Electric Thruster Test and Evaluation

Edward J. Beiting, Ronald B. Cohen, Mark W. Crofton, Kevin Diamant, James E. Pollard, Jun Qian, and Michael L. Garrett

Evaluation in ground test facilities plays a critical role in the development and qualification of new thruster systems.

Spacecraft electric thrusters are responsible for the critical functions of orbit transfer, on-orbit station keeping, and, in recent applications, interplanetary propulsion. Electric systems generate thrust by using electric and magnetic processes to heat and/or accelerate a propellant or plasma. Chemical systems create thrust through chemical reactions that generate expansive exhaust. Electric thrusters have an exhaust velocity normally 2 to 10 times higher than chemical thrusters, which means their efficiency with respect to propellant usage is greater. Payloads can therefore be augmented or launched on smaller, cheaper vehicles.

On the other hand, the testing and evaluation of electric thrusters is more challenging than for chemical systems. The amount of energy per expelled particle, the overall complexity, and the required lifetime is typically much greater. In addition, the considerable potential for sputtered particle deposition, energetic ion impingement, electromagnetic interference, and other interactions that could adversely affect the spacecraft and its subsystems must be addressed. The potential for adverse interaction tends to increase along with electrical power input and propellant flow rate, which have risen as available onboard power has increased.

In addition, ground test facilities interact with electric thrusters in ways that can skew test results. For example, facility background pressure can affect electromagnetic interference, thruster erosion, and electrical breakdown characteristics. Wall surfaces near the thruster produce contamination, thermal interactions, artificial plume neutralization, and perturbations on electric fields and plasma migration. Contamination from walls is one of the most insidious effects: The deposition of a conductive layer on thruster surfaces can cause the loss of insulator functions, upset emission characteristics, and modify the apparent contamination potential of the thruster.

The T5 (UK-10) ion engine operating in the Aerospace facility, revealing the fluorescent ion beam and focusing effect of the extraction grids. The ion beam is reexpanding as it strikes the beam stop on the right.

In view of all these considerations, electric propulsion test and evaluation techniques and facilities have been forced to grow in sophistication and scope.

Aerospace has a lengthy history of testing and evaluating thruster technologies for diverse space programs and has played an important role in the development and qualification of new thruster designs. The recent emphasis on high-power electric propulsion is pushing the envelope on system performance and service life. Through its Advanced Propulsion Diagnostic Facility, Aerospace is working to ensure that the next generation of electric thruster systems will achieve the envisioned power and efficiency without sacrificing reliability.

History of Electric Propulsion

Development of electric propulsion systems has already spanned more than four decades. After the invention of the gridded ion engine in 1960, many inside and outside the space community believed that the benefits of this technology would soon be realized. With ion propulsion, charged particles are accelerated by passing them through highly charged electrode grids. In theory, a relatively high specific impulse can be achieved, but at the expense of high energy requirements. The early years of frenzied activity included space tests—one even used electricity derived from a nuclear reactor. These tests were not entirely successful, and research continued in the 1970s and 1980s at a slower pace. At the same time, several forms of electric propulsion were developed.

The first among these was the resistojet, a relatively simple form of electric thruster. It operates by passing gaseous propellant, usually hydrazine, over a resistive heater and expanding it through a conventional nozzle.

The resistojet was followed by the arcjet, which passes the hydrazine through an electric arc that heats it before it expands through a nozzle. Developed for flight applications in the 1990s, the hydrazine arcjet offers a more substantial specific impulse boost compared with standard chemical thrusters.

The pulsed-plasma thruster, despite its modest efficiency, found niche applications on spacecraft, thanks to its flexibility and simplicity. In a pulsed-plasma thruster, a capacitor discharge creates a pulsed arc across the face of a block of solid propellant. A small amount of the material is ablated and ionized to form a plasma that is accelerated in a magnetic field.

A Hall-effect thruster in a fiberglass testing chamber.

