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At least 919 records · Page 51
All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

Direct thrust measurement of a 30-cm ion thruster

A direct thrust measurement of a 30-cm diameter ion thruster was accomplished by means of a laser interferometer thrust stand. The thruster was supported in a pendulum manner by three 3.65-m long wires. Electrical power was provided by means of 18 mercury filled pots. A movable 23-button planar probe rake was used to determine thrust loss due to ion beam divergence. Values of thrust, thrust loss due to ion beam divergence, and thrust loss due to multiple ionization were measured for ion beam currents ranging from 0.5 to 2.5 A. Measured thrust values indicate an accuracy of approximately 1% and are in good agreement with thrust values calculated by indirect measurements.

Banks, B.↗
NASA NTRS1975

Designing solid propellant pulsed plasma thrusters

In this paper fundamental definitions and a few semiempirical correlations of experimental data of solid propellant pulsed plasma thrusters are used to illustrate the design analysis of such propulsion systems. The design analysis presented facilitates the design of a thruster to meet a specific performance level and provides the spacecraft designer with a methodology to select a pulsed plasma propulsion system to meet his mission requirements.

Guman, W. J.↗
NASA NTRS1975

15 cm cusped magnetic field mercury ion thruster research

The importance of achieving a uniform current density in the exhaust beam of an electrostatic ion thruster is discussed in terms of thrust level and accelerator grid lifetime. A neutral residence time approach is used to propose a magnetic field geometry which should produce a highly uniform beam current density. The discharge chamber length to diameter ratio is shown to be an important optimization parameter and experimental evaluation of the cusped field thruster over a wide range of this parameter is presented. Beam profile measurements 5 cm downstream of the accelerator grid indicate a beam profile flatness parameter which is 25% greater than the SERT II value. Flatness parameters extrapolated to the plane of the accelerator grid are demonstrated to be as high as 0.9.

Beattie, J. R.↗
NASA NTRS1975

Performance mapping of a 30 cm engineering model thruster

A 30 cm thruster representative of the engineering model design has been tested over a wide range of operating parameters to document performance characteristics such as electrical and propellant efficiencies, double ion and beam divergence thrust loss, component equilibrium temperatures, operational stability, etc. Data obtained show that optimum power throttling, in terms of maximum thruster efficiency, is not highly sensitive to parameter selection. Consequently, considerations of stability, discharge chamber erosion, thrust losses, etc. can be made the determining factors for parameter selection in power throttling operations. Options in parameter selection based on these considerations are discussed.

Poeschel, R. L.↗
NASA NTRS1975

An experimental investigation of multiple ion processes in mercury bombardment thrusters

Utilizing a collimated E x B mass separator probe, the ratio of double to single ion current densities has been measured in the beam of a 30 cm ion thruster over a wide range of beam and discharge conditions. For a specified thruster geometry, the mass utilization efficiency has been found to be the governing parameter in the determination of the doubly-charged ion content. By a reduction in the open area of the accelerator grid, a reduction in the doubly-charged ion ratio could be achieved while maintaining overall efficiency constant. The results are examined in terms of ion densities and neutral loss rates.

Vahrenkamp, R.↗
NASA NTRS1975

Solutions for discharge chamber sputtering and anode deposit spalling in small mercury ion thrusters

Small electron bombardment mercury ion thrusters are considered for attitude-control and station-keeping operations in the case of earth synchronous satellites. The required operational lifetimes of the order of 20,000 hours could present problems because of the occurrence of sputter erosion of thruster discharge chamber components. The investigation reported is concerned with a study of these problems and the development of approaches for overcoming them.

