Ammonia Arcjet Research Articles

Orbital Performance Measurements of Air Force Electric Propulsion Space Experiment Ammonia Arcjet

During the Electric Propulsion Space Experiment mission, eight e rings of the 26-kW ammonia arcjet were performed. Data from onboard systems,including an accelerometerandglobal positioning system unit, are used to determine thruster performance. In addition, ground-based tracking is used to determine velocity change during these e rings. The mean values of thrust, specie c impulse, and thrust efe ciency are estimated to be 1.93 § 0.06 N, 786.2§ 43.0s,and 0.267 § 0.021,respectively.Thismeasuredperformanceislowerthanexpectedbasedon ground test.Themostlikelycauseofthisdiscrepancyiserrorinonboardmeasurementofdischargepowerduetoa6%drift inthepowerprocessing unitcurrent shunt.Atthecorrected power, performancefallswithintheexpectedenvelope.

Effect of ESEX 26-kW Arcjet Operation on Spacecraft Communications

Tests designed to observe the electromagnetic compatibility of the Electric Propulsion Space Experiment 26-kW ammonia arcjet on normal spacecraft communications and operations showed no conclusive adverse effect. Two onboard antennas sensitive to the 2-, 4-, 8-, and 12-GHz frequencies detected no increase in signal amplitude clearly identifiable with arcjet operation. Analysis of the bit-error-rate test data revealed no obvious correlation between arcjet operation and the observed increases in bit-error rate. Finally, a series of qualitative observations consistently indicated the benign nature of arcjet operation on normal spacecraft events. For example, commands uplinked without abnormal rejection rate and telemetry downlinked successfully during arcjet operation.

Solar Cell Degradation During the 26-Kilowatt Electric Propulsion Space Experiment Flight

The Electric Propulsion Space Experiment (ESEX) was launched and operated in early 1999 to demonstrate the compatibility and readiness of a 30-kW class ammonia arcjet for satellite propulsion applications. As part of thisflight, an array of onboard contamination sensors assessed the effects of the arcjet on the spacecraft. The sensors consisted of microbalances to measure material deposition, radiometers to assess material degradation due to thermal radiation, and solar cell segments to investigate solar array degradation. During firings, the solar cell segments showed decreasing open-circuit voltages, probably due to additional electrical load imposed by currents through the plume plasma. Over eight firings of the ESEX arcjet, and 33-min, 44-s operating time, the solar cells exhibited a 3% decrease in short-circuit current, attributable to decreased solar transmission of the cover glass. The spacecraft's main solar arrays, however, exhibited no degradation in performance. Contamination affected only the solar cell sensor segments placed near the thruster exhaust nozzle. In the backplane of the thruster, where the main arrays are located, no deleterious effects occurred, indicating that although engineering measures may be required for equipment in the immediate vicinity of the thruster, the arcjet environment is generally benign.

Mass Deposition Measurements from 26-Kilowatt Electric Propulsion Space Experiment Flight

The Electric Propulsion Space Experiment (ESEX) was launched and operated in early 1999 to demonstrate the compatibility and readiness of a 30-kW class ammonia arcjet for satellite propulsion applications. As part of the onboard sensor array, thermoelectric quartz crystal microbalances were used to measure material deposition at selected locations on the spacecraft surface. The sensors were held at a temperature that would allow them to condense metallic materials eroded from the electrodes, but that are insensitive to possible deposition of the propellant gas. The ESEX arcjet was fired eight times, constituting 33 min and 26 s of operating time. No material deposition was observed that could be attributed to nominal the arcjet firings, although during the first firing, significant deposition was observed near the arcjet nozzle. This deposition is attributed to contaminants within the arcjet body, collected during handling and storage, which were ejected during the first firing. For future programs, although engineering measures may be needed to protect spacecraft equipment in the immediate vicinity of the thruster body, the arcjet environment causes negligible deposition of electrode material.

