Investigation of the arcjet plume near field using electrostatic probes

The near field plum of a 1 kW class arcjet thruster was investigated using electrostatic probes of various geometries. The electron number densities and temperatures were determined in a simulated hydrazine plume at axial distances between 3 cm (1.2 in) and 15 cm (5.9 in) and radial distances extending to 10 cm (3.9 in) off centerline. Values of electron number densities obtained using cylindrical and spherical probes of different geometries agreed very well. The electron density on centerline followed a source flow approximation for axial distances as near as 3 cm (1.2 in) from the nozzle exit plane. The model agreed well with previously obtained data in the far field. The effects of propellant mass flow rate and input power level were also studied. Cylindrical probes were used to obtain ion streamlines by changing the probe orientation with respect to the flow. The effects of electrical configuration on the plasma characteristics of the plume were also investigated by using a segmented anode/nozzle thruster. The results showed that the electrical configuration in the nozzle affected the distribution of electrons in the plume

Investigation of the Arcjet near Field Plume Using Electrostatic Probes

The near field plume of a 1 kW class arcjet thruster was investigated using electrostatic probes of various geometries. The electron number densities and temperatures were determined in a simulated hydrazine plume at axial distances between 3 cm (1.2 in.) and 15 cm (5.9 in.) and radial distances extending to 10 cm (3.9 in.) off centerline. Values of electron number densities obtained using cylindrical and spherical probes of different geometries agreed very well. The electron density on centerline followed a source flow approximation for axial distances as near as 3 cm (1.2 in.) from the nozzle exit plane. The model agreed well with previously obtained data in the far field. The effects of propellant mass flow rate and input power level were also studied. Cylindrical probes were used to obtain ion streamlines by changing the probe orientation with respect to the flow. The effects of electrical configuration on the plasma characteristics of the plume were also investigated by using a segmented anode/nozzle thruster. The results showed that the electrical configuration in the nozzle affected the distribution of electrons in the plume

The effects of arcjet operating condition and constrictor geometry on the plasma plume

Measurements of plasma number density and electron temperature were obtained in the plumes of lab arcjet thrusters using electrostatic probes of both spherical and cylindrical geometry. The two arcjet thrusters used had different constrictor and/or nozzle geometries and operated on mixtures of nitrogen, hydrogen, and ammonia to simulate the decomposition products of hydrazine and ammonia. An increase in the measured electron density was observed for both geometries with increasing arc power at a constant mass flow rate and with increasing mass flow rate at a constant arc current. For a given operating condition, the electron number density decreased exponentially off centerline and followed an inverse distance squared relationship along the thrust axis. Typical measured electron temperatures ranged from 0.1 to 0.2 eV

Arcjet thermal characteristics

The performance of water-cooled and radiation cooled arcjet thrusters operating on both 1:2 nitrogen/hydrogen mixtures at 1 to 2 kW and on pure hydrogen at 1 to 4 kW are compared. To investigate the effects of test facility background pressure on performance, data were taken for both thruster operating on nitrogen/hydrogen mixtures at facility background pressures nominally at 0.20 Pa and 20 Pa. It is shown that increasing the background pressure decreased the thruster performance, and simple pressure area corrections alone could not account for observed degradation in performance

Effects of anode material on arcjet performance

Anodes fabricated from four different materials were tested in a modular arcjet thruster at 1 kW power level on nitrogen/hydrogen mixtures. A two-percent thoriated tungsten anode served as the control. Graphite was chosen for its ease in fabrication, but experienced severe erosion in the constrictor and diverging side. Hafnium carbide and lanthanum hexaboride were chosen for their low work functions but failed due to thermal stress and reacted with the propellant. When compared to the thoriated tungsten nozzle, thruster performance was significantly lower for the lanthanum hexaboride insert and the graphite nozzle, but was slightly higher for the hafnium carbide nozzle. Both the lanthanum hexaboride and hafnium carbide nozzle operated at higher voltages. An attempt was made to duplicate higher performance hafnium carbide results, but repeated attempts at machining a second anode insert were unsuccessful. Graphite, hafnium carbide, and lanthanum hexaboride do not appear viable anode materials for low power arcjet thrusters

