM is operating from 0.7 to 3.5 AU's. The modules are 0.61 M (2 ft) wide by 2.29 M (7.5 ft) long and have a dry weight including propellant tank of 54.4 kg (120 lb). The propellant tank will hold 145.1 kg (320 lb) of mercury.

Sharp, G. R.↗
NASA NTRS1975

Determination of the extent of ion thruster efflux deposition on spacecraft surfaces from the SERT II flight thermal data

The flight thermal data from the SERT II satellite, the only electric propulsion mission with an extensive thruster operational history, is reviewed specifically to see if there is any evidence of contamination that could be attributed to the 5860 hours of mercury bombardment ion thruster operation. This evaluation shows that the only evidence of deposition occurred on the contamination experiment solar cells, which are located at the edge of the thruster exhaust beam. There is no evidence of any deposition of ion thruster efflux on any other surface of the vehicle.

Stevens, N. J.↗
NASA NTRS1975

Optical properties of mercury ion thruster exhaust plumes Significance for candidate SEP science instruments

Emission from the exhaust plume of a 30 cm mercury ion thruster was measured from 160 to 600 nm as a function of axial and radial distance from the thruster discharge chamber. The spectrally dispersed absolute intensities were used to construct an empirical volume emission rate function. The function was integrated along a typical instrument field of view, and the resulting apparent brightness was compared with instrument sensitivities to evaluate the extent of optical interference. The intensity levels degraded rapidly with distance from the thruster so that optical interference was negligible for fields of view not intercepting the beam axis. The operation of only one instrument, a zodiacal photopolarimeter was considered incompatible with simultaneous thruster operation.

Goldstein, R.↗
NASA NTRS1975

A 20,000-hour endurance test of a structurally and thermally integrated 5-cm diameter ion thruster main cathode

A 5-cm diameter mercury ion thruster main cathode has completed over 20,000 hours of operation in an ongoing lifetime endurance test. The cathode operating parameters remained at acceptable performance levels throughout the test, the first 9175 hours of which were part of a thruster endurance test. After 20,000 hours, the cathode discharge was easily restarted, the tip orifice indicated negligible erosion and the tip heater showed no degradation. The cathode-isolator-vaporizer assembly, a major thruster subsystem, has thus successfully demonstrated an operational lifetime capability of 20,000 hours, which is the lifetime goal of the 8-cm diameter auxiliary propulsion ion thruster.

Wintucky, E. G.↗
NASA NTRS1975

Studies of internal sputtering in a 30-cm ion thruster

Initial studies have been made of the sputtering and deposition phenomena in a 30-cm thruster. Sputtering rates, of the cathode baffle, one of the main sources of sputtered material in a thruster, have beem measured by weight loss as a function of several thruster parameters. Sputtering rates were found to increase with both cathode flow rate and beam current when constant discharge voltage of 37 volts and power losses of 185 ev/ion were maintained. Sputtering rates were reduced 24% as discharge voltage was decreased from 37 to 33 volts while keeping discharge power constant. Qualitative agreement was found between sputtering rates obtained by the weight loss and those implied by spectroscopically observed line intensities of the excited iron sputtered atoms. After the completion of the sputtering tests, deposition and sputtering sites inside the thruster were identified.

Mantenieks, M. A.↗
NASA NTRS1975

A reliability tradeoff study of interconnecting power processors to mercury ion thrusters

A Monte Carlo mission simulation program has been developed to optimize the solar electric propulsion (SEP) thrust subsystem configuration for an Encke comet rendezvous mission. This program examines several possible options for interconnecting power processors to the mercury ion thrusters in order to enhance mission reliability and to reduce SEP subsystem weight. The quantity of power processors and ion thrusters required to perform the mission successfully depends not only on the total required thrust, but also on the individual thruster and power processor performance and reliability. Based on these considerations, it was necessary to determine the quantity of the active power processors and ion thrusters required at any time and the means of providing redundancy. The results provide the basis for selecting an optimum SEP thrust subsystem for this mission and other missions.

