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At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Electrostatic thrusters.

The current status of research and development programs on electrostatic thrusters is reviewed. Current programs that utilize mercury electron-bombardment thrusters range from 5- to 30-cm in diameter. Recent progress on the 5-cm thruster has emphasized durability, with accelerator time exceeding 6300 hours and total time on the rest of the thruster exceeding 8300 hours. Recent progress on the 30-cm thruster has been outstanding in dished-grid accelerator systems. Ion beams up to 5 amperes have been obtained for short periods with 1000 volts net accelerating potential difference. The cesium electron-bombardment and cesium contact programs are also described.

Kaufman, H. R.↗
NASA NTRS2000

Mission Advantages of Constant Power, Variable Isp Electrostatic Thrusters

Electric propulsion has moved from station-keeping capability for spacecraft to primary propulsion with the advent of both the Deep Space One asteroid flyby and geosynchronous spacecraft orbit insertion. In both cases notably more payload was delivered than would have been possible with chemical propulsion. To provide even greater improvements electrostatic thruster performance could be varied in specific impulse, but kept at constant power to provide better payload or trip time performance for different mission phases. Such variable specific impulse mission applications include geosynchronous and low earth orbit spacecraft stationkeeping and orbit insertion, geosynchronous reusable tug missions, and interplanetary probes. The application of variable specific impulse devices is shown to add from 5 to 15% payload for these missions. The challenges to building such devices include variable voltage power supplies and extending fuel throughput capabilities across the specific impulse range.

Oleson, Steven R.↗
NASA NTRS1984

Ion beamlet steering for two-grid electrostatic thrusters

An experimental study of ion beamlet steering in which the direction of beamlets emitted from a two grid aperture system is controlled by relative translation of the grids, is described. The results can be used to design electrostatic accelerating devices for which the direction and focus of emerging beamlets are important. Deflection and divergence angle data are presented for two grid systems as a function of the relative lateral displacement of the holes in these grids. At large displacements, accelerator grid impingements become excessive and this determines the maximum allowable displacement and as a result the useful range of beamlet deflection. Beamlet deflection is shown to vary linearly with grid offset angle over this range. The divergence of the beamlets is found to be unaffected by deflection over the useful range of beamlet deflection. The grids of a typical dished grid ion thruster are examined to determine the effects of thermally induced grid distortion and prescribed offsets of grid hole centerlines on the characteristics of the emerging beamlets. The results are used to determine the region on the grid surface where ion beamlet deflections exceed the useful range. Over this region high accelerator grid impingement currents and rapid grid erosion are predicted.

Homa, J. M.↗
NASA NTRS2006

Miniature Free-Space Electrostatic Ion Thrusters

A miniature electrostatic ion thruster is proposed for maneuvering small spacecraft. In a thruster based on this concept, one or more propellant gases would be introduced into an ionizer based on the same principles as those of the device described in an earlier article, "Miniature Bipolar Electrostatic Ion Thruster". On the front side, positive ions leaving an ionizer element would be accelerated to high momentum by an electric field between the ionizer and an accelerator grid around the periphery of the concave laminate structure. On the front side, electrons leaving an ionizer element 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 a thruster design consisting of multiple membrane ionizers in a thin laminate structure with a peripheral accelerator grid, the direction of thrust could then be controlled (without need for moving parts in the thruster) by regulating the supply of gas to specific ionizer.

Hartley, Frank T.↗
NASA NTRS2006

Miniature Electrostatic Ion Thruster With Magnet

A miniature electrostatic ion thruster is proposed that, with one exception, would be based on the same principles as those of the device described in the previous article, "Miniature Bipolar Electrostatic Ion Thruster". The exceptional feature of this thruster would be that, in addition to using electric fields for linear acceleration of ions and electrons, it would use a magnetic field to rotationally accelerate slow electrons into the ion stream to neutralize the ions.

Hartley, Frank T.↗
NASA NTRS2019

Characterization of Low-Temperature, Co-Fired Ceramic-Manufactured Electrostatic Thruster-Closeout Report

