Electrothermal hydrazine thruster analyses and performance evaluation.
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Callens, R. A.↗
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A review of the impact of ground test-related facility effects on gridded ion thruster operation and performance
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John E. Foster↗
Ion thruster thermal characteristics and performance.
Experimental and analytical studies of mercury-fed hollow cathode ion thrusters are presented. Pertinent thermal properties and plasma heating characteristics of the system have been determined through correlation and integration of temperature measurement data with a single stage Wiener-Kalman estimation technique. The resultant analytical network model is applied to predict the thermal performance of specified thruster systems under various operating conditions. The effects of solar intensity levels on the thruster thermal behavior, the dependency on the thruster dimension and multithruster design arrangements have been examined.
Wen, L.↗
Performance of a 15 cm ion thruster with reliable restart capability
Experimental evaluation of a modified SERT II thruster with dished grids capable of producing a 600 ma ion beam at 2690 sec specific impulse and achieving rapid, reliable startup is described. In addition to steady-state performance curves, thruster performance is examined over the duration of rapid startup and one hour operating sequence. Rapid startup utilizing high voltage spark ignition of the main and neutralizer cathodes is described. Minimal deterioration of the cathodes and the ignition system during 1000 neutralizer cathode startups and 2000 main cathode startups is demonstrated. The need for heaters and a vaporizer control system to facilitate more rapid startup is identified.
Wilbur, P. J.↗
Impact of reverse flow induced by a sawtooth anode on the performance of an argon Hall thruster
This study reports the performance comparison of the RAIJIN-66 TAL-type Hall thruster with argon propellant, using two hollow anode designs, one with a simple straight shape and another one with a sawtooth shape. Simulations of the rarefied gas in the sawtooth hollow anode region suggest a successful reversal of the direction of neutral particle flow, with an average increase in the neutral particle density by 15.5% when compared to the straight-shaped hollow anode. This contributes to the improvement in performance across a range of discharge voltages that was measured experimentally. When the thruster is operated with a discharge voltage of 150 V and a flow rate of 70 SCCM, the propellant utilization efficiency with the straight anode is 14%, while with the sawtooth anode it is 25.1%. The anode efficiency reaches in the vicinity of 15% with the sawtooth anode in high voltage conditions, exceeding the efficiencies achieved with the straight anode in the same operating conditions. The optimal magnetic field condition for argon operation and the tuning parameters of the sawtooth anode design for RAIJIN-66 are also discussed.
Satpathy, Dibyesh [Univ. of Tokyo (Japan)] (ORCID:↗
Inert gas ion thruster development
Two 12 cm magneto-electrostatic containment (MESC) ion thrusters were performance mapped with argon and xenon. The first, hexagonal, thruster produced optimized performance of 48.5to 79 percent argon mass utilization efficiencies at discharge energies of 240 to 425 eV/ion, respectively, Xenon mass utilization efficiencies of 78 to 95 percent were observed at discharge energies of 220 to 290 eV/ion with the same optimized hexagonal thruster. Changes to the cathode baffle reduced the discharge anode potential during xenon operation from approximately 40 volts to about 30 volts. Preliminary tests conducted with the second, hemispherical, MESC thruster showed a nonuniform anode magnetic field adversely affected thruster performance. This performance degradation was partially overcome by changes in the boundary anode placement. Conclusions drawn the hemispherical thruster tests gave insights into the plasma processes in the MESC discharge that will aid in the design of future thrusters.
Ramsey, W. D.↗
Inert gas test of two 12-cm magnetostatic thrusters
Comparative performance tests were conducted with 12 cm line and ring magnetic cusp thrusters. Shell anode and magnetoelectrostatic containment boundary anode configurations were evaluated with each magnet array. The best performance was achieved with the 12-cm ring cusp-shell anode configuration. Argon operation of this configuration produced 65-81 percent mass utilization efficiency at 170-208 watts/single-charged-equivalent (SCE) ampere beam. Xenon test results showed 75-95 percent utilization at 162-188 watts/SCE ampere beam.
