Search NASA⌕ Search

SEARCH · Search NASA

Results for “gridded ion thruster”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS2008

Qualification of Commercial XIPS(R) 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 solar electric propulsion (SEP), which will approach 40 satellites by year-end, demonstrates the significant level of technical maturity and spaceflight heritage achieved by the commercial IPS systems. A program to delta-qualify XIPS(R) 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-centimeter electromagnetic (EM) thruster over DAWN-like vibe and temperature profiles, and wear testing of the thruster cathodes to demonstrate the life and benchmark the model results. This paper will present the delta-qualification status of the XIPS thruster and discuss the life and reliability with respect to known failure mechanisms.

electric propulsion↗
NASA NTRS1984

Ion thruster performance model

A model of ion thruster performance is developed for high flux density, cusped magnetic field thruster designs. This model is formulated in terms of the average energy required to produce an ion in the discharge chamber plasma and the fraction of these ions that are extracted to form the beam. The direct loss of high energy (primary) electrons from the plasma to the anode is shown to have a major effect on thruster performance. The model provides simple algebraic equations enabling one to calculate the beam ion energy cost, the average discharge chamber plasma ion energy cost, the primary electron density, the primary-to-Maxwellian electron density ratio and the Maxwellian electron temperature. Experiments indicate that the model correctly predicts the variation in plasma ion energy cost for changes in propellant gas (Ar, Kr and Xe), grid transparency to neutral atoms, beam extraction area, discharge voltage, and discharge chamber wall temperature. The model and experiments indicate that thruster performance may be described in terms of only four thruster configuration dependent parameters and two operating parameters. The model also suggests that improved performance should be exhibited by thruster designs which extract a large fraction of the ions produced in the discharge chamber, which have good primary electron and neutral atom containment and which operate at high propellant flow rates.

Brophy, J. R.↗
NASA NTRS1971

Design and development of a small structurally integrated ion thruster system

A 5 cm structurally integrated ion thruster has been developed for attitude control and stationkeeping of synchronous satellites. With 2-D vectorable grids, thrust is 0.56 mlb at a beam voltage of 1200 V, total mass efficiency is 64%, and electrical efficiency is 46.8%. Nonvectoring thrust of 0.41 is demonstrated at very low specific impulse with dielectric coated grids. Structural integrity is demonstrated with dielectric coated grids for shock (30 G), sinusoidal (9 G), and random (19 G squared/Hz) accelerations. System envelope is 31.6 cm long by 14 cm flange B.C. with a mass of 8.5 kg including 6.2 Kg mercury propellant.

Hyman, J., Jr.↗
NASA NTRS1981

SERT II thrusters - Still ticking after eleven years

The Space Electric Rocket Test II (SERT II) spacecraft was launched in 1970 with a primary objective of demonstrating long-term operation of a space electric thruster system. An overview is presented of all the SERT II testing conducted during the time from 1970 to 1981. Thruster testing and interaction results are considered, taking into account ion beam thrusting, distant neutralization, and the plasma beam thrust. In a discussion of durability testing, attention is given to the main cathodes, the neutralizer cathodes, the main keeper insulator, the H.V. grid insulators, the neutralizer propellant tanks, and the main propellant tanks. The most important result of the study is related to the confidence gained that mercury bombardment ion thruster systems can be built and operated in space on a routine basis with the same lifetime and performance as measured in ground testing.

Kerslake, W. R.↗
NASA NTRS2004

The Investigation Of Carbon Contamination And Sputtering Effects Of Xenon Ion Thrusters

The Electro-Physics Branch of the NASA Glenn Research Center investigates the effect of atomic oxygen, environmental durability of high performance power materials and surfaces, and low earth orbit. One of its current projects involves the analysis of ion thrusters. Ion thrusters are devices that initiate a beam of ions to a target area. The type of ion thruster that I have been working with this Summer of 2004 emits positively charged Xenon (Xe(+)) atoms through two grids, the screen grid and the accelerator grid, after it enters an ionization chamber. Insulators are used to mechanically hold and separate these two grids. A propellant isolator, an instrument that closely resembles insulators, is placed in front of the ionization chamber. Both the insulator and isolator are made with a ceramic compound and filled with insulating beads. The main difference between the two devices is that the propellant isolator allows gas to flow through, in this case, the gas is Xe(+) and the insulators do not. In order to avoid carbon deposits and other contaminating chemicals to settle on the insulators and propellant isolator, a metal shadow shield is placed around them. These shadow shields function as a protectant and can be shaped in numerous configurations. Part of my job responsibility this summer is to investigate the effectiveness of different shadow shields that are utilized on three different ion engines: the NSTAR (NASA Solar Electric Propulsion Technology Application Readiness), JIMO (Jupiter Icy Moons Orbiter), and NEXIS (Nuclear Electric Xenon Ion System). Using calculus and other mathematical tactics, I was asked to find the total flux of carbon contamination that was able to pass the protectant shadow shield. I familiarized myself with the software program, MathCad2004, to help perform some mathematical computations such as complex integration. Another method of studying the probability of contamination is by experimental simulation. After attaining the precise parameters of the actual shadow shields, I created replicas of three types of shadow shielding to be used to undergo testing. It will be placed in a machine that produces carbon atoms at a high temperature of 200 C. or beam is aimed at a targeted material. As a result of this collision, atoms and other particles are ejected out of the target surface. Another part of my internship consisted of research on sputter ejection, or the angle distribution of sputtered material. This research entailed finding the past results of sputter ejection investigation as well as creating another type of mock simulation. Other minor projects include calculating the path of Xe(+) gas through the insulating beads of the isolators and assisting my mentor in collecting data for his paper for the Joint Propulsion Conference & Exhibit to be held July 11-14,2004 in Fort Lauderdale, Florida.

