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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 Oct. 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 per cent 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. An independent test of the propellant storage and cathode-isolator-vaporizer subsystem has demonstrated good reliability under simulated thruster operating conditions.

Nakanishi, S.↗

NASA NTRS2018

Computational Investigation of the Near-Field Plasma Plume in Ion-Ion Propulsion

A two-fluid computational model of plasma flows was developed to investigate the plume of an ion-ion propulsion system. The densities of positive and negative ions, along with the associated values of net charge, electric field, and electric potential were calculated throughout the domain. The computational domain was chosen to be large enough (25 thruster diameters downstream of the accelerating grids) to examine the neutralization of the plume. The resulting plasma electric potential and charge neutrality at the downstream end of the domain are shown. The results from this simulation are compared to existing literature on ion-ion plasma thrusters.

Sankaran, K.↗

NASA NTRS1979

Sputtering in mercury ion thrusters

A model, which assumes that chemisorption is the dominant mechanism, is applied to the sputtering rate measurements of the screen grid of a 30 cm thruster in the presence of nitrogen. The model utilizes inputs from a variety of experimental and analytical sources. The model of environmental effects on sputtering was applied to thruster conditions of low discharge voltage and a discussion of the comparison of theory and experiment is presented.

Mantenieks, M. A.↗

NASA NTRS2000

Performance of Titanium Optics on a NASA 30 Cm Ion Thruster

The results of performance tests with two titanium optics sets are presented and compared to those of molybdenum optics. All tests were conducted on a 30 cm ion thruster that was nearly identical to the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) thruster design. Optics performance tests were conducted over a thruster input power range of 0.5 to 4.6 kW. Optics performance including impingement-limited total voltages, electron backstreaming limits, screen grid ion transparencies, near-field beam current density profiles, beam divergence angles, and beam divergence thrust correction factors were determined throughout this power range. The impingement-limited total voltages for titanium optics were within 10 to 55 V of those for molybdenum optics. Electron backstreaming limit magnitude as a function of peak beam current density for both molybdenum and titanium optics were within a few volts of each other, indicating similar hot grid gaps for these two grid materials during steady-state operation. Beam divergence half-angles at 90 percent of the total beam current and thrust correction factors for both titanium optics sets were within one degree and one percent, respectively, of those for molybdenum optics. When thruster power was increased to 2.3 kW immediately following discharge ignition, the titanium screen grid came into contact with the accelerator grid within five minutes of ignition. Relative to molybdenum, titanium's larger thermal expansion and smaller thermal conductivity likely caused the screen grid to thermally expand more relative to the accelerator grid during startup.

Soulas, George C.↗

NASA NTRS2002

Performance Evaluation of 40 cm Ion Optics for the NEXT Ion Engine

The results of performance tests with two 40 cm ion optics sets are presented and compared to those of 30 cm ion optics with similar aperture geometries. The 40 cm ion optics utilized both NSTAR and TAG (Thick-Accelerator-Grid) aperture geometries. All 40 cm ion optics tests were conducted on a NEXT (NASA's Evolutionary Xenon Thruster) laboratory model ion engine. Ion optics performance tests were conducted over a beam current range of 1.20 to 3.52 A and an engine input power range of 1.1 to 6.9 kW. Measured ion optics' performance parameters included near-field radial beam current density profiles, impingement-limited total voltages, electron backstreaming limits, screen grid ion transparencies, beam divergence angles, and start-up transients. Impingement-limited total voltages for 40 cm ion optics with the NSTAR aperture geometry were 60 to 90 V lower than those with the TAG aperture geometry. This difference was speculated to be due to an incomplete burn-in of the TAG ion optics. Electron backstreaming limits for the 40 cm ion optics with the TAG aperture geometry were 8 to 19 V higher than those with the NSTAR aperture geometry due to the thicker accelerator grid of the TAG geometry. Because the NEXT ion engine provided beam flatness parameters that were 40 to 63 percent higher than those of the NSTAR ion engine, the 40 cm ion optics outperformed the 30 cm ion optics.