System Engineering, Design, Integration, and Qualification of Electric Propulsion Space Experiment

Electric arcjet propulsion systems provide an attractive option for enhancing spacecraft payload mass, increasing on-orbit lifetimes, and reducing launch costs as long as the mission can accommodate the low-thrust levels inherent to electric propulsion. To demonstrate this technology, a 30-kW class ammonia arcjet flight experiment was designed, built, qualified, and operated in space as part of the 30-kW Class Arcjet Advanced Technology Transition Demonstration Program. The engineering qualification and spacecraft integration of the propulsion system are summarized.

Radiometric Analysis from the 26-Kilowatt Electric Propulsion Space Experiment Flight

The U.S. Air Force Research Laboratory's Electric Propulsion Space Experiment (ESEX) was launched and operated in early 1999 to demonstrate the compatibility and readiness of a 30-kW class ammonia arcjet for satellite propulsion applications. As part of this flight, an array of onboard contamination sensors was used to assess the effect of the arcjet and other environments on the spacecraft. The sensors consisted of microbalances to measure material deposition, radiometers to assess material degradation due to thermal radiation or contamination, and solar cell segments to investigate solar array degradation. Over eight firings of the ESEX arcjet, and 33 min, 26 s operating time, the radiometer near the thruster, viewing the arcjet plume and body, experiences a change in the thermal properties of its coating from the firings. Radiometers with no view of the arcjet, or a view of only the plume, show no change. In general, degradation effects are observed only on sensors near the thruster exhaust nozzle, a location unlikely to be used in an operational high-power electric propulsion system. No degradation effects are observed in the backplane of the thruster. For future programs, although engineering measures may be needed for spacecraft equipment in the immediate vicinity of the thruster body, the arcjet environment is generally benign.

Development and Flight Qualification of a 26-Kilowatt Arcjet Propulsion Subsystem

The Electric Propulsion Space Experiment (ESEX) flew in 1999 on the Air Force Space Test Program Advanced Research and Global Observation Satellite. ESEX was initiated to address potential in-space operational issues associated with the use of high-power electric propulsion. Potential issues included plume deposition, electromagnetic interference, and thermal radiation. Because of physical constraints of even the largest test facilities, that is, vacuum chamber size, these phenomena are best measured in space. The propulsion component for the ESEX demonstration, a 26-kW ammonia arcjet propulsion subsystem, including thruster, power processing, and propellant feed, was developed and flight qualified on the Air Force Research Laboratory'Arcjet Advanced Technology Transition Demonstration program. Each of the propulsion system elements, the principal design challenges, and the program of risk reduction and qualification testing that provided preflight verification of the arcjet propulsion subsystem are described.

Electric Propulsion Space Experiment (ESEX) On-Orbit Results

The U.S. Air Force Research Laboratory's Electric Propulsion Space Experiment was launched and successfully operated, demonstrating the compatibility and readiness of a 26-kW ammonia arcjet subsystem for satellite applications. The Electric Propulsion Space Experiment is one of nine experiments on the U.S. Air Force's Advanced Research and Global Observation Satellite. Data were acquired to characterize the thruster in four different areas: electromagnetic interactions, contamination effects, optical properties of the plume, and thruster system performance. The results demonstrated that the critical system components including the arcjet, power processor, and propellant system operated well and verified the interoperability of high-power electric propulsion with generic satellite operations.