Additional research focused on magnetoplasmadynamic thrusters, which function by passing a large current through a neutral plasma from a central cathode to an annular anode. The radial current induces a circular magnetic field that accelerates the plasma along the axis of the electrode structure. Operational efficiency is poor if input power is below 50 kilowatts, but nuclear-powered spacecraft may make attractive platforms if other design issues can be resolved.

In the 1990s, work on ion devices such as the Hall-effect thruster intensified. In a Hall-effect thruster, neutral atoms from a heavy gas such as xenon are ionized by collision with high-energy electrons whose movement is confined by a radial magnetic field. The ions are largely unaffected by the magnetic field but are accelerated by the electric field between the anode and cathode. Development of the Hall-effect thruster had come to a stop in the United States, but the former Soviet Union spearheaded an intensive development program that led to extensive flight application.

Ion engines became operational as commercial devices in 1997, with the launch of the Galaxy 11 communications satellite. The success of NASA's Deep Space 1 mission, the first beyond Earth orbit to use an ion engine, further established the viability of this technology. Ion propulsion will be used on the Wideband Gapfiller and Advanced EHF military communication satellites. The first will carry a type of gridded ion engine, and the latter will use Hall-effect thrusters.

Aerospace Role

Aerospace has been a central force in electric thruster test and evaluation since 1989, when the company's Advanced Propulsion Diagnostic Facility became operational. The cylindrical test chamber, 2.4 meters in diameter by 4.8 meters long, was intended to handle gases such as hydrogen and nitrogen (principal products of hydrazine decomposition) that would be exhausted by a resistojet or arcjet. The chamber was equipped with an integrated molecular-velocity analyzer that could quickly obtain the velocity distributions of individual plume species. It was a unique instrument in the electric propulsion community.

Important projects in the early years included a 1-kilowatt simulated hydrazine arcjet designed by NASA. Detailed measurements were made of thrust, plume dissociation fraction, rotational and vibrational temperatures, molecular velocity, and emission characteristics. These measurements were made with various propellants and for multiple operating points. The NASA arcjet development program led to the successful operational use of arcjets, beginning in 1993.

The 1.5 meter diameter by 3.0 meter long test chamber used for near-field and performance studies. The facility is applied to the performance testing of small ion thrusters and development of new diagnostics.

During a three-year period starting in 1992, Aerospace conducted an intensive test and evaluation of a British ion engine, eventually flown on the Artemis communications satellite. The project entailed a minor upgrade of the facility, along with the introduction of various diagnostic techniques, including several that were developed specifically for the project. In addition to quantifying basic electrical and flow parameters, Aerospace was able to evaluate the thrust-vector direction and magnitude, grid deformation during operation, beam divergence, plasma density, plasma potential, electron temperature, ion charge distribution, ion velocity distribution, xenon neutral density, metal erosion rates, ultraviolet and visible emission, radio-frequency and microwave emission, infrared emission, component temperature, microwave phase shift, and surface modification of spacecraft materials. This effort resulted in the most comprehensive set of evaluation tools for an ion engine anywhere in the world and was a vital factor in establishing a baseline for ion propulsion in military communications satellites. These tools were directly applicable to the testing of small ion engines and Hall-effect thrusters for military and commercial programs. During the next few years, Aerospace performed detailed evaluations for various programs and also designed, constructed, and employed a low-power laboratory-model Hall-effect thruster to evaluate engineering trades and to assist in diagnostic development.

During the same period, Aerospace began component-level evaluations and established small supporting facilities for component work and one-off specialized measurements. In the mid 1990s, for example, coherent anti-Stokes Raman scattering was used in a small vacuum chamber to measure the velocity and kinetic temperature of molecular hydrogen both inside and outside the nozzle of an operating resistojet. Since then, Aerospace has installed several smaller test chambers for component work, small thruster efforts, and thruster and diagnostic development. A near-field facility, for example, is applied to the performance testing of small ion thrusters and development of new diagnostics. Aerospace has also devoted considerable effort to thruster hollow-cathode and extraction-grid components and to the study of alternative propellants and novel thruster designs.