Power, J. L.↗
NASA NTRS1975

A structural and thermal packaging approach for power processing units for 30-cm ion thrusters

Solar Electric Propulsion (SEP) is currently being studied for possible use in a number of near earth and planetary missions. The thruster subsystem for these missions would consist of 30 centimeter ion thrusters with Power Processor Units (PPU) clustered in assemblies of from two to ten units. A preliminary design study of the electronic packaging of the PPU has been completed at Lewis Research Center of NASA. This study evaluates designs meeting the competing requirements of low system weight and overall mission flexibility. These requirements are evaluated regarding structural and thermal design, electrical efficiency, and integration of the electrical circuits into a functional PPU layout.

Maloy, J. E.↗
NASA NTRS1975

Mathematical model of the current density for the 30-cm engineering model thruster

Mathematical models are presented for both the singly and doubly charged ion current densities downstream of the 30-cm engineering model thruster with 0.5% compensated dished grids. These models are based on the experimental measurements of Vahrenkamp at a 2-amp ion beam operating condition. The cylindrically symmetric beam of constant velocity ions is modeled with continuous radial source and focusing functions across 'plane' grids with similar angular distribution functions. A computer program is used to evaluate the double integral for current densities in the near field and to obtain a far field approximation beyond 10 grid radii. The utility of the model is demonstrated for (1) calculating the directed thrust and (2) determining the impingement levels on various spacecraft surfaces from a two-axis gimballed, 2 x 3 thruster array.

Cuffel, R. F.↗
NASA NTRS1976

The effects of an ion-thruster exhaust plume on S-band carrier transmission

The magnitude of the effects of an ion thruster plume on S-band signals is measured. Modeling techniques are developed to predict the effects. Results show that the RF signal transmitted through an ion thruster plume is reduced in amplitude and shifted in phase. An increase in noise is also experienced.

Ackerknecht, W. E., III↗
NASA NTRS1976

The ion-optics of a two-grid electron-bombardment thruster

A detailed experimental investigation was performed to determine the ion-optical performance of an electron-bombardment ion thruster as a function of grid geometry changes. The results show that each geometrical grid parameter independently affects one aspect of ion-optical performance. These observations are used in developing a graphical technique to predict the ion-optical performance of an arbitrary ion source and grid geometry combination. The usefulness of this technique is demonstrated by comparing predicted ion-optical performance of the 30-cm diameter engineering model ion thruster with independent experimental determinations. Good agreement is shown between predicted and experimental results.

Aston, G.↗
NASA NTRS1976

A 15,000-hour cyclic endurance test of an 8-centimeter-diameter electron bombardment mercury ion thruster

A laboratory model 8 cm thruster with improvements to minimize ion chamber erosion and peeling of sputtered metal was subjected to a cyclic endurance test for 15,040 hours and 460 restarts. A charted history of several thruster operating variables and off-normal events are shown in 600-hour segments at three points in the test. The transient behavior of these variables during a typical start-stop cycle is presented. Finding of the post-test inspection confirmed most of the expected results. Charge exchange ions caused normal accelerator grid erosion. The workability of the various design features was substantiated, and attainable improvements in propellant utilization efficiency should significantly reduce accelerator erosion.

Nakanishi, S.↗
NASA NTRS1976

Cycle life testing of 8-cm mercury ion thruster cathodes

Two main cathodes have successfully completed 2800 and 1980 cycles and three neutralizers, 3928, 3050, and 2850 cycles in ongoing cycle life tests of flight-type cathode-isolator-vaporizer and neutralizer-isolator-vaporizer assemblies for the 4.45 mN 8-cm Hg ion thruster system. Each cycle included one hour of cathode operation. Starting and operating conditions simulated those expected in a typical auxiliary propulsion mission duty cycle. The cycle life test results are presented along with results of an insert comparison test which led to the selection of a rolled foil insert type for the 8-cm Engineering Model Thruster cathodes.