A Soft-Start Circuit for Arcjet Ignition

The reduced propellant flow rates associated with high performance arcjets have placed new emphasis on electrode erosion, especially at startup. A soft-start current profile was defined which limited current overshoot during the initial 30 to 50 ms of operation, and maintained significantly lower than the nominal arc current for the first eight seconds of operation. A 2-5 kW arcjet PPU was modified to provide this current profile, and a 500 cycle test using simulated fully decomposed hydrazine was conducted to determine the electrode erosion during startup. Electrode geometry and mass flow rates were selected based on requirements for a 600 second specific impulse mission average arcjet system. The flow rate was varied throughout the test to simulate the blow down of a flight propellant system. Electrode damage was negligible at flow rates above 33 mg/s, and minor chamfering of the constrictor occurred at flow rates of 33 to 30 mg/s, corresponding to flow rates expected in the last 40 percent of the mission. Constrictor diameter remained unchanged and the thruster remained operable at the completion of the test. The soft-start current profile significantly reduced electrode damage when compared to state of the art starting techniques

Multiphase Flow Technology Impacts on Thermal Control Systems for Exploration

The Two-Phase Flow Facility (TPHIFFy) Project focused on bridging the critical knowledge gap by developing and demonstrating critical multiphase fluid products for advanced life support, thermal management and power conversion systems that are required to enable the Vision for Space Exploration. Safety and reliability of future systems will be enhanced by addressing critical microgravity fluid physics issues associated with flow boiling, condensation, phase separation, and system stability. The project included concept development, normal gravity testing, and reduced gravity aircraft flight campaigns, in preparation for the development of a space flight experiment implementation. Data will be utilized to develop predictive models that could be used for system design and operation. A single fluid, two-phase closed thermodynamic loop test bed was designed, assembled and tested. The major components in this test bed include: a boiler, a condenser, a phase separator and a circulating pump. The test loop was instrumented with flow meters, thermocouples, pressure transducers and both high speed and normal speed video cameras. A low boiling point surrogate fluid, FC-72, was selected based on scaling analyses using preliminary designs for operational systems. Preliminary results are presented which include flow regime transitions and some observations regarding system stability

Hydrogen arcjet technology

During the 1960's, a substantial research effort was centered on the development of arcjets for space propulsion applications. The majority of the work was at the 30 kW power level with some work at 1-2 kW. At the end of the research effort, the hydrogen arcjet had demonstrated over 700 hours of life in a continuous endurance test at 30 kW, at a specific impulse over 1000 s, and at an efficiency of 0.41. Another high power design demonstrated 500 h life with an efficiency of over 0.50 at the same specific impulse and power levels. At lower power levels, a life of 150 hours was demonstrated at 2 kW with an efficiency of 0.31 and a specific impulse of 935 s. Lack of a space power source hindered arcjet acceptance and research ceased. Over three decades after the first research began, renewed interest exists for hydrogen arcjets. The new approach includes concurrent development of the power processing technology with the arcjet thruster. Performance data were recently obtained over a power range of 0.3-30 kW. The 2 kW performance has been repeated; however, the present high power performance is lower than that obtained in the 1960's at 30 kW, and lifetimes of present thrusters have not yet been demonstrated. Laboratory power processing units have been developed and operated with hydrogen arcjets for the 0.1 kW to 5 kW power range. A 10 kW power processing unit is under development and has been operated at design power into a resistive load

Method and apparatus for pressure pulse arcjet starting

The invention disclosed is directed to a method and apparatus for an arcjet starter. The invention discloses a method of moving an arc from the subsonic region of the thruster to the supersonic region by introducing a pressurized propellant into the casuty of the anode

Benefits of Low-Power Electrothermal Propulsion

Mission analyses were completed to show the benefits of low-power electrothermal propulsion systems for three classes'of LEO smallsat missions. Three different electrothermal systems were considered: (1) a 40 W ammonia resistojet system, (2) a 600 W hydrazine arcjet system, and (3) a 300 W ammonia resistojet. The benefits of using two 40 W ammonia resistojet systems were analyzed for three months of drag makeup of a Shuttle-launched 100 kg spacecraft in a 297 km orbit. The two 46 W resistojets decreased the propulsion system wet mass by 50% when compared to state-of-art hydrazine monopropellant thrusters. The 600 W arcjet system was used for a 300 km sun synchronous makeup mission of a 1000 kg satellite and was found to decrease the wet propulsion mass by 30%. Finally, the 300 W arcjet system was used on a 200 kg Earth-orbiting spacecraft for both orbit transfer from 300 to 400 km, two years of drag makeup, and a final orbit rise to 700 km. The arcjet system was determined to halve the propulsion system wet mass required for that scenario as compared to hydrazine monopropellant thrusters