Measurement configuration used for contamination and grid-life assessment of T5 (UK-10) ion engine.

A major upgrade of the diagnostic facility in 1999—which doubled the length of the test chamber—enabled Aerospace to perform high-fidelity measurements on medium-power thrusters. Since then, Aerospace has evaluated most of the advanced electric thruster systems in the world. Much of this work has been proprietary to individual customers. In some cases, evaluations have been quite comprehensive, and in the other extreme, limited to one specific measurement result.

Diagnostic Capabilities

Thruster performance, life-limiting characteristics, and interactions between spacecraft and exhaust plumes are best understood through a combination of measurements and modeling. In the case of arcjets in particular, the combination of measurement and modeling has resulted in an unusually complete understanding of many aspects of the device physics and performance.

Modeling can help place measurement data in a framework that lends better predictability for changes in parameters or system design. These models are complex, and typically, the accuracy of one model can have direct bearing on the accuracy of another. For example, numerical plume-propagation models need as inputs the flow properties at the exit plane, which are predicted by a separate model of the propellant acceleration zone. Measurements of near-field plume properties are essential for validating the acceleration-zone model and for controlling the erosion rate of thruster components. Far-field measurements are required for validating the plume-propagation model and assessing interactions with spacecraft materials and sensor payloads.

Contaminant deposition rate measured by a quartz-crystal microbalance in the T5 (UK-10) ion engine plume.

These measurements can be obtained through various methods, depending on the nature of the thruster and plume. Each thruster type has sets of particles with intrinsic velocity, density, and temperature distributions that are determined by complex physical processes. Gas kinetic behavior produces a more diffuse density distribution of lighter particles, as opposed to more massive ones, with a pronounced difference in many cases. Plasma devices generate ions, which are normally fast, and neutral particles, which are normally slow; however, scattering effects produce a degree of homogenization, such that a small percentage of fast ions become slow and slow neutrals become fast (as a result of charge exchange). Scattering also produces ions and neutrals having moderate velocities, and these are directed away from the plume centerline; these need to be considered in spacecraft erosion and contamination models. Slow ions can find themselves in the thruster backflow region, where they can impinge on spacecraft surfaces. Charge transfer can occur between ions and neutral particles of the propellant, and between propellant ions and contaminant particles that were sputtered from thruster components. Densities are always low in backflow regions, where detectability is usually an overriding consideration. The detectability of various species in any region and the measurement of more general properties, as a function of the thruster and the diagnostic employed, is an important consideration.

To address these concerns, Aerospace has developed a comprehensive array of diagnostic capabilities, representing a large investment in equipment and expertise (see table, Diagnostic Capabilities). Within the testing chamber, a wide range of test configurations have been implemented, including movable diagnostic devices and rake-mounted sample holders. Small movable probes are suspended on a rotating arm that samples the plume over a 360-degree range at a distance of up to 1.1 meters, extending to a larger radius if the angular range is reduced. View ports allow access for laser-induced fluorescence and video cameras. Beam profiling, thrust-vector tracking, and spatially resolved laser-induced fluorescence are performed by mounting the thruster on a multiple-axis microstepper positioning system, which is temperature controlled to counteract the radiative cooling effect of nearby cryopanels.

The typical configuration for laser-based diagnostics. Beam profiling, thrust-vector tracking, and spatially resolved laser-induced fluorescence are performed by mounting the thruster on a multiple-axis microstepper positioning system, which is temperature-controlled to counteract the radiative cooling effect of nearby cryopanels.

