John R. Anderson

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.

Numerical simulations of ion thruster accelerator grid erosion

The highly successful demonstration of ion propulsion on Deep Space 1 has stimulated the study of more demanding applications of this technology. These future applications require ion thrusters capable of providing significantly greater specific impulses and total impulses than the current state-of-the-art. Higher specific impulses aggravate the known wear out mechanisms of the ion accelerator system. Computer simulations of the ion accelerator system operation and erosion are essential tools for the development of ion thrusters to meet the demand for higher specific impulse and longer life. Two-dimensional and three-dimensional computer codes have been developed at JPL and are used to provide insight into the processes limiting the life of the accelerator grid. The 2D code described herein was used to identify a key feature of operation at high Isp. That is, as the beam voltage is increased the energy of the charge-exchange ions hitting the hole walls increases in proportion to the beam voltage and more rapidly than the increase in the magnitude of the accelerator grid voltage. This has serious implications for the design of long-life, high specific impulse thrusters.

Validation of the NSTAR Ion Propulsion System on the Deep Space One Mission: Overview and Initial Results

Deep Space 1 is the first interplanetary spacecraf t to use an ion propulsion system for the p r imary delta-v maneuvers. The purpose of the mission is to validate a number of technologies, including ion propulsion and a high degree of spacecraft autonomy, on a fiyby of an asteroid and two comets. The ion propulsion system has operated now for a total of 1791 hours at engine power levels ranging from 0.48 to 1.94 kW and has completed the deterministic thrusting requi red for an encoun te r w i th the asteroid 1992KD in late July, 1999. The system has worked extremely well after an initial grid short was cleared after launch. O p eration during this primary mission phase has demonstrated all ion propuls ion system and au tonomous navigation functions. All propulsion system operating parameters are ve ry close to the expected values wi th the excep tion of the thrust at higher power levels, which is about 2 percent lower than calculated values. This paper provides an overview of the system and presents the first flight validation data on an ion propulsion system i n interplanetary space.

Numerical Simulation of Two-Grid Ion Optics Using a 3D Code

A three-dimensional ion optics code has been developed under NASA’s Project Prometheus to model two grid ion optics systems. The code computes the f low of positive ions from the discharge chamber through the ion optics and into the beam downstream of the thruster. The rate at which beam ions interact w ith background neutral gas to form charge exchange ions is also computed. Charge exchange ion trajectories are computed to determine where they strike the i o n optics grid surfaces and to determine the extent of sputter erosion they cause. The code has been used to compute predictions of the erosion pattern and wear rate on the NSTAR ion optics system; the code predicts the shape of the eroded pattern but overestimates the initial wear rate by about 50%. An example o f use of the code to estimate the NEXIS thruster accelerator grid life is a lso presented.

Results of an On-Going Long Duration Ground Test of the DS1 Flight Spare Engine

Ground testing of the DS1 night spare thruster (FT2) is presently being conducted. To date, the thruster has accumulated over 4500 hours of operation. Comparison of FT2 with the performance of the engineering model thruster 2 (EMT2) during the 8.2 khr test shows a transient, lasting for about 3000 hours, during which the discharge chamber efficiency decreases for both thrusters. The flow rates are 2% lower for FT2 than for EMT2 and the discharge chamber performance is 4.5% lower for FT2 during the transient. Sensitivity data obtained during the test show that the lower flow rate accounts for about half of the observed difference. After the initial transients decay, the performance of both thrusters is comparable with the exception of the electron backstreaming margin--which is 6 V lower for FT2.

Environmental Testing of the NEXT PM1R Ion Engine

Abstract The NEXT propulsion system is an advanced ion propulsion system presently under development that is oriented towards robotic exploration of the solar system using solar electric power. The subsystem includes an ion engine, power processing unit, feed system components, and thruster gimbal. The Prototype Model engine PM1 was subjected to qualification-level environmental testing in 2006 to demonstrate compatibility with environments representative of anticipated mission requirements. Although the testing was largely successful, several issues were identified including the fragmentation of potting cement on the discharge and neutralizer cathode heater terminations during vibration which led to abbreviated thermal testing, and generation of particulate contamination from manufacturing processes and engine materials. The engine was reworked to address most of these findings, renamed PM1R, and the environmental test sequence was repeated. Thruster functional testing was performed before and after the vibration and thermal-vacuum tests. Random vibration testing, conducted with the thruster mated to the breadboard gimbal, was executed at 10.0 G