Soulas, George C.↗

NASA NTRS2006

Miniature Bipolar Electrostatic Ion Thruster

The figure presents a concept of a bipolar miniature electrostatic ion thruster for maneuvering a small spacecraft. The ionization device in the proposed thruster would be a 0.1-micron-thick dielectric membrane with metal electrodes on both sides. Small conical holes would be micromachined through the membrane and electrodes. An electric potential of the order of a volt applied between the membrane electrodes would give rise to an electric field of the order of several mega-volts per meter in the submicron gap between the electrodes. An electric field of this magnitude would be sufficient to ionize all the molecules that enter the holes. In a thruster-based on this concept, one or more propellant gases would be introduced into such a membrane ionizer. Unlike in larger prior ion thrusters, all of the propellant molecules would be ionized. This thruster would be capable of bipolar operation. There would be two accelerator grids - one located forward and one located aft of the membrane ionizer. In one mode of operation, which one could denote the forward mode, positive ions leaving the ionizer on the backside would be accelerated to high momentum by an electric field between the ionizer and an accelerator grid. Electrons leaving the ionizer on the front side would be ejected into free space by a smaller accelerating field. The equality of the ion and electron currents would eliminate the need for an additional electron- or ion-emitting device to keep the spacecraft charge-neutral. In another mode of operation, which could denote the reverse mode, the polarities of the voltages applied to the accelerator grids and to the electrodes of the membrane ionizer would be the reverse of those of the forward mode. The reversal of electric fields would cause the ion and electrons to be ejected in the reverse of their forward mode directions, thereby giving rise to thrust in the direction opposite that of the forward mode.

Hartley, Frank T.↗

NASA NTRS1972

Change in transmittance of fused silica as a means of detecting material sputtered from components on a 5-cm ion thruster

Two endurance tests of a 5-cm mercury bombardment thruster are reported. Both tests used a translational screen-grid system with the beam vectored 10 degrees. The first test lasted 141 hours and the second test operated for 2026 hours. In each test two fused silica samples (solar cell covers), 2.0 cm by 2.1 cm, were placed in shielded holders to detect materials sputtered from the thruster. Spectral optical properties between 0.398 and 2.16 microns were measured on each sample, both before and after the endurance tests. The deposition on each sample was spectrographically analyzed to determine the type of materials sputtered from the thruster. It was found that sputtering from the neutralizer is highly dependent on its position with respect to the beam edge. The sputtering from the accelerator grid of the translational screen-grid system of the 2026 hour test was sufficient to form an opaque film on the sample located in the direction opposite to the vectored beam.

Weigand, A. J.↗

NASA NTRS1993

Experimental and analytical ion thruster research

The results of further spectroscopic studies on the plume from a 3 cm ion source operated on an argon propellant is reported on. In particular, it is shown that it should be possible to use the spectroscopic technique to measure the plasma density of the ion plume close to the grids, where it is difficult to use electrical probe measurements. How the technique, along with electrical probe measurements in the far downstream region of the plume, can be used to characterize the operation of a three-grid, 15 cm diameter thruster from NASA JPL is outlined. Pumping speed measurements on the Vacuum Research Facility have shown that this facility should be adequate for testing the JPL thruster at pressures in the low 10(exp -5) Torr range. Finally, we describe a simple analytical model which can be used to calculate the grid impingement current which results from charge-exchange collisions in the ion plume.

Ruyten, Wilhelmus M.↗

NASA NTRS2004

An Overview of the Nuclear Electric Xenon Ion System (NEXIS) Activity

The Nuclear Electric Xenon Ion System (NEXIS) research and development activity within NASA's Project Prometheus, was one of three proposals selected by NASA to develop thruster technologies for long life, high power, high specific impulse nuclear electric propulsion systems that would enable more robust and ambitious science exploration missions to the outer solar system. NEXIS technology represents a dramatic improvement in the state-of-the-art for ion propulsion and is designed to achieve propellant throughput capabilities >= 2000 kg and efficiencies >= 78% while increasing the thruster power to >= 20 kW and specific impulse to >= 6000 s. The NEXIS technology uses erosion resistant carbon-carbon grids, a graphite keeper, a new reservoir hollow cathode, a 65-cm diameter chamber masked to produce a 57-cm diameter ion beam, and a shared neutralizer architecture to achieve these goals. The accomplishments of the NEXIS activity so far include performance testing of a laboratory model thruster, successful completion of a proof of concept reservoir cathode 2000 hour wear test, structural and thermal analysis of a completed development model thruster design, fabrication of most of the development model piece parts, and the nearly complete vacuum facility modifications to allow long duration wear testing of high power ion thrusters.

ion thrusters↗

NASA NTRS2006

Comparison of NASA's 30-cm Ion Thruster Capabilities with the Dawn Mission Requirements

The Dawn ion thrusters meet all of the performance and life requirements for the Dawn mission. FT001 has two magnets in sideways in the middle magnet ring, but this has only a minor impact on the thruster performance. Electron-backstreaming is the first wear-out failure mechanism for the Dawn ion thrusters - PFA Modeling Indicates: a) The mission cannot be completed with a single thruster; b) The probability for mission success based on accelerator grid wear is > 99% for the worst case thruster usage for nominal thruster beamlet profiles; and c) FT002 has the most peaked beam profile which impacts the thruster s throughput capability. PFA results indicate that this reduces the probability for mission success based on accelerator grid wear to 97% based on the worst case thruster usage. The wear-out probability is zero if all three thrusters are used during the mission.

