Charles E. Garner

An Overview of the Results from the 30,000 Hr Life Test of Deep Space 1 Flight Spare Ion Engine

The extended life test of the Deep Space 1 (DS1) spare flight ion thruster (FT2) was voluntarily terminated on June 26th 2003. Although the engine had not yet reached its end of life at the conclusion of the test, the decision to terminate and begin the post-test destructive analyses was made to benefit near term ion engine development programs. During its 5-year run, the thruster operated for a total of 30,352 hours, processed 235.1 kg of Xenon propellant, and acquired several thousand hours of operation at each of the four independent throttled conditions investigated. The objectives of the test were to characterize previously observed failure modes, identify unknown failure modes, and quantify thruster performance as a function of engine wear and throttle level. Several performance variations and degradation processes were observed and monitored during the course of the test. Degradation processes included erosion of the discharge cathode keeper, ion optics grid sputter erosion, and deposition of material in the neutralizer cathode at low power that later cleared with the return to full power operation. Performance degradation was limited to reduction in measured thrust at the full power point for the final 1000 hours of operation, most likely due to electronbackstreaming. Post-test inspection of the engine was initiated immediately following the test termination, to ascertain causes of the wear, and to look for any previously unknown wear processes. The ion engine consists of various internal systems, and the post-test analysis effort has been divided into separate analysis efforts of the ion optics system, discharge and neutralizer cathode assemblies, and the discharge chamber. Post-test inspection of the ion optics system revealed significant sputter erosion of the accelerator grid and measurable erosion of the screen grid. Post-test inspection results include the presence of through pits in the accelerator grid webbing, but no formation of rogue holes. Inspection of the discharge cathode indicates significant cathode orifice plate sputter erosion, as a result of exposure to the discharge plasma following removal of the keeper, but no erosion of or deposition near the orifice itself. Inspection of the neutralizer revealed no erosion to the keeper, orifice plate, or heater, and an orifice free of the deposits that were previously observed during the minimum power test segment. Post-test inspection indicates the discharge chamber experienced no measurable sputter erosion, although a substantial number of loose molybdenum coated carbon flakes were present on the sputter containment mesh. It is believed that the bulk of the flakes are due to back-sputtered beam target material, subsequently coated by sputter eroded grid material, a facility induced effect. A summary of the BOL and EOL performance, results of the system level inspection, an implication of findings to the ultimate service life capability of the 30-cm technology are presented.

The Ion Propulsion System For Dawn

Summary The Dawn mission will be the first use of ion propulsion on a full up science mission for NASA. The ion propulsion system for Dawn is based on that demonstrated on Deep Space 1 with modifications necessary to accommodate multiple thrusters, to make the system single fault tolerant, to reduce the mass of the mechanical gimbals, and to accommodate a much larger propellant load. To rendezvous with the two heaviest main-belt asteroids, Vesta and Ceres, Dawn will carry 450 kg of xenon, whereas DS 1 carried only 8 1.5 kg of xenon. Acknowledgements The work described in this paper was conducted, in part, by the Jet Propulsion Laboratory, California Institute of Technology, under contract to the National Aeronautics and Space Administration. References 1. 2. 3. 4. 5. 6. M. D. Rayman, P. Varghese, D. H. Lehman, and L. L. Livesay, “Results From The Deep Space 1 Technology Validation Mission,’D IAA-99- IAA. 1 1.2.01, Presented at the 50th International Astronautical Congress, Amster#am, The Netherlands, 4-8 October, 1999, Acta Astronautica 47, p. 475 (2000). M. D. Rayman and P. Varghese, “The Deep Space 1 Extended Mission,” Acta Astronautica Brophy, J. R., et al., “Ion Propulsion System (NSTAR) DS 1 Technology Validation Report,” JPL Publication 00-1 0, October 2000. J. E. Polk, et al., “Validation of the NSTAR Ion Propulsion System on the Deep Space One Mission: Overview and Initial Results,” AIAA 99-2274, presented at the 35th Joint Propulsion Conference and Exhibit, 20-24 June 1999, Los Angeles, California. J. E. Polk, et al., “Demonstration of the NSTAR Ion Propulsion System on the Deep Space One Mission, IEPC-01-075, Presented @t the 27th International Electric Propulsion iconference, Pasadena, CA, 15- 19 October, 2001. Brophy,