Prak, Moline K.↗
NASA NTRS1980

Inert gas thrusters

Some advances in component technology for inert gas thrusters are described. The maximum electron emission of a hollow cathode with Ar was increased 60-70% by the use of an enclosed keeper configuration. Operation with Ar, but without emissive oxide, was also obtained. A 30 cm thruster operated with Ar at moderate discharge voltages give double-ion measurements consistent with a double ion correlation developed previously using 15 cm thruster data. An attempt was made to reduce discharge losses by biasing anodes positive of the discharge plasma. The reason this attempt was unsuccessful is not yet clear. The performance of a single-grid ion-optics configuration was evaluated. The ion impingement on the single grid accelerator was found to approach the value expected from the projected blockage when the sheath thickness next to the accelerator was 2-3 times the aperture diameter.

Kaufman, H. R.↗
NASA NTRS2004

Discharge Chamber Primary Electron Modeling Activities in Three-Dimensions

Designing discharge chambers for ion thrusters involves many geometric configuration decisions. Various decisions will impact discharge chamber performance with respect to propellant utilization efficiency, ion production costs, and grid lifetime. These hardware design decisions can benefit from the assistance of computational modeling. Computational modeling for discharge chambers has been limited to two-dimensional codes that leveraged symmetry for interpretation into three-dimensional analysis. This paper presents model development activities towards a three-dimensional discharge chamber simulation to aid discharge chamber design decisions. Specifically, of the many geometric configuration decisions toward attainment of a worthy discharge chamber, this paper focuses on addressing magnetic circuit considerations with a three-dimensional discharge chamber simulation as a tool. With this tool, candidate discharge chamber magnetic circuit designs can be analyzed computationally to gain insight into factors that may influence discharge chamber performance such as: primary electron loss width in magnetic cusps, cathode tip position with respect to the low magnetic field volume, definition of a low magnetic field region, and maintenance of a low magnetic field region across the grid span. Corroborating experimental data will be obtained from mockup hardware tests. Initially, simulated candidate magnetic circuit designs will resemble previous successful thruster designs. To provide opportunity to improve beyond previous performance benchmarks, off-design modifications will be simulated and experimentally tested.

Steuber, Thomas J.↗
NASA NTRS2001

Design and Performance of 40 cm Ion Optics

A 40 cm ion thruster is being developed at the NASA Glenn Research Center to obtain input power and propellant throughput capabilities of 10 kW and 550 kg. respectively. The technical approach here is a continuation of the "derating" technique used for the NSTAR ion thruster. The 40 cm ion thruster presently utilizes the NSTAR ion optics aperture geometry to take advantage of the large database of lifetime and performance data already available. Dome-shaped grids were chosen for the design of the 40 cm ion optics because this design is naturally suited for large-area ion optics. Ion extraction capabilities and electron backstreaming limits for the 40 cm ion optics were estimated by utilizing NSTAR 30 cm ion optics data. A preliminary service life assessment showed that the propellant throughput goal of 550 kg of xenon may be possible with molybdenum 40 cm ion optics. One 40 cm ion optics' set has been successfully fabricated to date. Additional ion optics' sets are presently being fabricated. Preliminary performance tests were conducted on a laboratory model 40 cm ion thruster.

Soulas, George C.↗
NASA NTRSDate not supplied

Characterizing Electrical Arc Behavior in Ion Engines

Test results using an engineering model Advanced NEXT thruster are presented to identify different arc types utilizing ground test hardware. Three distinct arc types were observed: grid-to-grid arcs, high voltage-to-chassis arcs, and a combination of both. The frequency of high voltage-to-chassis arcs ranged from 4-30%, depending on operating conditions and hardware configuration. Higher beam currents were typically associated with more frequent high voltage-to-chassis arcs. During these arcs, the measured chassis current levels ranged from 20-70 A, however a significant portion of the output beam current returned to the beam supply via the neutralizer cathode plasma, which may be a ground test facility effect. The paper also discusses conducted susceptibility tests that can be performed at the spacecraft level to address the transients associated with arcing events.

electric propulsion↗
NASA NTRS2005

Results of a 2000 Hour Wear Tof the NEXIS Ion Engine

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.

