Roger M. Myers

Launch vehicle and power level impacts on electric GEO insertion

Solar Electric Propulsion (SEP) has been shown to increase net geosynchronous spacecraft mass when used for station keeping and final orbit insertion. The impact of launch vehicle selection and power level on the benefits of this approach were examined for 20 and 25 kW systems launched using the Ariane 5, Atlas IIAR, Long March, Proton, and Sea Launch vehicles. Two advanced on-board propulsion technologies, 5 kW ion and Hall thruster systems, were used to establish the relative merits of the technologies and launch vehicles. GaAs solar arrays were assumed. The analysis identifies the optimal starting orbits for the SEP orbit raising/plane changing while considering the impacts of radiation degradation in the Van Allen belts, shading, power degradation, and oblateness. This use of SEP to provide part of the orbit insertion results in net mass increases of 15 - 38% and 18 - 46% for one to two month trip times, respectively, over just using SEP for 15 years of north/south station keeping. SEP technology was shown to have a greater impact on net masses of launch vehicles with higher launch latitudes when avoidance of solar array and payload degradation is desired. This greater impact of SEP could help reduce the plane changing disadvantage of high latitude launch sites. Comparison with results for 10 and 15 kW systems show clear benefits of incremental increases in SEP power level, suggesting that an evolutionary approach to high power SEP for geosynchronous spacecraft is possible.

Pulsed Plasma Thruster Contamination

Pulsed Plasma Thrusters (PPTs) are currently baselined for the Air Force Mightysat II.1 flight in 1999 and are under consideration for a number of other missions for primary propulsion, precision positioning, and attitude control functions. In this work, PPT plumes were characterized to assess their contamination characteristics. Diagnostics included planar and cylindrical Langmuir probes and a large number of collimated quartz contamination sensors. Measurements were made using a LES 8/9 flight PPT at 0.24, 0.39, 0.55, and 1.2 m from the thruster, as well as in the backflow region behind the thruster. Plasma measurements revealed a peak centerline ion density and velocity of approx. 6 x 10(exp 12) cm(exp -3) and 42,000 m/s, respectively. Optical transmittance measurements of the quartz sensors after 2 x 10(exp 5) pulses showed a rapid decrease in plume contamination with increasing angle from the plume axis, with a barely measurable transmittance decrease in the ultraviolet at 90 deg. No change in optical properties was detected for sensors in the backflow region.

Preliminary plume characteristics of an arcjet thruster

An experimental investigation of a low power arcjet plume was conducted using emission spectroscopy. A laboratory model arcjet incorporating a segmented anode was run on simulated hydrazine at a flow rate of 5 x 10(exp -5) kg/s. The complete visible spectrum measured in the exit plane of the arcjet showed the presence of N2, N2(+), NH, and H. Radial intensity profiles for the H alpha, H sub beta, and the NH A(sup 3)Pi yields X(sup 3)Sigma(0,0) transitions at four different axial locations were measured. These line of sight intensity measurements, spaced 0.05 mm apart, were deconvoluted to give the radial intensity distribution using an inverse Abel transformation. The ratio between the intensities from the H sub alpha and H sub beta transitions indicated a non-Boltzmann energy distribution between excited states in the plume. Axial intensity profiles taken on center line indicated the decay rate of excited states in the plume. An electron number density of 2 x 10(exp 13)/cu cm at the exit plane was determined based on Stark broadening of the H sub beta line. Rotational temperatures of 750 K, 1750 K, and 2500 K were determined for N2, N2(+), and NH respectively. The results demonstrate that the location of the current attachment on the anode has a measurable effect on the electronically excited species in the plume and that dissociation is the dominant frozen flow loss mechanism in low power arcjets.

Thermal Nonequilibrium in a Low-Power Arcjet Nozzle

Emission spectroscopy measurements were made of the plasma flow inside the nozzle of a 1-kW-class arcjet thruster. The thruster propellant was a hydrogen-nitrogen mixture used to simulate fully decomposed hydrazine. Several 0.25-mm-diam holes were drilled into the 12-mm-long diverging section of the tungsten thruster nozzle to provide side-on optical access to the internal flow. Electron excitation for atomic species and molecular vibrational and rotational temperatures were determined for the expanding plasma using relative line ratio techniques. The atomic excitation temperature decreased from 18,000 K at a location 3-mm downstream of the constrictor to 9000 K at a location 9 mm from the constrictor, while the molecular vibrational and rotational temperatures decreased from 6500 to 3000 K and 8000 to 3000 K, respectively, between the same locations. The electron density, measured using H^ line Stark broadening, decreased from ~1021 m~3 to 2 x 10 20 m~3 during the expansion in the nozzle. The results show that the plasma is in a nonequilibrium state throughout most of the nozzle, with relaxation times close to or larger than the particle residence time.