The goal of the project was to evaluate prototypes of an experimental thruster developed by the University of Arkansas (UA), Fayetteville, AR. The design under evaluation is a radio frequency (RF) electrostatic thruster that was fabricated using the low-temperature, co-fired ceramic (LTCC) materials and fabrication process. This materials system is analogous to printed circuit board (PCB) technology with the most significant difference being that the laminate is replaced by a ceramic material and the copper layer is replaced by printed sinterable silver paste. LTCC designs are baked after fabrication and assembled to realize an entirely monolithic structure with internal conductors, vias, and cavities. In this process, the LTCC electrostatic thruster (LTCC-ET) that is the subject of the present work becomes a monolithic ceramic thruster capable of withstanding temperatures in excess of 500 ºC. The UA and NASA Marshall Space Flight Center (MSFC) jointly performed prototype testing on the LTCC-ET under a NASA Cooperative Agreement Notice (CAN) award. The LTCC-ET was tested at MSFC in May 2018 over a 1-week period. There were two goals for the test program: (1) Testing to determine the operating parameters required to create plasma ignition in the test articles. This was explored by setting a propellant flowrate and increasing RF power until plasma ignition was observed. Testing was conducted with both argon and krypton. (2) Investigate the thrust and specific impulse (Isp) performance of the thruster as a function of propellant flowrate and grid voltage. This goal was not met during the project as technical challenges in maintaining stable plasma ignition arose due to stress and heating of the RF power feed. In summary, a prototype thruster design (consisting of three packaged units) was fabricated by UA and tested for the first time under vacuum conditions at MSFC to experimentally determine basic performance metrics and functionality. It was found that the design was not sufficiently optimized or robust enough in its initial iteration to support a significant test campaign or characterization program. It was concluded that the propellant outlet channels must be reduced in size with the flowpaths adjusted to increase propellant residence time in the thruster, and that the RF connector must be replaced with a version capable of handling higher power throughput and heating. However, even in its unoptimized form, a plasma could be produced in the LTCCET, demonstrating the efficacy of the design approach. The design is especially compelling due to its low cost to manufacture and, more importantly, its scalability of size and power throughput. Low cost and scalability are also important in that additional functionalities, such as thrust vectoring and plume charge neutralization, can be integrated into future designs with minimal additional cost. This project has matured the LTCC-ET development Technology Readiness Level (TRL) from 1 to 2. The low-cost RF plasma source portion of the LTCC device was matured from TRL 2 to 4 through the demonstration of RF plasma ignition under vacuum conditions.

Huang, P.-H. A.↗
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 NTRS2013

Development And Testing Of The Inertial Electrostatic Confinement Diffusion Thruster

The Inertial Electrostatic Confinement (IEC) diffusion thruster is an experiment in active development that takes advantage of physical phenomenon that occurs during operation of an IEC device. The IEC device has been proposed as a fusion reactor design that relies on traditional electrostatic ion acceleration and is typically arranged in a spherical geometry. The design incorporates two radially-symmetric spherical electrodes. Often the inner electrode utilizes a grid of wire shaped in a sphere with a radius 15 to 50 percent of the radius of the outer electrode. The inner electrode traditionally has 90 percent or more transparency to allow particles (ions) to pass to the center of the spheres and collide/recombine in the dense plasma core at r=0. When operating the IEC, an unsteady plasma leak is typically observed passing out one of the gaps in the lattice grid of the inner electrode. The IED diffusion thruster is based upon the idea that this plasma leak can be used for propulsive purposes. The IEC diffusion thruster utilizes the radial symmetry found in the IEC device. A cylindrical configuration is employed here as it will produce a dense core of plasma the length of the cylindrical grid while promoting the plasma leak to exhaust through an electromagnetic nozzle at one end of the apparatus. A proof-of-concept IEC diffusion thruster is operational and under testing using argon as propellant (Figure 1).

Becnel, Mark D.↗
NASA NTRS1981

Electric propulsion and its applications to space missions

Consideration is given the NASA, Air Force and European electric propulsion programs, the characteristics of primary electric propulsion systems, nuclear electric orbit transfer vehicles, and such topics in the fundamental processes of electrostatic thrusters as sputtering in mercury ion thrusters, the screen hole plasma sheath of an ion accelerator system, and the modelling of ion beam neutralization and nitrogen chemisorption. Also considered are electrostatic thruster components and systems, electromagnetic thrusters such as MPD and RIT systems, electric rail guns and mass drivers, power sources which include solar and nuclear alternatives, power conversion systems and their cooling apparatus, and the environmental interactions between spacecraft and their electric propulsion systems.

Finke, R. C.↗
NASA NTRS1992

Electrostatic ion thruster optics calculations

Calculations have been performed which encompass both a self-consistent ion source extraction plasma sheath and the primary ion optics including sheath and electrode-induced aberrations. Particular attention is given to the effects of beam space charge, accelerator geometry, and properties of the downstream plasma sheath on the position of the electrostatic potential saddle point near the extractor electrode. The electron blocking potential blocking is described as a function of electrode thickness and secondary plasma processes.