Wintucky, E. G.↗
NASA NTRS1976

Sputtering phenomena of discharge chamber components in a 30-cm diameter Hg ion thruster

Sputtering and deposition rates were measured for discharge chamber components of a 30-cm diameter mercury ion thruster. It was found that sputtering rates of the screen grid and cathode baffle were strongly affected by geometry of the baffle holder. Sputtering rates of the baffle and screen grid were reduced to 80 and 125 A/hr, respectively, by combination of appropriate geometry and materials selections. Sputtering rates such as these are commensurate with thruster lifetimes of 15,000 hours or more. A semiempirical sputtering model showed good agreement with the measured values.

Mantenieks, M. A.↗
NASA NTRS1976

High reliability cathode heaters for ion thrusters

A number of space missions were proposed which utilize 30-cm mercury bombardment ion thrusters and also require a large number of thruster restarts. A test program was carried out to determine thermal cycle life of several different cathode heater designs. Plasma/flame sprayed heaters and swaged type heaters were tested. Four of the five plasma/flame sprayed heaters tested failed in a comparatively short time. Four tantalum swaged heaters that were brazed to the tantalum cathode tube were successfully tested and met the goals that were set at the start of the test.

Mueller, L. A.↗
NASA NTRS1976

Inert gas thrusters

Inert gases are of interest as possible alternatives to the usual electric thruster propellants of mercury and cesium. The multipole discharge chamber investigated was shown capable of low discharge chamber losses and flat ion beam profiles with a minimum of optimization. Minimum discharge losses were 200 to 250 eV/ion for xenon and 300 to 350 eV/ion for argon, while flatness parameters in the plane of the accelerator grid were 0.85 to 0.95. The design used employs low magnetic field strengths, which permits the use of sheet-metal parts. The corner problem of the discharge chamber was resolved with recessed corner anodes, which approximately equalized both the magnetic field above the anodes and the electron currents to these anodes. Argon hollow cathodes were investigated at currents up to about 5 amperes using internal thermionic emitters. Cathode chamber diameter optimized in the 1.0 to 2.5 cm range, while orifices diameter optimized in the 0.5 to 5 mm range. The use of a bias voltage for the internal emitter extended the operating range and facilitated starting. The masses of 15 and 30 cm flight type thrusters were estimated at about 4.2 and 10.8 kg.

Kaufman, H. R.↗
NASA NTRS1977

The 6670-Newton attitude-control thruster using hydrogen-oxygen propellant

The development of a reusable, attitude-control propulsion system for the space transportation system is discussed. A flight weight, gaseous oxygen attitude control thruster assembly was tested to obtain data on cyclic life, thermal and hydraulic characteristics, pulse response, and performance. The basic thruster components were tested in excess of 51,000 pulses and 660 seconds, steady state, with no degradation of the 93 percent characteristic exhaust velocity efficiency level. Nominal operating conditions were a chamber pressure of 207 N sq cm (300 psia), a mixture ratio of 4.0, a pulse width of 100 ms, and a pulse frequency of 2 Hz.

Gordon, L. H.↗
NASA NTRS1977

Solar array maximum power tracking with closed-loop control of a 30-centimeter ion thruster

A new solar array/ion thruster system control concept has been developed and demonstrated. An ion thruster beam load is used to automatically and continuously operate an unregulated solar array at its maximum power point independent of variations in solar array voltage and current. Preliminary tests were run which verified that this method of control can be implemented with a few, physically small, signal level components dissipating less than two watts.

Gruber, R. P.↗
NASA NTRS1976

Cycle life testing of 8-cm mercury ion thruster cathodes

Two main cathodes have successfully completed 2800 and 1980 cycles and three neutralizers, 3928, 3050, and 2850 cycles in ongoing cycle life tests of flight-type cathode-isolator-vaporizer and neutralizer-isolator-vaporizer assemblies for the 4.45 mN 8-cm Hg ion thruster system. Each cycle included one hour of cathode operation. Starting and operating conditions simulated those expected in a typical auxiliary propulsion mission duty cycle. This paper presents the cycle life test results and also results of an insert comparison test which led to the selection of a rolled foil insert type for the 8-cm Engineering Model Thruster cathodes.

Wintucky, E. G.↗