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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.
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The ST4/Champollion mission is designed to rendezvous with and land on the comet Tempel 1 and return data from the first-ever sampling of a comet surface.
One of the first science missions to use solar electric propulsion (SEP) may be the Comet Nucleus Sample Return (CNSR) mission which is designed to return samples of volatiles and dust from the nucleus of a comet. The use of SEP for CNSR makes the mission affordable and scientifically more attractive because it enables the use of a smaller, lower cost launch vehicle and it significantly reduces the time required to return samples to the Earth.
This paper describes the development and integration of a one millipound thrust ion thruster system on the NASA Goddard ATS-F spacecraft with the objective of outlining for the propulsion system engineer: (1) system interactions or potential problems associated with the application of ion thruster systems, and (2) progress which has been made and is being made in identifying and solving these problems. Principal areas of concern are the compatibility of the thruster exhaust (ions and atoms) with sensitive spacecraft surfaces and possible electromagnetic interactions of the thruster and its exhaust with other spacecraft subsystems.
Electric propulsion systems for rocket vehicles, particularly dynamics and flight mechanics of ion motor
In this paper we present ion thruster design concepts created using the new computer codes that model performance limiting and erosion mechanisms. Presently, the codes model extraction grid ion optics and both discharge and neutralizer hollow cathodes.
New design approach for high-power (100-kW class or greater) ion engines conceptually divides single engine into combination of smaller discharge chambers integrated to operate as single large engine. Analogous to multicylinder automobile engine, benefits include reduction in required accelerator system span-to-gap ratio for large-area engines, reduction in required hollow-cathode emission current, mitigation of plasma-uniformity problem, increased tolerance to accelerator system faults, and reduction in vacuum-system pumping speed.
Outer-planet mission analysis was performed using three next generation solar-electric ion thruster models. Optimal trajectories are presented that maximize the delivered mass to the designated outer planet. Trajectories to Saturn and Neptune with a single Venus gravity assist are investigated. For each thruster model, the delivered mass versus flight time curve was generated to obtain thruster model performance. The effects of power to the thrusters and resonance ratio of Venutian orbital periods to spacecraft period were also studied. Multiple locally optimal trajectories to Saturn and Neptune have been discovered in different regions of the parameter search space. The characteristics of each trajectory are noted.
Spacecraft thruster systems are disclosed. In some instances, a spacecraft thruster system may include stacked ion thrusters and/or ion thruster layers. The ion thrusters and/or ion thruster layers may be sequentially activated and jettisoned from the thruster system after use.
This paper describes the different electro static and electro magnetic emissions of the ion engine for each of the the thrust levels the engine has operated in space and in the test chamber.
Upon the completion of two long-duration life tests of a 30-cm ion engine, the orifice channel of the neutralizer hollow cathode was eroded away to as much as twice its original diameter. Whereas the neutralizer cathode orifice opened significantly, no noticeable erosion of the discharge cathode orifice was observed. Noquantitative explanation of these erosion trends has been established since the completion of the two life tests. A two-dimensional model of the partially ionized gas inside these devices has been developed and applied to the neutralizer hollow cathode. The numerical simulations show that the main mechanism responsible for the channel erosion is sputtering by Xe+. These ions are accelerated by the sheath along the channel and bombard the surface with kinetic energy/charge of about 17 V at the beginning of cathode life. The density of the ions inside the neutralizer orifice is computed to be as high as 2.1 x 10(sup 22) m(sup -3). Because of the 3.5-times larger diameter of the discharge cathode orifice, the ion density inside the orifice is more than 40 times lower and the sheath drop 7 V lower compared with the values in the neutralizer. At these conditions, Xe+ can cause no significant sputtering of the surface.
Impulse scaling during magnetic reconnection, the magnetic energy conversion to kinetic energy, via direct Mach probe measurements in Magnetic Reconnection Experiment is examined. Ion exhaust velocity and impulse scalings with reconnecting magnetic field during the push phase of driven reconnection are presented. The outflows and impulse measurements are compared with global MHD simulations. Both measurements and simulations reveal a favorable scaling, greater than linear, of impulse with reconnecting field. These scaling results establish that magnetic reconnection could be utilized for plasma propulsion.
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Thermionic reactor electric propulsion for unmanned outer planets exploration, discussing spacecraft design, launch vehicle, weight factors, etc
A nuclear thermionic reactor as the electric power source for an electric propulsion spacecraft introduces a nuclear radiation environment that affects the spacecraft configuration, the use and location of electrical insulators and the science experiments. The spacecraft is conceptually configured to minimize the nuclear shield weight by: (1) a large length to diameter spacecraft; (2) eliminating piping penetrations through the shield; and (3) using the mercury propellant as gamma shield. Since the alumina material is damaged by the high nuclear radiation environment in the reactor it is desirable to locate the alumina insulator outside the reflector or develop a more radiation resistant insulator.
The unmanned planetary exploration to be conducted in the last two decades of this century includes many higher energy missions which tax all presently available propulsion systems beyond their limit. One candidate with the versatility and performance to meet these mission objectives is nuclear electric propulsion (NEP). Additionally, the NEP System is feasible in orbit raising operations with the Shuttle or Shuttle/Tug combination. A representative planetary mission is described (Uranus-Neptune flyby with probe), and geocentric performance and tradeoffs are discussed. The NEP System is described in more detail with particular emphasis on the power subsystem consisting of the thermionic reactor, heat rejection subsystem, and neutron shield.
An improvement in thruster performance was achieved by reducing the diameter of the accelerator grid holes. The smaller accelerator grid holes resulted in a reduction in neutral mercury atoms escaping the discharge chamber, which in turn enhanced the discharge propellant utilization from approximately 68 percent to 92 percent. The accelerator grids were fabricated by ion machining with an 8-centimeter-diameter thruster, and the screen grid holes individually focused ion beamlets onto the blank accelerator grid. The resulting accelerator grid holes are less than 1.12 millimeters in diameter, while previously used accelerator grids had hole diameters of 1.69 millimeters. The thruster could be operated with the small-hole accelerator grid at neutralizer potential.
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