Thruster performance↗
NASA NTRS1972

Durability tests of a five centimeter diameter ion thruster system

A modified Hughes SIT-5 system is being tested for durability at the Lewis Research Center. As of October 1, 1972, the thruster subsystem has logged over 8000 hours of operation. The initial 2023 hours were run with a translating screen thrust vector grid. The thruster is currently operating with an electrostatic type vector grid. Profiles and maps taken at widely separated intervals show that performance and operating characteristics have remained essentially constant. Overall efficiency is about 32 percent and power to thrust ratio is 170 watts per millipound at a specific impulse of 2500 seconds. Telescopic examination of the vector grid shows some sputtering erosion due to charge exchange and direct impingement ions.

Nakanishi, S.↗
NASA NTRS1972

SIT-5 system development.

A 5-cm structurally integrated ion thruster (SIT-5) has been developed for attitude control and stationkeeping of synchronous satellites. With two-dimension thrust-vectoring grids, a first generation unit has demonstrated a thrust of 0.56 mlb at a beam voltage of 1200 V, total mass efficiency of 64%, and electrical efficiency of 46.8%. Structural integrity is demonstrated with a dielectric-coated grid for shock (30 G), sinusoidal (9 G), and random (19.9 G rms) accelerations. System envelope is 31.8 cm long by 13.9 cm flange bolt circle, with a mass of 8.5 kg, including 6.2 kg mercury propellant. Characteristics of a second-generation unit indicate significant performance gains.

Hyman, J., Jr.↗
NASA NTRS1976

Status of SERT 2 thrusters and spacecraft 1976

The historical record of the SERT II ion thrusters and spacecraft performance for 6 1/2 years since the February 1970 launch is reviewed. The most recent ion thruster operation test shows no changes since 1974. Thruster 2 is fully operational with no performance degradation. Thruster 1 has a high voltage grid short, but continues to demonstrate cathode and discharge relight capability. Spacecraft orbit and dynamic analysis indicates a stable, sun-synchronous spacecraft orientation by 1979. An attitude adjustment maneuver was performed in August 1976 to achieve this orientation and provide sufficient continuous solar power for thruster operation in 1979.

Kerslake, W. R.↗
NASA NTRS2005

Results of a 2000 hour wear test of the NEXIS ion engine

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.

Nuclear Electric Xenon Ion System (NEXIS)Thruster ↗
NASA NTRS1980

Analysis and design of ion thrusters for large space systems

This study undertakes the analysis and conceptual design of a 0.5 Newton electrostatic ion thruster suitable for use on large space system missions in the next decade. Either argon or xenon gas shall be used as propellant. A 50 cm diameter discharge chamber was selected to meet stipulated performance goals. The discharge plasma is contained at the boundary by a periodic structure of alternating permanent magnets generating a series of line cusps. Anode strips between the magnets collect Maxwellian electrons generated by a central cathode. Ion extraction utilizes either two or three grid optics at the user's choice. An extensive analysis was undertaken to investigate optics behavior in the high power environment of this large thruster. A plasma bridge neutralizer operating on inert gas provides charge neutralizing electrons to complete the design. The resulting conceptual thruster and the necessary power management and control requirements are described.

James, E. L.↗
NASA NTRS1988

An approach to the parametric design of ion thrusters

A methodology that can be used to determine which of several physical constraints can limit ion thruster power and thrust, under various design and operating conditions, is presented. The methodology is exercised to demonstrate typical limitations imposed by grid system span-to-gap ratio, intragrid electric field, discharge chamber power per unit beam area, screen grid lifetime, and accelerator grid lifetime constraints. Limitations on power and thrust for a thruster defined by typical discharge chamber and grid system parameters when it is operated at maximum thrust-to-power are discussed. It is pointed out that other operational objectives such as optimization of payload fraction or mission duration can be substituted for the thrust-to-power objective and that the methodology can be used as a tool for mission analysis.

Wilbur, Paul J.↗
NASA NTRS1988

An approach to the parametric design of ion thrusters

A methodology that can be used to determine which of several physical constraints can limit ion thruster power and thrust, under various design and operating conditions, is presented. The methodology is exercised to demonstrate typical limitations imposed by grid system span-to-gap ratio, intragrid electric field, discharge chamber power per unit beam area, screen grid lifetime and accelerator grid lifetime constraints. Limitations on power and thrust for a thruster defined by typical discharge chamber and grid system parameters when it is operated at maximum thrust-to-power are discussed.

Wilbur, Paul J.↗
NASA NTRS2004

Progress in NEXT Ion Optics Modeling

Results are presented from an ion optics simulation code applied to the NEXT ion thruster geometry. The error in the potential field solver of the code is characterized, and methods and requirements for reducing this error are given. Results from a study on electron backstreaming using the improved field solver are given and shown to compare much better to experimental results than previous studies. Results are also presented on a study of the beamlet behavior in the outer radial apertures of the NEXT thruster. The low beamlet currents in this region allow over-focusing of the beam, causing direct impingement of ions on the accelerator grid aperture wall. Different possibilities for reducing this direct impingement are analyzed, with the conclusion that, of the methods studied, decreasing the screen grid aperture diameter eliminates direct impingement most effectively.

Emhoff, Jerold W.↗