Tests of high current hollow cathodes for ion engines

The development of high power, high thrust ion engines requires the development of long life hollow cathodes capable of producing emission currents of hundreds of amperes. This paper describes results obtained through the extended testing of two 12.7 mm diameter hollow cathodes at high emission currents. The first cathode, operated on xenon for a total of 170 hours at an emission current of 100 A, sustained very little damage except for an increase in the cathode orifice diameter, from 0.953 to 1.71 mm. The second cathode was operated on xenon at 150 A emission current continuously for 24 hours, and for over 1000 hours at 100 A on argon with essentially no change in the cathode operating characteristics. Operation of this cathode was terminated due to severe erosion of several discharge chamber components located downstream of the cathode. The erosion of these components occurred despite operation at discharge voltages less than 24 volts, and poses a serious challenge to the sucessful implementation of multi-hundred ampere cathodes and high thrust ion engines.

Dawn Ion Propulsion System - Initial Checkout After Launch

The first 80 days after launch of the Dawn mission were dedicated to the checkout of the spacecraft with a major emphasis on the ion propulsion system. All three ion thrusters, all three thruster-gimbal assemblies, both power processor units, both digital interface and control units, and the entire xenon feed system were completely checked out, and every component was found to be in good health. Direct thrust measurements agreed well with preflight expected values for all three thrusters over the entire throttle range. Measurements of the thruster-produced roll- torque verified that each thruster produced less than the maximum allowed value of 60 μNm at full power. Thruster electrical operating parameters and power processor unit efficiencies also agreed well with preflight expected values based on acceptance test data. Two of the three ion thrusters were fully checked out within 30 days after launch. Checkout of all three thrusters was completed 64 days after launch. Deterministic thrusting with the ion propulsion system began on 17 December 2007.

Fabrication and testing of carbon-carbon grids for ion optics

Ion optics measuring 16.5 cm in diameter and 1.0 mm in thickness were fabricated from carbon-carbon composites that were woven from a high-tensile-modulus carbon fiber. Plate flatness varied by less than 0.05 mm. Several methods were investigated for forming ion-extraction apertures in the carbon-carbon plates, including laser machining, mechanical drilling, and conventional electric discharge machining. Tests conducted using a quartz dilatometer indicated that the coefficient of thermal expansion of the carbon-carbon plates varied between -0.51 to -1.8 x 10 exp -6/degree C at plate temperatures between 173-773 K. Sputter-erosion experiments indicate that carbon-carbon erodes at a rate approximately 25 percent below molybdenum under the same conditions. These material properties indicate that carbon-carbon may be superior to molybdenum for use as ion optics electrodes for ion engines.

Deep Space 1 Flight Spare Ion Thruster 30,000-Hour Life Test

The extended-life test of the Deep Space 1 flight spare ion thruster was voluntarily terminated on 26 June 2003. During its five-year run, the thruster operated for a total of 30,352 h, processed 235.1 kg of xenon propellant, and demonstrated extended operation at multiple throttled conditions. The objectives of the test were to characterize failure modes and quantify thruster performance as a function of engine wear and throttle level. Degradation processes included erosion of the discharge cathode keeper, accelerator-grid sputter erosion, and deposition of material in the neutralizer cathode at low power. Performance degradation was limited to a reduction in measured thrust at the full-power point for the final 1000 h of operation. Posttest inspection of the enginewas initiated following the test termination to ascertain causes of the wear and to look for any previously unknown wear processes. Significant findings included facility-induced flakes in the discharge chamber, the presence of through-pits in the accelerator-grid webbing, significant erosion of the discharge cathode orifice plate, and healthy cathode inserts. A summary of the beginning-of-test and end-of-test performances and results of the posttest destructive evaluation are presented.