Near-field plume measurements are performed with a fast intrusive probe near the exit plane. The ion current is collected with a long, electrically biased wire that crosses through the plume. The collected data are converted into flux contour maps. Angle-resolved laser-induced fluorescence is used to measure the longitudinal and azimuthal velocity of neutral particles and ions. Fluorescence of single-charge xenon ions or other suitable species allows a determination of the translational temperature and the most probable velocity vector over a grid of measurement points close to the exit plane. Fluorescence of plume metals, such as low-density grid-sputtered molybdenum, generates contaminant density and velocity maps. Laser absorption determines absolute column-averaged density for suitable species of sufficient abundance. Two-photon laser-induced fluorescence measurement using a 225-nanometer excitation wavelength is used to map density of neutral xenon and hydrogen atoms and, with modeling, determine propellant utilization efficiency.

Far-field measurements have been made using a Faraday probe to determine the flux-versus-angle and a time-of-flight parallel-plate electrostatic deflector to determine energy and xenon charge distributions. Electron density measurements in the far field are performed using Langmuir probes and radio-frequency resonator probes. Measurements of plume optical radiation and electromagnetic compatibility can be tailored to support specific requirements.

Angle-resolved laser-induced fluorescence (LIF) is used to measure the longitudinal and azimuthal velocity of neutral particles and ions. Fluorescence of single-charge xenon ions or other suitable species allows a determination of the translational temperature and the most probable velocity vector over a grid of measurement points close to the exit plane.

To supplement conventional surface-effect tests that use spacecraft material coupons arrayed in the far-field plume, Aerospace researchers determine the sample deposition or erosion rate as a function of angle using several temperature-controlled collimated quartz-crystal microbalances at set distances. Another bulk property of interest is the accommodated heat flux caused by plume impingement, which is measured using an instrumented copper disk coated with a reference material. Plume heat flux can be evaluated as a function of the angle of incidence on the probe surface and as a function of the probe position relative to thruster centerline.

A Case Study

Ion flux vs. angle for the BPT-4000 Hall current thruster at steady-state conditions. Flux scans at 100-centimeter radius were measured with a retarding potential analyzer. A high degree of symmetry is evident about the thruster's physical centerline at 0 degrees. Within 35 degrees from centerline, the flux is predominantly fast ions. Beyond 35 degrees, the contributions from ion-neutral elastic scattering and from charge-exchange production of slow ions become more important, yielding the wings on the flux curves (view larger image).

The Aerojet BPT-4000 is a 4.5-kilowatt xenon Hall-effect thruster that will be used for orbit insertion, orbit maintenance, and repositioning of geosynchronous satellites such as Advanced EHF. A requirement of the BPT-4000 flight-qualification program is to demonstrate that the thruster will survive repeated cycles from the minimum on-orbit temperature to the maximum steady-state temperature. Another requirement is to measure the thrust-vector angle from the physical centerline during startup and at steady state.

Aerospace performed thermal cycling and thrust-vector alignment tests. The thruster was placed in a copper shroud cooled by liquid nitrogen to simulate the orbital environment. The shroud door was closed for thruster cooling and opened just before the start of each firing cycle. An aluminum mounting bracket fitted with cartridge heaters maintained the desired interface temperature. The thruster operated nominally through 10 thermal cycles starting at the minimum expected on-orbit temperature and finishing in 3.5–4 hours at the hot steady-state condition. The shroud returned the thruster to the cold steady state in 9–10 hours. Temperature variability between cycles was minimal.

Angle-dependent ion flux scans at 100-centimeter radius were measured with a retarding potential analyzer. All scans displayed a high degree of symmetry about the thruster's physical centerline at 0 degrees. Within 35 degrees from centerline, the flux was predominantly fast ions. Beyond 35 degrees, the contributions from ion-neutral elastic scattering and from charge-exchange production of slow ions became more important, as evident in the flux curves. A high degree of symmetry about the centerline was observed at all operating points, and none of the measured thrust vector angles exceeded 0.7 degrees. Steady-state angles varied by no more than 0.2 degrees between operating points, with no clear dependence on discharge power or voltage. Thrust vector motion was typically 0.25 degrees during the first two hours of operation and 0.05 degrees during the second two hours. Based on the reproducibility between cycles, the random error in the thruster vector measurements was plus or minus 0.05 degrees.