Costogue, E. N.↗
NASA NTRS1975

One-millipound mercury ion thruster

A mercury ion thruster has been developed for efficient operation at the nominal 1-mlb thrust level with a specific impulse of about 3,000 sec and a total power consumption of about 120 W. At a beam voltage of 1,200 V and beam current of 72 mA, the discharge chamber operates with a propellant efficiency of 93.8% at an ion-generation energy of 276 eV/ion. The 8-cm diameter thruster advances proven component technology to assure the capability for thruster operation over an accumulated beam-on time in excess of 20,000 hours with a capability for 10,000 on-off duty cycles. Discharge chamber optimization has combined stable current-voltage characteristics with high performance efficiency by careful placement of the discharge cathode near the location of a magnetic-field zero just upstream of the thruster endplate.

Hyman, J., Jr.↗
NASA NTRS1975

Factors in the design of spacecraft utilizing multiple electric thrusters

An analysis has been performed which describes the electrical interactions between simultaneously operating ion thrusters and between the ion thrusters and a spacecraft bus. The problem is reduced to a relatively simple equivalent circuit including all leakage paths. The influence of the various circuits elements on a spacecraft potential is studied. A thruster grounding scheme is introduced which is designed to minimize the electrical interactions and includes protection for both thrusters and spacecraft from possible damage due to neutralizer failure. The paper includes the results of various experiments which have been conducted to quantitatively define the elements of the equivalent circuit.

Fitzgerald, D. J.↗
NASA NTRS1974

Study of monopropellants for electrothermal thrusters

A 333 mN electrothermal thruster designed to use MIL-grade hydrazine was demonstrated to be suitable for operation with low freezing point monopropellants containing hydrazine azide, monomethylhydrazine, unsymmetrical-dimethylhydrazine and ammonia. The steady-state specific impulse was greater than 200 sec for all propellants. The pulsed-mode specific impulse for an azide blend exceeded 175 sec for pulse widths greater than 50 msec; propellants containing carbonaceous species delivered 175 sec pulsed-mode specific impulses for pulse widths greater than 100 msec. Longer thrust chamber residence times were required for the carbonaceous propellants; the original thruster design was modified by increasing the characteristic chamber length and screen packing density. Specific recommendations were made for the work required to design and develop flight worthy thrusters, including methods to increase propellant dispersal at injection, thruster geometry changes to reduce holding power levels and methods to initiate the rapid decomposition of the carbonaceous propellants.

Kuenzly, J. D.↗
NASA NTRS1976

Ion beam thruster shield

An ion thruster beam shield is provided that comprises a cylindrical housing that extends downstream from the ion thruster and a plurality of annular vanes which are spaced along the length of the housing, and extend inwardly from the interior wall of the housing. The shield intercepts and stops all charge exchange and beam ions, neutral propellant, and sputter products formed due to the interaction of beam and shield emanating from the ion thruster outside of a fixed conical angle from the thruster axis. Further, the shield prevents the sputter products formed during the operation of the engine from escaping the interior volume of the shield.

Power, J. L.↗
NASA NTRS1976

Status of 30-centimeter-diameter mercury ion thruster isolator development

Results are presented of several 30 cm diameter mercury ion thruster isolator life tests that show that the onset and exponential increase of leakage current problems observed in earlier thruster operations and isolator tests have been solved. A 10,006 hour life test of a main isolator vaporizer operated with no mercury flow at 320 C and 1500 volts was found to have no onset of leakage current during the test. A cathode-isolator vaporizer operated with a mercury discharge at 340 to 360 C and 1200 volts for 18,000 hours, was found to have a small increase of leakage current with time. A 10,000 hour thruster life test exhibited no increase of leakage current during the life test. Isolators have been developed which will satisfy 30 cm mercury ion thruster mission requirements.

Mantenieks, M. A.↗
NASA NTRS1976

Interaction of a solar array with an ion thruster due to the charge-exchange plasma

The generation of a charge exchange plasma by a thruster, the transport of this plasma to the solar array, and the interaction of the solar array with the plasma after it arrives are all described. The generation of this plasma is described accurately from thruster geometry and operating conditions. The transport of the charge exchange plasma was studied experimentally with a 15 cm thruster. A model was developed for simple thruster array configurations. A variety of experiments were surveyed for the interaction of the plasma at the solar array.

