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At least 1,081 records · Page 60
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Variable-speed Generators with Flux Weakening

A cost-competitive, permanent-magnet 20 kW generator is designed such that the following criteria are satisfied: an (over) load capability of at least 30 kW over the entire speed range of 60-120 rpm, generator weight of about 550 lbs with a maximum radial stator flux density of 0.82 T at low speed, unity power factor operation, acceptably small synchronous reactances and operation without a gear box. To justify this final design four different generator designs are investigated: the first two designs are studied to obtain a speed range from 20 to 200 rpm employing rotor field weakening, and the latter two are investigated to obtain a maximum speed range of 40 to 160 rpm based on field weakening via the stator excitation. The generator reactances and induced voltages are computed using finite element/difference solutions. Generator losses and efficiencies are presented for all four designs at rated temperature of Tr=120C.

Fardoun, A. A.↗
NASA NTRS1993

NEP technology: FY 1992 milestones (NASA LeRC)

A discussion of Nuclear Electric Propulsion (NEP) thrusters and facilities is presented in vugraph form. The NEP thrusters are discussed in the context of the following three items: (1) establishing a 100 H test capability for 100-kW magnetoplasmadynamic (MPD) thrusters; (2) demonstrating a lightweight 20-kW krypton ion thruster; and (3) the optimization of the design of low-mass power processor transformers. The primary accomplishment at NEP facilities was the completion of the Electric Propulsion Laboratory's (EPL's) tank 5 cryopump upgrade.

Sovey, Jim↗
NASA NTRS1993

Evaluation of actuator energy storage and power sources for spacecraft applications

The objective of this evaluation is to determine an optimum energy storage/power source combination for electrical actuation systems for existing (Solid Rocket Booster (SRB), Shuttle) and future (Advanced Launch System (ALS), Shuttle Derivative) vehicles. Characteristic of these applications is the requirement for high power pulses (50-200 kW) for short times (milliseconds to seconds), coupled with longer-term base or 'housekeeping' requirements (5-16 kW). Specific study parameters (e.g., weight, volume, etc.) as stated in the proposal and specified in the Statement of Work (SOW) are included.

Simon, William E.↗
NASA NTRS1993

Preliminary investigation of high power microwave plasmas for electrothermal thruster use

Results are reported from preliminary tests to evaluate the high power microwave electrothermal thruster (MET) concept, which employs a free-floating plasma discharge maintained by applied CW microwave power to heat a propellant gas flow. Stable plasmas have been created and maintained in helium (He), nitrogen (N2), and hydrogen (H2) as propellants in both the TM(sub 011) and TM(sub 012) modes at discharge pressures from 10 Pa to 69 kPa. Reproducible starting conditions of pressure and power have been documented for all the plasmas. Vortical inflow of the propellant gas was observed to cause the formation of on-axis 'spike' plasmas. The formation and unformation conditions of these plasmas were studied. Operation in the spike plasma condition enables maximum power absorption with minimum wall heating and offers maximum efficiency in heating the propellant gas. In the spike condition, plasmas of the three propellant gases were investigated in an open channel configuration to a maximum applied power level of 11.2 kW (in N2). Microwave power coupling efficiencies of over 90 percent were routinely obtained at absorbed power levels up to 2 kW. Magnetic nozzle effects were investigated with a superconducting solenoid Al magnet applying a high magnetic field to the plasmas in and exiting from the discharge tube.

Power, John L.↗
NASA NTRS1992

Low-Isp derated ion thruster operation

The performance and lifetime expectations of 30 cm xenon ion thruster technology at low values of specific impulse were evaluated, with emphasis on 1000-2500 s operation. Power levels of up to 2.0 kW, appropriate for auxiliary and orbit maneuvering propulsion, were processed at thrust-to-power ratios up to 57 mN/kW. These tests were conducted using a derated 30 cm ion thruster with high-perveance design two-grid ion optics with xenon propellant. Lifetime projections were made based on a simple analysis of critical component erosion rates, and it was found that a strong correlation exists with the ratio of the specific impulse-to-input power. Under all operating conditions for which the projected thruster lifetime is less than 10,000 hrs, the life-limiting component of this technology is erosion of the accelerator grid due to charge-exchange ions. The use of alternative grid materials such as carbon is estimated to increase useful thruster lifetimes by as much as an order of magnitude and may enable long-life high thrust-density, sub-2500 s Isp operation. The performance and life of the derated thruster appears similar to that of the Russian SPT-100 thruster in the 1.0-2.0 kW, 1600-2000 s operational envelope.