Test Facilities for High-Power Electric Propulsion

Electric propulsion has applications for orbit raising, maneuvering of large space systems, and interplanetary missions. These missions involve propulsion power levels from tenths to tens of megawatts, depending upon the application. General facility requirements for testing high power electric propulsion at the component and thrust systems level are defined. The characteristics and pumping capabilities of many large vacuum chambers in the United States are reviewed and compared with the requirements for high power electric propulsion testing.

Geometric Scaling of Applied-Field Magnetoplasmadynamic Thrusters

Eight magnetoplasmadynamic thruster configurations were tested with argon propellant at power levels between 20 -130 kW to study the effects of geometry and applied magnetic field strength on thruster performance. The discharge voltage, thrust efficiency, and specific impulse increased monotonically with increasing applied-field strength for all geometries. The highest measured performance was 23% thrust efficiency at 2300s specific impulse at a power of 113 kW. Both cathode and anode radii fundamentally influenced the efficiencyspecific impulse relationship, whereas their lengths influenced only the magnitude of the applied magnetic field required to reach a given performance level. At a given specific impulse, large electrode radii resulted in lower efficiencies, although the rate at which efficiency increased with applied-field strength was higher with larger anode radii. Anode power deposition was the largest efficiency loss, and represented between 50-80% of the input power. The fraction of the input power deposited into the anode decreased with increasing applied-field strength and anode radius.

Nonequilibrium in a low power arcjet nozzle

Emission spectroscopy measurements were made of the plasma flow inside the nozzle of a 1 kW class arcjet thruster. The thruster propellant was a hydrogen-nitrogen mixture used to simulate fully decomposed hydrazine. The 0.25 mm diameter holes were drilled into the diverging section of the tungsten thruster nozzle to provide optical access to the internal flow. Atomic electron excitation, vibrational, and rotational temperatures were determined for the expanding plasma using relative line intensity techniques. The atomic excitation temperatures decreased from 18,000K at a location 3 mm downstream of the constrictor to 9,000K at a location 9 mm from the constrictor, while the molecular vibrational and rotational temperatures decreased from 6,500K to 2,500K and from 8,000K to 3,000K, respectively, between the same locations. The electron density measured using hydrogen H line Stark broadening decreased from about 10(exp 15) cm(-3) to about 2 times 10(exp 14) cm(-3) during the expansion. The results show that the plasma is highly nonequilibrium throughout the nozzle, with most relaxation times equal or exceeding the particle residence time.

Techniques for spectroscopic measurements in an arcjet nozzle

A successful attempt has been made to gain optical access to the inside of an arcjet nozzle without changing internal thruster design or affecting performance characteristics. Both fiber optics and small open holes have been used for emission spectroscopy of a small, confined, high-temperature plasma source. The plasma was found to be in a highly nonequilibrium state, with electron excitation temperatures more than double the rotational or vibrational temperatures.

100-kW class applied-field thruster component wear

Component erosion and material deposition sites were identified and analyzed during tests of various configurations of 100 kW class, applied-field, water-cooled magnetoplasmadynamic (MPD) thrusters. Severe erosion of the cathode and the boron nitride insulator was observed for the first series of tests, which was significantly decreased by reducing the levels of propellant contamination. Severe erosion of the copper anode resulting from sputtering by the propellant was also observed. This is the first observation of this phenomenon in MPD thrusters. The anode erosion indicates that development of long life MPD thrusters requires the use of light gas propellants such as hydrogen, deuterium, or lithium.

NSTAR Ion Thruster Plume Impact Assessments

Tests were performed to establish 30-cm ion thruster plume impacts, including plume characterizations via near and farfield ion current measurements, contamination, and sputtering assessments. Current density measurements show that 95% of the beam was enclosed within a 22 deg half-angle and that the thrust vector shifted by less than 0.3 deg during throttling from 2.3 to 0.5 kW. The beam flatness parameter was found to be 0.47, and the ratio of doubly charged to singly charged ion current density decreased from 15% at 2.3 kW to 5% at 0.5 kW. Quartz sample erosion measurements showed that the samples eroded at a rate of between 11 and 13 pm/khr at 25 deg from the thruster axis, and that the rate dropped by a factor of four at 40 deg. Good agreement was obtained between extrapolated current densities and those calculated from tantalum target erosion measurements. Quartz crystal microbalance and witness plate measurements showed that ion beam sputtering of the tank resulted in a facility material backflux rate of -10 A/hr in a large space simulation chamber.