ion thrusters↗

NASA NTRS1993

Operating Characteristics of a 15-cm dia. Ion Engine for Small Planetary Spacecraft

A 15-cm diameter, scaled-down version of the NASA light-weight 30-cm ion engine has been developed for potential application to very small planetary spacecraft. Integration of the 15-cm ion source into 4x15-cm segmented engine configuration results in a 30-cm equivalent engine which can be throttled over a 7-to-1 input power variation with a constant beam current in each of the four segments. Throttling the segmented engine by turning off individual segments can result in a significant decrease in the required service life (and qualification requirements) of the ion source components.

grids segmented ion engine↗

NASA NTRS1994

Segmented Ion Engine Operation and Performance

The continuing trend toward smaller and smaller planetary spacecraft to enable the use of smaller, less expensive launch vehicles has motivated an examination of new approaches to reduce the size and mass of xenon ion propulsion systems for these new spacecraft. A system is proposed which is based on the use of a single 4x15-cm segmented ion thruster operated from a single internally redundant power processing unit. An engine input power throttling range of 380 to 4640 W is projected. The design and operation over a power range of 600 to 2400 W of a laboratory model segmented ion engine equipped with carbon-carbon grids is presented. There appear to be no problems associated with interactions with multiple ion sources operating from a single set of high voltage power supplies, neutralization of multiple ion sources from a single centrally located neutralizer, or operation of flat, thin carbon-carbon grids.

ion↗

NASA NTRS1981

Inert-gas thruster technology

Attention is given to recent advances in component technology for inert-gas thrusters. It is noted that the maximum electron emission of a hollow cathode with Ar can be increased 60-70% by using an enclosed keeper configuration. Operation with Ar but without emissive oxide has also been attained. A 30-cm thruster operated with Ar at moderate discharge voltages is found to give double-ion measurements consistent with a double-ion correlation developed earlier on the basis of 15-cm thruster data. An attempt is made to reduce discharge losses by biasing anodes positive of the discharge plasma. The performance of a single-grid ion-optics configuration is assessed. The ion impingement on the single-grid accelerator is found to approach the value expected from the projected blockage when the sheath thickness next to the accelerator is 2-3 times the aperture diameter.

Kaufman, H. R.↗

NASA NTRS1975

The effects of exposure to LN2 temperatures and 2.5 suns solar radiation on 30-cm ion thruster performance

An experimental test program was developed to demonstrate all 30 cm Hg-ion bombardment thruster functions over the thermal environment of several proposed missions. A 30 cm thruster with grids dished 1.25 cm and instrumented with 31 thermocouples, was placed in a vacuum tank equipped with minus 196 C walls. Cold storage of a thruster was simulated and temperatures as low as minus 100 C were attained on the thruster. The thruster started successfully from these cold conditions. The thruster operating at both half and full beam power was exposed to 2.5 suns on axis solar simulation. Various thruster thermal configurations, used to simulate multiple thruster operation, were tested at the above conditions. The results of these tests are reported herein.

Mirtich, M. J.↗

NASA NTRS1975

The effects of exposure to LN2 temperatures and 2.5 suns solar radiation on 30-cm ion thruster performance

An experimental test program was developed to demonstrate all 30 cm Hg-ion bombardment thruster functions over the thermal environment of several proposed missions. A 30 cm thruster with grids dished 1.25 cm and instrumented with 31 thermocouples, was placed in a vacuum tank equipped with -196 C walls. Cold storage of a thruster was simulated and temperatures as low as -100 C were attained on the thruster. The thruster started successfully from these cold conditions. The thruster operating at both half and full beam power was exposed to 2.5 suns on axis solar simulation. Various thruster thermal configurations, used to simulate multiple thruster operation, were tested at the above conditions. The results of these tests are reported herein.

Mirtich, M. J.↗

NASA NTRS2001

Inter-cusp Ion and Electron Transport in a Nstar-derivative Ion Thruster

Diffusion of electrons and ions to anode surfaces between the magnetic cusps of a NASA Solar Electric Propulsion Technology Application Readiness ion thruster has been characterized. Ion flux measurements were made at the anode and at the screen grid electrode. The measurements indicated that the average ion current density at the anode and at the screen grid were approximately equal. Additionally, it was found that the electron flux to the anode between cusps is best described by the classical cross-field diffusion coefficient.

Foster, John E.↗

NASA NTRS2001

Ion and Electron Transport in an Nstar-derivative Ion Thruster

Foster, John E.↗

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↗