Performance characteristics of the deep space 1 flight spare ion thruster long duration test, the first 21,300 hours of operation

A long duration test of the DSl flight spare ion thruster (FT2) is presently being conducted at the Jet Propulsion Laboratory. To, date the thruster has accumulated over 23,500 hours of operation, and 190 kg of Xenon propellant, over 230% of the initial design life. The primary objectives of the test include the processing of 200 kg of Xenon propellant, the identification of unknown failure modes, the characterization and drivers of these failure modes, and to measure performance degradation as the thruster wears. The test is fitted with an extensive array of diagnostics to measure engine wear and performance degradation. To date the most notable erosion processes include severe discharge cathode keeper erosion, accelerator grid erosion, reduction in electrical isolation of the neutralizer assembly, and deposit formation within the neutralizer orifice, reducing margin from plume mode. Over the past 23,500 hours of operation, performance degradation has been minimal, and it is anticipated that the above erosion processes will not preclude the thruster from processing over 200 kg of Xenon.

Thermal Development Test of the NEXT PM1 Ion Engine

NASA’s Evolutionary Xenon Thruster (NEXT) is a next-generation high-power ion propulsion system under development by NASA as a part of the In-Space Propulsion Technology Program. NEXT is designed for use on robotic exploration missions of the solar system using solar electric power. Potential missio n destinations that could benefit from a NEXT Solar Electric Propulsion (SEP) system include inner planets, small bodies, and outer planets and their moons. This range of robotic expl oration missions generally calls for ion propulsion systems with deep throttling capability and system input power ranging from 0.6 to 25 kW, as referenced to solar array output at 1 Astronomical Unit (AU). Thermal development testing of the NEXT prototype model 1 (PM1) was conducted at JPL to assist in developing and validating a thruster thermal model and assessing the thermal design margins. NEXT PM1 performance prior to, during and subsequent to thermal testing are presented. Test results are compared to the predict ed hot and cold environments expected missions and the functionality of the thruster for these missions is discussed.

Qualification of Commercial XIPS ® Ion Thrusters for NASA Deep Space Missions

‡‡ Electric propulsion systems based on commercial ion and Hall thrusters have the potential for significantly reducing the cost and schedule-risk of Ion Propulsion Systems (IPS) for deep space missions. The large fleet of geosynchronous communication satellites that use SEP, which will approach 40 satellites by year-end, demonstrates the significant level of technical maturity and spaceflight heritage achieved by the commercial electric propulsion systems. A program to delta-qualify XIPS ® ion thrusters for deep space missions is underway at JPL. This program includes modeling of the thruster grid and cathode life, environmental testing of a 25-cm EM thruster over anticipated vibe and thermal requirements for deep space missions, and wear testing of the thruster cathodes to demonstrate the life and benchmark the model results. This paper will present the deltaqualification status of the XIPS thruster and discuss the life and reliability with respect to known failure mechanisms.

A THREE-DIMENSIONAL NUMERICAL SOLUTION FOR THE SHAPE OF A ROTATIONALLY DISTORTED POLYTROPE OF INDEX UNITY

We present a new three-dimensional numerical method for calculating the non-spherical shape and internal structure of a model of a rapidly rotating gaseous body with a polytropic index of unity. The calculation is based on a finite-element method and accounts for the full effects of rotation. After validating the numerical approach against the asymptotic solution of Chandrasekhar that is valid only for a slowly rotating gaseous body, we apply it to models of Jupiter and a rapidly rotating, highly flattened star (α Eridani). In the case of Jupiter, the two-dimensional distributions of density and pressure are determined via a hybrid inverse approach by adjusting an a priori unknown coefficient in the equation of state until the model shape matches the observed shape of Jupiter. After obtaining the two-dimensional distribution of density, we then compute the zonal gravity coefficients and the total mass from the non-spherical model that takes full account of rotation-induced shape change. Our non-spherical model with a polytropic index of unity is able to produce the known mass of Jupiter with about 4% accuracy and the zonal gravitational coefficient J 2 of Jupiter with better than 2% accuracy, a reasonable result considering that there is only one parameter in the model. For α Eridani, we calculate its rotationally distorted shape and internal structure based on the observationally deduced rotation rate and size of the star by using a similar hybrid inverse approach. Our model of the star closely approximates the observed flattening.