Ramsey, W. D.↗
High-Power Ion Thruster Technology
Performance data are presented for the NASA/Hughes 30-cm-diam 'common' thruster operated over the power range from 600 W to 4.6 kW. At the 4.6-kW power level, the thruster produces 172 mN of thrust at a specific impulse of just under 4000 s. Xenon pressure and temperature measurements are presented for a 6.4-mm-diam hollow cathode operated at emission currents ranging from 5 to 30 A and flow rates of 4 sccm and 8 sccm. Highly reproducible results show that the cathode temperature is a linear function of emission current, ranging from approx. 1000 C to 1150 C over this same current range. Laser-induced fluorescence (LIF) measurements obtained from a 30-cm-diam thruster are presented, suggesting that LIF could be a valuable diagnostic for real-time assessment of accelerator-arid erosion. Calibration results of laminar-thin-film (LTF) erosion badges with bulk molybdenum are presented for 300-eV xenon, krypton, and argon sputtering ions. Facility-pressure effects on the charge-exchange ion current collected by 8-cm-diam and 30-cm-diam thrusters operated on xenon propellant are presented to show that accel current is nearly independent of facility pressure at low pressures, but increases rapidly under high-background-pressure conditions.
Beattie, J. R.↗
Plasma properties and performance of mercury ion thrusters.
Electron bombardment mercury ion thrustor plasma properties and performance, computing ion beam current, discharge losses and propellant utilization efficiency
Masek, T. D.↗
Plasma properties and performance of mercury ion thrusters
Electron bombardment mercury ion thrustor plasma properties and performance, computing ion beam current, discharge losses and propellant utilization efficiency
Masek, T. D.↗
Environmental Testing of the HERMeS TDU-2 Hall Thruster
The Hall Effect Rocket with Magnetic Shielding (HERMeS) is a 12.5 kW magnetically shielded Hall thruster presently under development for deep-space missions. A development model HERMeS thruster, TDU-2, with borosilicate channel walls, has successfully undergone proto-qualification testing including random vibration and thermal-vacuum, demonstrating the thruster can survive critical launch and operational environmental requirements. The three-axis random vibrational tests were performed with the thruster body mounted directly to the shaker table using a modified loading schedule to compensate for the vibration isolation dampers under design for the flight thruster. Initial random vibration testing revealed issues with fragmentation of magnetic coil potting material and thruster spool mount fabrication weaknesses that were addressed with an alternate potting material and fabrication revisions that were then successfully retested with an additional round of random vibration tests. Thermal-vacuum testing involved three cycles of full-power 600 V 12.5 kW thruster operation from -121°C to +373°C which included eight hours at peak temperature as well as a hot restart. A thermal model of the test configuration was developed and validated with the experiment results within an average difference of 5°C. Before and after the random vibrational and thermal-vacuum tests, the thruster performance, plasma plume, thrust vector angle, discharge characteristics, and magnetic field profile were measured. The thruster performance, magnetic field, and most other parameters remained invariant throughout the environmental testing campaign. During the thermal testing, an increase in the discharge oscillation amplitude of 11% was observed as well as subtle thrust vector changes of ±0.2° as the thruster warmed from -121°C to +373°C.
Sorenson, Ryan↗
Experiments and analysis of a compact electrothermal thruster
The description and experimental performance of a compact microwave electrothermal thruster (MET) are presented. This thruster uses a coaxial applicator to couple microwave power into a high pressure discharge. Unlike earlier experiments, it uses no fused quartz in the discharge chamber or the nozzle. This allows high temperatures in the discharge chamber without quartz erosion and melting, thereby improving thruster performance and lifetime. The thruster design is compact, enhancing its potential as a space engine. Experimental tests using nitrogen and helium propellants with input powers levels of 200 W to 1.5 kW are presented. Experimental results, which produce energy efficiencies of 20 to 60 percent and specific impulse of 250 to 450 sec, compare favorably to previous experimental MET performance.
Asmussen, Jes↗
Performance characteristics of ring-cusp thrusters with xenon propellant
The performance characteristics and operating envelope of several 30-cm ring-cusp ion thrusters with xenon propellant were investigated. Results indicate a strong performance dependence on the discharge chamber boundary magnetic fields and resultant distribution of electron currents. Significant improvements in discharge performance over J-series divergent-field thrusters were achieved for large throttling ranges, which translate into reduced cathode emission currents and reduced power dissipation which should be of significant benefit for operation at thruster power levels in excess of 10 kW. Mass spectrometry of the ion beam was documented for both the ring-cusp and J-series thrusters with xenon propellant for determination of overall thruster efficiency, and lifetime. Based on the lower centerline values of doubly charged ions in the ion beam and the lower operating discharge voltage, the screen grid erosion rate of the ring-cusp thruster is expected to be lower than the divergent-field J-series thruster by a factor of 2.