Kaufman, H. R.↗
NASA NTRS1977

High-power and 2.5 kW advanced-technology ion thruster

Investigations for improving ion thruster components in the 30 cm engineering model thruster (EMT) resulted in the demonstration of useful techniques for grid short removal and discharge chamber erosion monitoring, establishment of relationships between double ion production and thruster operating parameters, verification of satisfactory specifications on porous tungsten vaporizer material and barium impregnated porous tungsten inserts, demonstration of a new hollow cathode configuration, and specification of magnetic circuit requirements for reproducing desired magnetic mappings. The capacity of a 30 cm EMT to operate at higher beam voltages and currents (higher power) was determined. Operation at 2 A beam current and higher beam voltage is shown to be essentially equivalent to operation at 1.1 kV with regard to efficiency, lifetime and operating conditions. The only additional requirement is an improvement in high voltage insulation and propellant isolator capacity. Operation at minimum voltage and higher beam currents is shown to increase thruster discharge chamber erosion in proportion to beam current. Studies to find alternatives to molybdenum for manufacturing ion optics grids are also reported.

Poeschel, R. L.↗
NASA NTRS1977

Operational compatibility of 30-centimeter-diameter ion thruster with integrally regulated solar array power source

System tests were performed in which Integrally Regulated Solar Arrays (IRSA's) were used to directly power the beam and accelerator loads of a 30-cm-diameter, electron bombardment, mercury ion thruster. The remaining thruster loads were supplied from conventional power-processing circuits. This combination of IRSA's and conventional circuits formed a hybrid power processor. Thruster performance was evaluated at 3/4- and 1-A beam currents with both the IRSA-hybrid and conventional power processors and was found to be identical for both systems. Power processing is significantly more efficient with the hybrid system. System dynamics and IRSA response to thruster arcs are also examined.

Gooder, S. T.↗
NASA NTRS1976

The interactions of solar arrays with electric thrusters

The generation of a charge-exchange plasma by a thruster, the transport of this plasma to the solar array, and the interaction of the solar array with the plasma after it arrives are all described. The generation of this plasma can be described accurately from thruster geometry and operating conditions. The transport of the charge-exchange plasma was studied experimentally with a 15 cm thruster. A model was developed for simple thruster-array configurations. A variety of experiments were surveyed for the interaction of the plasma at the solar array.

Kaufman, H. R.↗
NASA NTRS1976

Status of 30-centimeter-diameter mercury ion thruster isolator development

Results are presented for several 30-cm-diameter mercury-ion-thruster isolator life tests which show that the onset and exponential increase of leakage-current problems observed in earlier thruster operations and isolator tests have been solved. A 10,006-hour life test of a main isolator-vaporizer operated with no mercury flow at 320 C and 1500 volts was found to have no onset of leakage current during the test. A cathode-isolator-vaporizer (CIV) operated with a mercury discharge at 340 to 360 C and 1200 volts for 18,000 hours was found to have a small increase (0.004 micro A/hr) in leakage current with time. A 10,000-hour thruster life test exhibited no increase in leakage current during the life test. It is concluded that isolators have been developed which will satisfy 30-cm mercury-ion-thruster mission requirements.

Mantenieks, M. A.↗
NASA NTRS1976

Characteristics of a 30-cm thruster operated with small hole accelerator grid ion optics

Small hole accelerator grid ion optical systems have been tested as a possible means of improving 30-cm ion thruster performance. The effects of small hole grids on the critical aspects of thruster operation including discharge chamber performance, doubly-charged ion concentration, effluent beam characteristics, and plasma properties have been evaluated. In general, small hole accelerator grids are beneficial in improving thruster performance while maintaining low double ion ratios. However, extremely small accelerator aperture diameters tend to degrade beam divergence characteristics. A quantitative discussion of these advantages and disadvantages of small hole accelerator grids, as well as resulting variations in thruster operation characteristics, is presented.

Vahrenkamp, R. P.↗