EMC Test and Evaluation

Aerospace also tested the BPT-4000 thruster for electromagnetic compatibility—a particular area of expertise. Electromagnetic compatibility measurements consider four general categories: radiated emission through space, conducted emission onto the bus, susceptibility to radiation fields, and susceptibility to injected currents. Extensive measurements are carried out in all four areas, following military standard MIL-STD-461.

Emission from a BPT-4000 Hall-effect thruster operating at 4.5 kilowatts with discharge voltages of 300 and 400 volts. The frequency span shown (10 kilohertz to 18 gigahertz) was measured using four broadband antennas and a spectrum-analyzer-based receiver controlled by Aerospace-developed software. An increase of 20 decibels is equivalent to a 10X increase in measured electric field. Spacecraft EMC limits are payload specific and generally proprietary; the MIL-SPEC limit is shown for comparison (view larger image).

Electron plasma frequency emission from a BPT-4000 Hall-effect thruster (view larger images).

Hall-effect thrusters and gridded ion engines support complex plasma oscillations that can emit electromagnetic radiation from dc to frequencies above 20 gigahertz. These emissions can be quite strong, often exceeding MIL-STD-461 specifications for frequencies below 4 gigahertz. The low-frequency emissions may induce currents in satellite structures or cause problems directly with electronic components, while high-frequency emissions may interfere with communication channels. Additionally, dc and ac magnetic fields may affect operation of sensitive instruments that are part of the satellite payload. Selective shielding, repositioning of antennas, and modifications of operational procedures may be required to successfully integrate thruster and spacecraft.

The Aerospace approach to radiative emission measurements involves a closed cylinder of dielectric material—largely transparent to electromagnetic radiation—that houses the thruster and mates to the large vacuum chamber. This cylinder is surrounded by a semi-anechoic room, which shields the measurement space from the ambient electromagnetic fields and reduces the reflections of the thruster radiation from the metallic walls of the room. The plume of the thruster exhausts through a carbon-fiber grid into the main vacuum tank, terminating on a beam dump comprising a series of carbon-covered pyramids. The grid reduces the background radiation that leaks from the main chamber, and the pyramids reduce sputtering by the high-energy xenon ions and scattering of electromagnetic radiation by the metal walls of the vacuum tank.

This arrangement allows antennas to be placed 1 meter from the thruster, as required by MIL-STD-461, and remain unexposed to plasma or metallic surface reflections. Residual facility effects are measured from antenna response at various positions, with calibrated-emission transmitters at the location of the thruster.

Radiative Hall-effect thruster emissions at frequencies below a few hundred megahertz are largely understood, and many of the associated oscillations are required for the proper operation of the thruster. Emissions above 18 gigahertz are primarily caused by electron plasma oscillations. The strong emission seen in the 1–8-gigahertz range (the L, S, and C communication bands) exhibits complex temporal and spatial characteristics and is not currently understood. Thruster-to-thruster variations in the L, S, and C band and dependence on thruster age are also unknown, and are subjects of active research.

Looking to the Future

The ability to operate medium-power thrusters and apply a comprehensive suite of precision diagnostics has made Aerospace an important independent test and evaluation resource. Given the direction of development toward high-powered spacecraft, including those with nuclear electric power sources, Aerospace will soon need to perform high-fidelity characterizations of much more powerful thrusters. Continued facility upgrades will therefore be necessary.

While some of the more sophisticated forms of electric propulsion have finally entered service, refinement of current designs and progress toward higher power devices continues at a rapid pace. Ion propulsion use by NASA and the military is still in the initial phase. Many issues requiring detailed test and evaluation are being addressed for a variety of thruster systems and flight programs, and Aerospace will continue to play a central role in this work.