Engineering topics
Publications and source records attributed to Brophy, J. R..
This paper describes the an on-going study to examine the feasibility of using SEP based on derivatives of the Deep Space 1 ion propulsion system (IPS) technology to enable a Mars Sample Return (MSR mission to be performed from a single medium-class launch vehicle.
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.
Current interest in advanced propulsion within NASA and research activities in advanced propulsion concepts at the Jet Propulsion Laboratory are reviewed. The concepts, which include high power plasma thrusters such as lithuim-fueled Lorentz-Force-Accelerators, MEMS-scale propulsion systems, in-situ propellant utilization techniques, fusion propulsion systems and methods of using antimatter, offer the potential for either significantly enhancing space transportation capability as compared with that of traditional chemical propulsion, or enabling ambitious new missions.
Research activities in advanced propulsion concepts at the Jet Propulsion Laboratory are reviewed. The concepts were selected for study because each offers the potential for either significantly enhancing space transportation capability or enabling bold, ambitious new missions.
This paper presents an evaluation of the use of solar sails based on inflatable-structures technologies to perform low-cost robotic planetary missions.
Two sets of 30-cm dia. dished carbon-carbon ion engine grid sets were fabricated, assembled and tested. The carbon-carbon grid blanks were fabricated from unidirectional prepreg tape, having a very high fiber volume fraction of 57 percent and resulting in the strongest quasi-isotropic carbon-carbon material known to the authors. The flexural modulus of the panel material was measured at 275 MPa, corresponding to 80 percent of the value of molybdenum.
Performance and endurance testing of a 1.35 kW thruster with anode layer (TAL), developed at the Central Research Institute of Machine Building (TsNIMASH) is described.
Solar electric xenon ion propulsion can be used to deliver a substantial quantity of science instruments in rendezvous missions to small bodies such as comets and main belt asteroids with an Atlas IIAS launch vehicle. The performance of the ion propulsion system enables it to deliver typically more than twice the total mass to the destination in less than half the trip time relative to a chemical/ballistic approach using the same launch vehicle.
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.
To determine the failure mechanism and erosion characteristics of an ion accelerator system due to erosion by charge-exchange ions a test was performed in which a 30-cm-diameter, 2-grid ion accelerator system was tested to failure. The erosion charcteristics observed in this test, however, imply significantly shorter accelerator grid life times than typically stated in the literature. Finally, the test suggests that structural failure is probably not the most likely first failure mechanism for the accelerator grid.
The accelerator grid in two-grid ion optics systems is subject to sputter erosion by charge-exchange ions created in the ion beam.
This paper describes an optical technique developed for measuring small differential grid displacements due to thermal expansion of an ion thruster accelerator system. The technique is based on confocal scanning optical microscope type II. For the measurements of small displacements where there are distances on the order of a meter or more between the lens plane and the sample, some of the optical components are moved while the sample is kept fixed. The feasibility of applying this technique to measure the thermally induced ion thruster grid displacements was demonstrated in a bench-top simulation. It is noted that this technique can also provide information on grid movement resulting from thermal transients such as the start-up.
A lightweight spacecraft design concept has been developed which employs solar-electric ion propulsion to conduct a lunar science mission. The design relies on off-the-shelf technology to achieve high reliability and near-term launch readiness status. Overall project costs are kept to a minimum through the concentration of mission objectives, as well as through the minimization of payload launch costs by resort to the Space Shuttle's Get-Away Special canister. Attention is given to the spacecraft's instrumentation configuration.
The design and operation of a subscale MPD device is described as part of a program to develop a multimegawatt engine. The device is used as a test bed for component development, primarily the cathode and anode. The first series of exploratory tests have been conducted and the engine has been operated for 8.5 hours in the steady state, over 74 runs, and at a power level as high as 72 kW. A diffuser based vacuum tank pumping scheme has been evaluated with the engine operating at 17 kW and shows a positive result.
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Design, development and preliminary testing of a two-engine xenon ion propulsion module for early flight evaluation is described. Extensive use is made of flight spare propellant system components and also engineering model 30-cm ion engines and assemblies originally developed for the Solar Electric Propulsion System program. Significant design features include a redundant propellant feed system that incorporates a novel gas pulse assembly for rapid and completely reliable engine startup, a central neutralizer subsystem with dual neutralizers for redundancy, and major ion engine performance improvements resulting in a nearly doubling of the 30-cm engine thrust in addition to operating on xenon rather than mercury propellant. At a module input power of 10.0 kw, maximum thrust and specific impulse are projected to be 0.4 N and 3,500 sec. respectively, for a total module efficiency of 67 percent. Total mass of the xenon ion module, including the propellant tank, is only 70.2 kg. The technical approaches taken towards developing and integrating the subsystems comprising this ion propulsion module are presented and discussed in detail.
Steady-state testing of advanced, multimegawatt MPD engines requires vacuum facilities with pumping speeds which are not presently available. Gas dynamic diffusers have been proposed as a possible solution to this problem. An analytical investigation into the feasibility of using a diffuser for pumping the MPD engine exhaust is presented here. This analysis uses a detailed equation of state based on the partition function for the high temperature argon exhaust. On the basis of the electron-ion momentum exchange collision frequency in the exhaust plasma, it was concluded that the diffuser gas dynamics could be modeled to a first approximation with ordinary continuum equations. Calculations of the stagnation pressure in the diffuser, downstream of a strong normal shock, yielded pressures on the order of 10 torr suggesting that the diffuser is feasible for pumping the MPD engine exhaust.
Attention is given to the design features of a radiation-cooled, 30-kW thermal arcjet thruster, whose laboratory tests have yielded specific impulses of up to 935 sec at 36-44 percent thrust efficiency, together with a cumulative lifetime of over 400 hours. All materials used, including seals, can sustain operation at temperatures sufficiently elevated to require the radiation of all waste heat. This electric propulsion system is ideally suited to missions such as the Space-Based Radar. A detailed consideration is conducted for the thruster's seals, which are the most critical element of the design.