Whealton, John H.↗
NASA NTRS2012

Development and Testing of a Green-Propellant Micro-Hybrid Thruster with Electrostatic Ignition

As early as 1937 German scientists at Peenemunde experimented with highly unstable fuel blends of nitrous oxide (N2O) and ethanol. These early tests mostly resulted in explosions and destroyed rocket engines. More recently several companies have developed experimental nitrous oxide fuel blends (NOFB) with Isp exceeding 300 sec. Although NOFBx has recently been cleared for tests on the International Space Station, this propellant remains highly experimental and has not been cleared for commercial transport by the US DOT. Recent work by Karabeyoglu et al. has raised concerns about the safety risks of mixing hydrocarbons with N2O. Liquid oxidizer/fuel blends are highly explosive and require extreme care in transport and servicing. By adding small amounts of a liquid organic fuel such as alcohol or a hydrocarbon, the odds of an explosive decomposition event are significantly increased.iv The proposed solution mitigates the explosion hazards of NOFB by separating the oxidizer from the hydrocarbon fuel formed as of a small cylindrical section of ABS thermoplastic. As N2O vapor flows across the grain segment, current enters a 1000 VDC high-tension lead in the ABS fuel grain and produces an inductive spark that vaporizes a small amount of the material. The ablated fuel vapor plus residual energy from the spark seed a localized exothermic N2O dissociation that produces sufficient heat to initiate combustion. The process is also effective when gaseous oxygen is used. A low TRL (2-3) prototype demonstrating the feasibility of controlled hydrocarbon-seeding was recently tested at Utah State University.v The unit features a miniature 2.5 cm ABS fuel grain fabricated using a Stratasys Dimension 3-D printer. The 9-N thruster was pulse-fired up to 27 consecutive times on a single ABS grain segment. Ignition was achieved by as little as 12-15 Joules energy input. This value is contrasted with the typical 30-minute pre-heat requirement for the ECAPS LMP-103S ADN-based monopropellant, requiring an energy input of 14,850 Joules for catalytic dissociation. The hydrocarbon-seeded micro-hybrid was also adapted as a non-pyrotechnic ignitor for a 900 N (200-lbf) thrust hybrid motor. The motor was successfully ignited 4 consecutive times with no hardware swaps or propellant additions. The amount of ABS seed material that can be fit into the injector cap is the only limit to the number of available repeat firings. This series of tests marks the first time a hybrid motor was ever ignited by other than a solid-propellant pyrotechnic charge or bi-propellant flame ignitor. Nitrous oxide hybrid motors are typically difficult to ignite and usually require multiple solid-propellant charges to initiate combustion, so this nonpyrotechnic ignition is a significant accomplishment. The controlled hydrocarbon-seeding approach is fundamentally different from all other green propellant solutions offered by the aerospace industry. Although the proposed system is more correctly a hybrid technology; the system retains all the simple features of a monopropellant design. To date no optimization study has been performed to identify the best grain geometry for electrostatic ignition. Fortunately, because the grain segments are fabricated using rapid-prototyping technology, changing the grain geometry is as simple as modifying the 3-D printer CAD-file. Vacuum Isp exceeding 270 seconds has been demonstrated (Ref v), a value significantly higher than those offered by competing green monopropellant options. The propellants of choice, N2O/GOX and ABS are 100% non-toxic, non-explosive, and environmentally benign. Because the inert oxidizer and fuel components are mixed only within the combustion chamber, the system retains the inherent safety of a hybrid rocket and can be piggy-backed as a secondary payload with no overall mission risk increase to the primary payload, an excellent characteristic for secondary launch systems.

Whitmore, Stephen A.↗
NASA NTRS1978

Advanced electrostatic ion thruster for space propulsion

The suitability of the baseline 30 cm thruster for future space missions was examined. Preliminary design concepts for several advanced thrusters were developed to assess the potential practical difficulties of a new design. Useful methodologies were produced for assessing both planetary and earth orbit missions. Payload performance as a function of propulsion system technology level and cost sensitivity to propulsion system technology level are among the topics assessed. A 50 cm diameter thruster designed to operate with a beam voltage of about 2400 V is suggested to satisfy most of the requirements of future space missions.

Masek, T. D.↗
NASA NTRS1990

Plasma particle simulation of electrostatic ion thrusters

Charge exchange collisons between beam ions and neutral propellant gas can result in erosion of the accelerator grid surfaces of an ion engine. A particle in cell (PIC) is developed along with a Monte Carlo method to simulate the ion dynamics and charge exchange processes in the grid region of an ion thruster. The simulation is two-dimensional axisymmetric and uses three velocity components (2d3v) to investigate the influence of charge exchange collisions on the ion sputtering of the accelerator grid surfaces. An example calculation has been performed for an ion thruster operated on xenon propellant. The simulation shows that the greatest sputtering occurs on the downstream surface of the grid, but some sputtering can also occur on the upstream surface as well as on the interior of the grid aperture.

Peng, Xiaohang↗