Patterson, M. J.↗
Performance characteristics of ring-cusp thrusters with xenon propellant
The performance characteristics and operating envelope of several 30-cm ring-cusp ion thrusters with xenon propellant were investigated. Results indicate a strong performance dependence on the discharge chamber boundary magnetic fields and resultant distribution of electron currents. Significant improvements in discharge performance over J-series divergent-field thrusters were achieved for large throttling ranges, which translate into reduced cathode emission currents and reduced power dissipation which should be of significant benefit for operation at thruster power levels in excess of 10 kW. Mass spectrometer of the ion beam was documented for both the ring-cusp and J-series thrusters with xenon propellant for determination of overall thruster efficiency, and lifetime. Based on the lower centerline values of doubly charged ions in the ion beam and the lower operating discharge voltage, the screen grid erosion rate of the ring-cusp thruster is expected to be lower than the divergent-field J-series thruster by a factor of 2.
Patterson, M. J.↗
Performance capabilities of the 12-centimeter Xenon ion thruster
The 8- and 12-cm mercury ion thruster systems were developed primarily to provide N-S station keeping of satellites with masses up to about 1800 to 3600 kg respectively. The on-orbit propulsion requirements of recently proposed Large Space Systems (LSS) are beyond the thrust capabilities of the baseline 8- and 12-cm thruster systems. This paper presents a characterization of the performance capabilities of the 12-cm Xenon ion thruster to enable an evaluation of its application to LSS auxiliary propulsion requirements. With minor thruster modifications and simplifications the thrust was increased to 64 mN, a factor of six over the baseline 12-cm mercury thruster performance. The thruster was operated over a range of specific impulse of about 2000 to 4000 seconds and at total efficiencies up to 68.0 percent. The operating levels reached in this study were found to be close to the operating limits of the thruster design in terms of perveance, grid breakdown voltage and thruster component temperatures such as those of the magnets and cathode baffle.
Mantenieks, M.↗
Performance capabilities of the 12-centimeter Xenon ion thruster
The 8- and 12-cm mercury ion thruster systems were developed primarily to provide N-S station keeping of satellites with masses up to about 1800 to 3600 kg respectively. The on-orbit propulsion requirements of recently proposed Large Space Systems (LSS) are beyond the thrust capabilities of the baseline 8- and 12-cm thruster systems. This paper presents a characterization of the performance capabilities of the 12-cm Xenon ion thruster to enable an evaluation of its application to LSS auxiliary propulsion requirements. With minor thruster modifications and simplifications the thrust was increased to 64 mN, a factor of six over the baseline 12-cm mercury thruster performance. The thruster was operated over a range of specific impulse of about 2000 to 4000 seconds and at total efficiencies up to 68.0 percent. The operating levels reached in this study were found to be close to the operating limits of the thruster design in terms of perveance, grid breakdown voltage and thruster component temperatures such as those of the magnets and cathode baffle.
Mantenieks, M.↗
Preliminary Results of Performance Measurements on a Cylindrical Hall-Effect Thruster with Magnetic Field Generated by Permanent Magnets
The performance of a low-power cylindrical Hall thruster, which more readily lends itself to miniaturization and low-power operation than a conventional (annular) Hall thruster, was measured using a planar plasma probe and a thrust stand. The field in the cylindrical thruster was produced using permanent magnets, promising a power reduction over previous cylindrical thruster iterations that employed electromagnets to generate the required magnetic field topology. Two sets of ring-shaped permanent magnets are used, and two different field configurations can be produced by reorienting the poles of one magnet relative to the other. A plasma probe measuring ion flux in the plume is used to estimate the current utilization for the two magnetic configurations. The measurements indicate that electron transport is impeded much more effectively in one configuration, implying a higher thrust efficiency. Preliminary thruster performance measurements on this configuration were obtained over a power range of 100-250 W. The thrust levels over this power range were 3.5-6.5 mN, with anode efficiencies and specific impulses spanning 14-19% and 875- 1425 s, respectively. The magnetic field in the thruster was lower for the thrust measurements than the plasma probe measurements due to heating and weakening of the permanent magnets, reducing the maximum field strength from 2 kG to roughly 750-800 G. The discharge current levels observed during thrust stand testing were anomalously high compared to those levels measured in previous experiments with this thruster.
Polzin, K. A.↗
Thruster array thermal control
This paper presents the results of a preliminary investigation of the thermal interactions between elements in the thruster array for a proposed Encke flyby mission. A clustered eight-thruster array was considered as the trial configuration in the current study. Electrically heated dummy thrusters were calibrated to represent the thruster thermal performance. The thermal behavior of the thruster array thermal mockup and the effects of thermal coating applications were investigated. Temperature measurements of the thruster array mockup were performed under various operational modes to substantiate the analytical results. A resultant analytical thermal network model was established to provide pertinent information for thruster array design tradeoff studies.
Wen, L.↗