ALBERT

Description: The Deep Space 1 (DSl) spare flight thruster (FT2) was operated for 30,352 hours during the extended life test (ELT). The test was performed to validate the service life of the thruster, study known and identify unknown life limiting modes. Several of the known life limiting modes involve the ion optics system. These include loss of structural integrity for either the screen grid or accelerator grid due to sputter erosion from energetic ions striking the grid, sputter erosion enlargement of the accelerator grid apertures to the point where the accelerator grid power supply can no longer prevent electron backstreaming, unclearable shorting between the grids causes by flakes of sputtered material, and rouge hole formation due to flakes of material defocusing the ion beam. Grid gap decrease, which increases the probability of electron backstreaming and of arcing between the grids, was identified as an additional life limiting mechanism after the test. A combination of accelerator grid aperture enlargement and grid gap decrease resulted in the inability to prevent electron backstreaming at full power at 26,000 hours of the ELT. Through pits had eroded through the accelerator grid webbing and grooves had penetrated through 45% of the grid thickness in the center of the grid. The upstream surface of the screen grid eroded in a chamfered pattern around the holes in the central portion of the grid. Sputter deposited material, from the accelerator grid, adhered to the downstream surface of the screen grid and did not spall to form flakes. Although a small amount of sputter deposited material protruded into the screen grid apertures, no rouge holes were found after the ELT.

Description: During the Extended Life Test of the DS1 flight spare ion thruster, the engine was subjected to sensitvity testing in order to characterize the macroscopic dependence of discharge chamber sensitivity to a +\-3% vatiation in main flow, cathode flow and beam current, and to +\5% variation in beam and accelerator voltage, was determined for the minimum- (THO), half- (TH8) and full power (TH15) throttle levels. For each power level investigared, 16 high/low operating conditions were chosen to vary the flows, beam current, and grid voltages in in a matrix that mapped out the entire parameter space. The matrix of data generated was used to determine the partial derivative or senitivity of the dependent parameters--discharge voltage, discharge current, discharge loss, double-to-single-ion current ratio, and neutralizer-keeper voltage--to the variation in the independent parameters--main flow, cathode flow, beam current, and beam voltage. The sensititivities of each dependent parameter with respect to each independent parameter were determined using a least-square fit routine. Variation in these sensitivities with thruster runtime was recorded over the duration of the ELT, to detemine if discharge performance changed with thruster wear. Several key findings have been ascertained from the sensitivity testing. Discharge operation is most sensitve to changes in cathode flow and to a lesser degree main flow. The data also confirms that for the NSTAR configuration plasma production is limited by primary electron input due to the fixed neutral population. Key sensitivities along with their change with thruster wear (operating time) will be presented. In addition double ion content measurements with an ExB probe will also be presented to illustrate beam ion production and content sensitivity to the discharge chamber operating parameteres.

Description: Interest in science objectives at the outer planets, specifically at the moons of Jupiter, has spurred the development of high-power electric propulsion systems under the Prometheus program.

Description: Destructive physical analysis of the discharge and neutralizer hollow cathode assemblies from the Deep Space 1 Flight Spare 30,000 Hr life test was performed to characterize physical and chemical evidence of operationally induced effects after 30,372 hours of operation with beam extraction. Post-test inspection of the discharge-cathode assembly was subdivided into detailed analyses at the subcomponent level. Detailed materials analysis and optical inspection of the insert, orifice plate, cathode tube, heater, keeper assembly, insulator, and low-voltage propellant isolator were performed. Energy dispersive X-ray (EDX) and scanning electron microscopy (SEW analyses were used to determine the extent and composition of regions of net deposition and erosion of both the discharge and neutralizer inserts. A comparative approach with an un-operated 4:1:1 insert was used to determine the extent of impregnate material depletion as a function of depth from the ID surface and axial position from the orifice plate. Analysis results are compared and contrasted with those obtained from similar analyses on components from shorter term tests, and provide insight regarding the prospect for successful longer-term operation consistent with SOA ion engine program life objectives at NASA.

Description: A series of experimental measurements on a modified laboratory NSTAR engine were used to validate a zero dimensional analytical discharge performance model of a ring cusp ion thruster. The model predicts the discharge performance of a ring cusp NSTAR thruster as a function the magnetic field configuration, thruster geometry, and throttle level. Analytical formalisms for electron and ion confinement are used to predict the ionization efficiency for a given thruster design. Explicit determination of discharge loss and volume averaged plasma parameters are also obtained. The model was used to predict the performance of the nominal and modified three and four ring cusp 30-cm ion thruster configurations operating at the full power (2.3 kW) NSTAR throttle level. Experimental measurements of the modified engine configuration discharge loss compare well with the predicted value for propellant utilizations from 80 to 95%. The theory, as validated by experiment, indicates that increasing the magnetic strength of the minimum closed reduces maxwellian electron diffusion and electrostatically confines the ion population and subsequent loss to the anode wall. The theory also indicates that increasing the cusp strength and minimizing the cusp area improves primary electron confinement increasing the probability of an ionization collision prior to loss at the cusp.

Description: The Nuclear Electric Xenon Ion System (NEXIS) ion thruster was developed for potential outer planet robotic missions under NASA's Prometheus program. This engine was designed to operate at power levels ranging from 16 to over 20 kWe at specific impulses of 6000 to 7500 s for burn times of up to 10 years, satisfying the requirements of nuclear electric propulsion systems such as that on the proposed Prometheus 1 mission to explore the icy moons of Jupiter. State-of-the-art performance and life assessment tools were used to design the thruster. Following the successful performance validation of a Laboratory Model thruster, Development Model hardware was fabricated and subjected to vibration and wear testing. The results of a 2000-hour wear test are reported herein. Thruster performance achieved the target requirements and was steady for the duration of the test. Ion optics performance was similarly stable. Discharge loss increases of 6 eV/ion were observed in the first 500 hours of the test and were attributed to primary electron energy decreases due to cathode insert conditioning. Relatively high recycle rates were observed and were identified to be high-voltage-to-ground arcs in the back of the thruster caused by wire insulation outgassing and electron penetration through the plasma screen. Field emission of electrons between the accelerator and screen grids was observed and attributed to evolution of field emitter sites at accelerator grid aperture edges caused by ion bombardment. Preliminary modeling and analysis indicates that the NEXIS engine can meet mission performance requirements over the required lifetime. Finally, successful validation of the NEXIS design methodology, design tools, and technologies with the results of the wear test and companion performance and vibration tests presents significant applicability of the NEXIS development effort to missions of near-term as well as long-term interest for NASA.

Description: In gridded electrostatic thrusters, ions are produced by electron bombardment in the discharge chamber. In most of these thrusters, a single, centrally located hollow cathode supplies the ionizing electrons. An applied magnetic field in the discharge chamber restricts the electrons leaving the hollow cathode to a very narrow channel. In this channel, the high electron current density ionizes both propellant gas flowing from the hollow cathode, and other neutrals from the main propellant flow from the plenum. The processes that occur just past the hollow cathode exit are very important. In recent engine tests, several cases of discharge cathode orifice place and keeper erosion have been reported. In this paper we present results from a new 1-D, variable area model of the plasma processes in the magnetized channel just downstream of the hollow cathode keeper. The model predicts plasma densities, and temperatures consistent with those reported in the literature for the NSTAR engine, and preliminary results from the model show a potential maximum just downstream of the cathode.