A torsional-type thrust stand has been designed and built to test pulsed plasma thrusters in both single shot and repetitive operating modes. Using this stand, momentum per pulse is determined strictly as a function of thrust stand deflection, spring stiffness, and natural frequency. No empirical corrections are required. The accuracy of the method was verified using a swinging impact pendulum. Momentum transfer data between the thrust stand and the pendulum were consistent to within 1%. Following initial calibrations, the stand was used to test a Lincoln experimental satellite (LES-8/9) thruster. The LES-8/9 system had a mass of approximately 7.5 kg, with a nominal thrust to weight ratio of 8.0×10−6. A total of 34 single shot thruster pulses was individually measured. The average impulse bit per pulse was 266 μN s, which was slightly less than the value of 300 μN s published in previous reports on this device. Repetitive pulse measurements were performed similar to ordinary steady-state thrust measuremen...

Thrust stand for pulsed plasma thrusters

space, starting with the Meteor I in 1969-1970.7.8 More recently, eight SPT-100 thrusters were flown on the Russian A cyclic endurance test of the Russian 1.35 kW Stationary GALS television satellite to perform both north-south and Plasma Thruster SPT-100 is described. The endurance test east-west station-keeping functions.9 was performed for 6,925 on/off cycles and 5,730.3 hours of operation at an input power to the thruster of 1.35 kW. In 1991 a team of electric propulsion specialists Each cycle was approximately 50 minutes of thruster onvisited the former U.S.S.R. to experimentally evaluate the time and 23 minutes of thruster off-time. Thruster efficiency performance of a 1.35-kW SPT at the Scientific Research 1,000 hours. The efficiency increased slowly over the next Bureau "Fakel" in Kaliningrad, Russia. 10 -1 1 The 1,000 hours and then slowly decreased to 45% by the end of evaluation verified that the actual performance of the thruster the wear test. The unused cathode ignitor and radiation was close to the claimed performance of 50% efficiency at a shieldswerefoundto erode anunexpectedy highrate. A specific impulse of 1600 s.10 Studies indicate that for shields were found to erode at an unexpectedly high north/south station keeping and Earth orbit raising short between the cathode emitter and cathode ignitor applications of electric propulsion, the optimum specific occurred at cycle 5,316; the short was cleared without impulse is in the range of 1.000-2.000 sec. 12 The opening the vacuum tank but the short reoccurred at cycle combination of the flight heritage of the SPT-70 and the 6,344. The vacuum chamber was opened after the thruster availability of thrusters and thruster data led to substantial had accumulated 6,346 cycles and 5,250 hours of operating interest in these thrusters by Western spacecraft time to clear this short and repair the thrust stand, which had manufacturers for primary and auxiliary propulsion failed at cycle 5,818. The wear test was continued and applications. voluntarily terminated when 5,002.5 hours of operation were accumulated on cathode #1. Data from this wear test indicate Space Systems/Loral is presently flight-qualifying that the likely first failure mechanism for this thruster design SPT-100 thrusters for north/south station keeping and Earth is shorting between the ignitor and cathode emitter in the orbit raising applications and plans to provide these thrusters unused cathode. The endurance test was performed under a on their spacecraft. 13 At Design Bureau "Fakel", a steadycooperative program between Space SystemsLoral, JL, and state life test was performedl 4 , and performance, plume and EIprogram betw en ste al, JP, and /RFI evaluations were conducted at NASA Lewis the Ballistic Missile Defense Organization (BMDO). Research Centr.1 AS-18 INTRODUCTION NTRO CTION A key aspect of the SPT-100 evaluation program Stationary plasma thrusters (S ) are griless ion was characterization of the long term operating behavior of thrusters that were originally developed in the US. in the the thruster. Typical mission applications of interest require tht we originy d d e U . in te operating times of several thousand hours. Potential use of early 1960's. 1 3 Although efforts in the U. S. to develop the thruste for north-south station keeping of commercial high thrust efficiencies failed, efforts in the former U.S.S.R. communication satellites will also require the capability for were quite successful. The SPT was successfully developed several thousand on/off cycles. To address these objectives a during the 1960's and 1970's by Morozov 4 and others5 ,6 to cyclic endurance test was performed at JPL under a obtain a unique combination of specific impulse and cooperative program between Space Systems/Loral, JPL efficiency. More than 50 SPT-70 thrusters have flown in (under the JPL Affiliates program) and the BMDO. The original goal for the test was to accumulate a total of 5,000 SMember of the Technical Staff, Advanced Propulsion hours and 6,000 cycles on the thruster; the test was extended Technology Group. to accumulate more than 5,000 hours of operating time on a + Member of the Technical Staff, Electric Power Systems single cathode. Several papers describe the wear test results Section. through the first 5,000 cycles. 19 -2 1 *Tis paper describes the ** Supervisor, Advanced Propulsion Technology Group. results of this cyclic endurance test for 6,925 cycles and 5,7303 hours of operation at 1350 watts.