Arcjet thruster development

For several years an intensive program has been in progress at the University of Stuttgart to investigate and develop thermal arcjets for propellants including ammonia, nitrogen-hydrogen mixtures simulating hydrazine, and hydrogen. Since hydrogen yields the highest specific impulse /sp and best efficiencies TJ, special emphasis was placed on this propellant. Arcjet power levels between 0.7-150 kW have been studied, including water- and radiation-cooled laboratory models and flight hardware. Results yielded a maximal attainable 7sp as a function of the design and power level and showed that increasing power increased /sp. Radiation-cooled arcjets show better 17 and 7sp than water-cooled devices, but raise technical problems because of the high temperatures of the thrusters, which require the use of special refractory materials. Proper arcjet optimization was done with a thorough thermal analysis, including the propellant flow. A further improvement of these thrusters was reached by regenerative cooling and by optimizing the constrictor contour. The constrictor flow is modeled by a three-channel model, the results of which are compared with experimental data. A new two-dimensional computational fluid dynamics (CFD) approach for hydrogen arcjet thrusters is presented. In 1996 a 0.7-kW ammonia arcjet is scheduled for a flight on the P3-D AMSAT satellite.

Low-power ammonia arcjet - Numerical simulations and laser-induced fluorescence measurements

Numerical simulations were performed for a 1-kW ammonia arcjet using the University of Tennessee Space Institute equilibrium arcjet computational code. Thrust predicted by the code was compared with experimental data obtained from a thrust stand, and calculations of exit plane flow conditions were compared with data obtained from multiplexed laser induced fluorescence (LIF) experiments. The code predictions for thrust follow the same trends as the experimentally measured values, but they significantly overpredict the absolute values. Multiplexed LIF experiments were performed using both atomic nitrogen and atomic hydrogen in the plume. When nitrogen is used the effects of scattering can be significantly reduced by spectral discrimination of the fluorescence. The experiments reveal that the two species have different velocities at a position 1 mm downstream from the nozzle exit plane. The code predictions of the velocity at the exit plane have been compared with the LIF experiments. These comparisons were not conclusive because the propellant velocity can change significantly between the exit plane, where the code predictions are made, and 1 mm downstream, where the LIF measurements are made. The discrepancy between these code predictions and the experimental results highlight the inadequacy of equilibrium computational codes to accurately simulate arcjet flows.

Frequency-Domain Electromagnetic Characteristics of a 26-Kilowatt Ammonia Arcjet

Abstract : Arcjet thrusters employ an arc discharge to heat propellant, which expands through a nozzle to produce thrust Spacecraft designers who desire to exploit the enhanced specific impulse of arcjets have expressed concern about the electromagnetic environment produced by the thrusters. Laboratory tests were performed to investigate the electromagnetic environment produced by a 26-kW ammonia arcjet in the frequency domain from dc to 10 GHz using antennas intended to characterize electric and magnetic fields. Results obtained with a 30-cm monopole antenna exposed to the arcjet plume correspond to those of similar ground tests at lower power. When the antenna was shielded from electrical contact with the plume by a Pyrex(R) cover, signal levels dropped sharply at all frequencies. With a covered 30-cm monopole antenna near the arcjet plume, arcjet-on signals exceeded ambient levels over the frequency range 10 kHz-5 MHz. The maximum signal level typically exceeded the ambient level of 40 dBumV/MHz by approximately 27 dBumV/MHz near 200 kHz. A likely explanation for The sharp reduction in signal level once the antenna electrode is covered is that electrons and ions from the plume are collected on the exposed antenna but not on the covered antenna. Signals on the shielded antenna may be interpreted as the result of propagating electric fields, but considerable uncertainty remains about the field source and the effects of the facility on the measurements. Although the electromagnetic fields produced by an arcjet operating as part of a satellite are probably lower than previously expected, these results raise new concerns about the character and Spacecraft compatibility of the plasma from an arcjet plume.

Effect of nozzle and cathode configuration on arcjet performance

Experimental investigations were conducted to evaluate the effects of a bell-shaped contoured nozzle and modified cathode tip shape on 30-kWe-class ammonia arcjet engine performance. The experiments were conducted in a fully instrumented facility which included a thrust stand. The performance data for an arcjet with a bell-shaped nozzle were compared to the performance data of an arcjet that had a 38-deg included-angle, conical nozzle. Thrust improvements of up to 10% were demonstrated which corresponded to a 10% improvement in thrust efficiency and up to a 35-s increase in specific impulse. The modified cathode tip had a reduced diameter and more acute conical tip with respect to the baseline cathode design. A uniform 15% decrease in arc voltage was noted over a mass flow range of 0.175-0.350 g/s using the modified cathode. However, there was no change in the engine performance characteristics.