Ion engine endurance testing at high background pressures

Ion engine endurance testing at vacuum chamber pressures in the low 10 exp -3 Pa range is enabled through the use of a three-grid accelerator system with the decelerator grid biased 50 to 100 volts negative of neutralizer cathode potential. The negative decelerator grid serves to collect the facility induced charge exchange ion current which normally results in rapid erosion of the accelerator grid during testing at elevated vacuum chamber pressures. This screen, accelerator, negative decelerator (SAND) grid configuration enables an order of magnitude reduction in vacuum chamber pumping speeds relative to that required for endurance testing of ion engines with conventional two-grid accelerator systems. A 900-hr test of a 30-cm diameter engine at 6.5 kW and a tank pressure of 3.7 x 10 exp -3 Pa was performed to test the feasibility of the three-grid SAND accelerator system technology. Grid erosion rates from this test are compared to those from a 200-hr test performed with the same discharge chamber, in the same test facility, and at the same background pressure with a conventional two grid accelerator system. The SAND optics resulted in greater than a factor of 100 reduction in accelerator grid erosion rate relative to the two-grid system.

Methods for Cryopumping Xenon

Design chamclcris(ics ofacryopun~p o~)ti\~]i~,d forxc]]on arc])rcscl)tcd. Sin.glc-stagc Gifford-McMahon hcliun~ cryoprrnrps that])rovidc acryc)rcfrigcratiol) capability of105Wat 50Ktochill coldplatcs arcuscd witioulbafflcsto cnablclargcpurnp spedsonxcn on. Shrouds arcusti torducc ticthcrt]lal loadtodlc cold ]~latcto cnablc. gr~tcrxcl~onpun)p spcrds, but it is notrcquircd tooWratc thcshrouds at1,N2tct))pcra( urcs,Ton ]axin]i~.cp ul])l)s lwtiorlx erlongascss uc.hasnitrogcn, neon and helium can not be cryotrappcd with these pumps. q’t)ccryo1)u~~)psarc52ct~] in length, 16cn~in diarmtc.randwcigh 15.4 Kg, however the cold head assembly is only 29 cm long and 8,3 cm in diameter. The cryorcfrigcralors can be mourr(cd using a 1 S-cm-dia ftangc, and a 10-cm-dia through-hole in the vacuum c}]amber. Alternatively, the cnlirc. cryopump can be mounted inside the vacuum chamber, although preliminary data indic.am that pump efficiency dccrcascs when the pumps arc operalcd in vacuum. I’cst

A 5,000 hour xenon hollow cathode life test

indicate that the full lhcorctical pump speed on xenon is achic.vcd, A pumping systcm designed for the 1,000 hr NSTAR validation test was opaatcd without shrouds or baffles. A poor] y controlled thermal radia[ion load 10 the cryopancls and uncxjxx%.xlly low cryopancl cocfficicnl of thermal conductivity rcsrrllcd in pressure spikes and a drop in the xenon pump speed. A cryopurnping syslcm u[ilixing three cryorcfrigcrators wilh a chilled shroud but without baffles is being used on (hc 8,000-hr NSTAR wear lest to bcticr control the thermal cnvironmcmt Each of [IICSC xenon cryopumps provides a pump speed on xenon of approximately 15,000 1,/s. The, small si~.c of the pumps, coupled with their abil ity to he opcramd in vacuum, provide the user with a grca[ deal of vcrsalili(y as far as locating ihe pumps with rcspca to thrus[crs, tank walls, shrouds, CIC. The COS1 and comp]cxity of installing and operating these cryopumps is a fraclion of the. cost and complexity to install and opera[c diffusion pumps or cryopumps thal deliver sitnilar pump spct.ds on xenon.

A solar sail design for a mission to the near-interstellar medium

A cathode life test voluntary terminated after 5024 hours of operation at 25 A is reported. The cathode including a 6.35-mm diameter by 57.12-mm long molybdenum tube with a nominal wall thickness of 0.38 mm is described along with a test facility and start-up procedure. It is shown that over the time of the experiment, the cathode-orifice diameter eroded from 1.80 mm to 2.08 mm, which corresponds to a ratio of the emission current to the orifice diameter at the end of the test of 12.0 A/mm. Tungsten deposits on the interior surface of the insert are detected in post-test analyses, and complete depletion of the original impregnate is suggested by X-ray diffraction analyses. A cathode-jet phenomenon in which energetic ions are produced during hollow-cathode operation at emission currents above 20 A is confirmed.