A 5,730-hr cyclic endurance test of the SPT-100

/pdf/an-extended-life-and-performance-test-of-a-low-power-arcjet-4axiz4u6i6.pdf

An extended life and performance test of a low-power arcjet

Arcjets and arc heaters are important to a number of government-sponsored aerospace programs and share several common features. Placing satellites in orbit, maintaining proper orbit, and repositioning are major spacecraft mass drivers for a wide range of missions. Arcjets can reduce mission propulsion requirements in many cases and are of interest to many satellite users. Arc heaters are critical to the future development of hypersonic combat aircraft. Design and construction of a hypersonic aircraft for speeds of 4-5 km/s (Mach 12-15) will be complex and will require the development of high-performance arc-heating facilities for the evaluation/validation of new airframe and propulsion technologies. The first objective of this overview is to provide a summary of the basic principles of arcjets and arc heaters for readers unfamiliar with the field. The second objective is to describe fundamental research required to overcome the limitations of both devices, with an emphasis on common research areas. This overview is not comprehensive, but does include national research and development activities. It is intended as a guide to the organization of scientific research serving both arcjet and arc-heater development communities.

Arcjets and arc heaters : an overview of research status and needs

Investigation of the plasma acceleration mechanism within a coaxial Hall thruster

An arcjet starting reliability test was performed to investigate one feasibility issue in the use of arcjets onboard a satellite for north-south stationkeeping. A 1 kW arcjet was run on hydrogen/nitrogen gas mixtures simulating decomposed hydrazine. A pulse width modulated power supply with an integral high voltage starting pulser was used for arc ignition and steady-state operation. The test was performed in four phases in order to determine if starting characteristics changed as a result of long term thruster operation. More than 300 successful starts were accumulated over an operating time of 18 hrs. Overall results indicate that there is a link between starting characteristics and long term thruster operation; however, the large number of starts had no effect on steady-state performance.

/pdf/arcjet-starting-reliability-a-multistart-test-on-hydrogen-1zpdqe2m2m.pdf

Arcjet starting reliability: A multistart test on hydrogen/nitrogen mixtures

A low power, dc arcjet thruster was tested for starting reliability using hydrogen-nitrogen mixtures simulating the decomposition products of hydrazine. More than 300 starts were accumulated in phases with extended burn-in periods interlaced. A high degree of flow stabilization was built into the arcjet and the power supply incorporated both rapid current regulation and a high voltage, pulsed starting circuit. A nominal current level of 10 A was maintained throughout the test. Photomicrographs of the cathode tip showed a rapid recession to a steady-state operating geometry. A target of 300 starts was selected, as this represents significantly more than anticipated (150 to 240), in missions of 10 yr or less duration. Weighings showed no apparent mass loss. Some anode erosion was observed, particularly at the entrance to the constrictor. This was attributed to the brief period during startup the arc mode attachment point spends in the high pressure region upstream of the nozzle. Based on the results obtained, startup does not appear to be performance or life limiting for the number of starts typical of operational satellite applications.

/pdf/arcjet-component-conditions-through-a-multistart-test-11g0qzdbmo.pdf

Arcjet component conditions through a multistart test

A segmented anode/nozzle for a low-power (1 kW) arc jet thruster was fabricated to investigate the effect of electrode configuration on the arc jet performance. The five segments of the nozzle, which could be isolated individually or in groups made it possible to observe current distribution from the constrictor through the diverging section of the nozzle; measurements of the potential difference between the cathode and any of the individual segments were possible. Results showed that the discharge initiates in the high pressure region of the nozzle upstream of the diverging section, and then rapidly moves to the diverging section of the nozzle. When the arc was allowed to seat in the diverging section, the anode fall voltage was between 10 to 20 volts. When the current was forced to the high-pressure section of the constrictor, the anode fall voltage increased to more than 40 volts.

The effect of electrode configuration on arcjet performance

Most space missions require on-board propulsion systems and these systems are often dominant spacecraft mass drivers. Presently, on-board systems account for more than half the injected mass for commercial communications systems and even greater mass fractions for ambitious planetary missions. Anticipated trends toward the use of both smaller spacecraft and launch vehicles will likely increase pressure on the performance of on-board propulsion systems. The acceptance of arcjet thrusters for operational use on commercial communications satellites ushered in a new era in on-board propulsion and exponential growth of electric propulsion across a broad spectrum of missions is anticipated. NASA recognizes the benefits of advanced propulsion and NASA's Office of Space Access and Technology supports an aggressive On-Board Propulsion program, including a strong electric propulsion element, to assure the availability of high performance propulsion systems to meet the goals of the ambitious missions envisioned in the next two decades. The program scope ranges from fundamental research for future generation systems through specific insertion efforts aimed at near term technology transfer. The On-Board propulsion program is committed to carrying technologies to levels required for customer acceptance and emphasizes direct interactions with the user community and the development of commercial sources. This paper provides a discussion of anticipated missions, propulsion functions, and electric propulsion impacts followed by an overview of the electric propulsion element of the NASA On-Board Propulsion program.

/pdf/electric-propulsion-applications-and-impacts-42fbr4lsn8.pdf

Frank M. Curran

Papers

Arcjet starting reliability: A multistart test on hydrogen/nitrogen mixtures

Arcjet component conditions through a multistart test

The effect of electrode configuration on arcjet performance

Electric Propulsion Applications and Impacts