Electric propulsion for constellation deployment and spacecraft maneuvering

This paper outlines the near-term (1990s) advantages of electric propulsion for two SDI missions: (1) the launch of a constellation of spacecraft, and (2) continual spacecraft defensive maneuvering. Ammonia arcjet and Xe-ion electric propulsion systems are compared to advanced chemical propulsion for each of these missions. The number of launch vehicles required for constellation deployment can be reduced by up to a factor of 2 when electric propulsion upper stages are used in place of advanced upper stages. Electric propulsion can provide significant benefits when used for continuous defensive maneuvering by enabling a large reduction in the initial spacecraft mass.

Electric propulsion for orbit transfer - A case study

This study assesses the viability and cost of using electric propulsion to transfer future global positioning system satellites from Shuttle parking orbits to mission orbits. Transportation-to-orbit life-cycle costs of a xenon-ion and ammonia arcjet propelled orbital transfer vehicle are compared to those of a conventional chemical upper stage. To insure a fair treatment of electric orbital transfer vehicle (EOTV) technology; worst-, median-, and best-case technology goals defining specific engineering and cost parameters were established. Use of the EOTV in lieu of a chemical upper stage potentially reduces the Space Shuttle cargo element mass by 45 to 73% and the transport-to-orbit life-cycle cost by 4 to 61%, or up to $21 million per flight, while still meeting a 90-day flight time requirement. This concept offers the potential to utilize the existing Space Shuttle fleet or expendable launch vehicles more efficiently. Heavier or higher altitude payloads can also be efficiently transported to their mission orbits by scaling up the EOTV size or accepting somewhat longer transfer times.

Spacecraft and mission design for the SP-100 flight experiment

The design and performance of a spacecraft employing arcjet nuclear electric propulsion, suitable for use in the SP-100 Space Reactor Power System (SRPS) Flight Experiment, are outlined. The vehicle design is based on a 93 kW(e) ammonia arcjet system operating at an experimentally measured specific impulse of 1031 s and an efficiency of 42.3 percent. The arcjet/gimbal assemblies, power conditioning subsystem, propellant feed system, propulsion system thermal control, spacecraft diagnostic instrumentation, and the telemetry requirements are described. A 100 kW(e) SRPS is assumed. The spacecraft mass is baselined at 5675 kg excluding the propellant and propellant feed system. Four mission scenarios are described which are capable of demonstrating the full capability of the SRPS. The missions considered include spacecraft deployment to possible surveillance platform orbits, a spacecraft storage mission, and an orbit raising round trip corresponding to possible orbit transfer vehicle (OTV) missions.

Arcjet propulsion system for an SP-100 flight experiment

The design and performance of a arcjet nuclear electric propulsion spacecraft, suitable for use in a space reactor power system (SRPS) flight experiment, are outlined. The vehicle design is based on a 92-kW ammonia arcjet system operating at a specific impulse of 1050 s and an efficiency of 45 percent. The arcjet/gimbal system, power processing unit, and propellant feed system are described. A 100-kW SRPS is assumed and the spacecraft mass is baselined at 5250 kg, excluding the propellant and propellant feed system. A radiation/arcjet efflux diagnostics package is included in the performance analysis. This spacecraft, assuming a Shuttle launch from Kennedy Space Center, can perform a 35-deg inclination change and reach a final orbit of 35,860 km with a 120-day trip time, thus providing a four-month active load for the SRPS. Alternatively, a Titan IV launch could provide a mass margin of 120 kg to a 1000km, 58-deg final orbit in 74 days.