Patterson, Michael J.↗
NASA NTRS1991

NASA mission planning for space nuclear power

An evaluation is conducted of those aspects of the Space Exploration Initiative which stand to gain from the use of nuclear powerplants. Low-power, less than 10 kW(e) missions in question encompass the Comet Rendezvous Asteroid Flyby, the Cassini mission to Saturn, the Mars Network mission, a solar probe, the Mars Rover Sample Return mission, the Rosetta comet nucleus sample return mission, and an outer planets orbiter/probe. Reactor power yielding 10-100 kW(e) can be used by advanced rovers and initial lunar and Martian outposts, as well as Jovian and Saturnian grand tours and sample-return missions.

Bennett, Gary L.↗
NASA NTRS1993

Investigation of a subsonic-arc-attachment thruster using segmented anodes

To investigate high frequency arc instabilities observed in subsonic-arc-attachment thrusters, a 3 kW, segmented-anode arcjet was designed and tested using hydrogen as the propellant. The thruster nozzle geometry was scaled from a 30 kW design previously tested in the 1960's. By observing the current to each segment and the arc voltage, it was determined that the 75-200 kHz instabilities were results of axial movements of the arc anode attachment point. The arc attachment point was fully contained in the subsonic portion of the nozzle for nearly all flow rates. The effects of isolating selected segments were investigated. In some cases, forcing the arc downstream caused the restrike to cease. Finally, decreasing the background pressure from 18 Pa to 0.05 Pa affected the pressure distribution in the nozzle, including the pressure in the subsonic arc chamber.

Berns, Darren H.↗
NASA NTRS1993

Preliminary investigation of high power microwave plasmas for electrothermal thruster use

Results are reported from preliminary tests to evaluate the high power microwave electrothermal thruster concept, which employs a free-floating plasma discharge maintained by applied CW microwave power to heat a propellant gas flow. Stable plasmas have been created and maintained in He, N2, and H2 as propellants in both the TM(sub 011) and TM(sub 012) modes at discharge pressures from 10 Pa to 69 kPa. Vortical inflow of the propellant gas was observed to cause the formation of on-axis 'spike' plasmas. Operation in the spike plasma condition yields maximum power absorption with minimum wall heating and maximum propellant-heating efficiency; plasmas of the three propellant gases were investigated in the spike condition in an open channel to a maximum applied power level of 11.2 kW (in N2). Microwave power coupling efficiencies of over 90 percent were routinely obtained at absorbed power levels up to 2 kW. Magnetic nozzle effects were investigated with a superconducting solenoid Al magnet applying a high magnetic field to the plasmas in, and exiting from, the discharge tube.

Power, John L.↗
NASA NTRS1993

Simplified power processing for inert gas ion thrusters

Significant simplifications to power processors for inert gas ion thrusters in the 1 to 5 kW range have been identified. They include elimination of all but three power supplies - one each for the neutralizer, main discharge, and beam. The neutralizer and discharge power supplies would provide both cathode heating and plasma generating functions. This dual-use power supply concept was validated via integration tests with a 30 cm diameter xenon ion thruster. The beam/accelerator power supply would have positive and negative outputs to allow a single power supply to provide both functions. The discharge and beam power supplies would incorporate full-bridge inverters similar to those proven for flight-ready arcjet propulsion systems. Operation of this simplified power processing scheme at an inverter frequency of 50 kHz results in a projected power processor design with low mass and high efficiency. A 2 kW reference point design has estimated values of specific mass of 5.4 kg/kW and an efficiency of 93 percent.

Rawlin, V. K.↗
NASA NTRS1993

BMDO electric space-propulsion program

Electric propulsion (EP) applications being considered include: orbit insertion, orbit repositioning, station keeping, and elusive maneuvering. Typically, 1 to 5 kW are available for EP. A class of thrusters, the Hall-effect thrusters, is extensively researched, developed and flown by Russia. These thrusters, using xenon propellant, perform reliably, e.g., at 1.35 kW, 600 s specific impulse, 50 percent efficiency and greater than 2000 hr life. This specific impulse and efficiency combination is superior to the present arcjets for several Ballistic Missile Defense Organization (BMDO) applications. Three versions of the Hall thrusters are part of the experimental evaluation. Since today's goals are within reach of available thrusters and power sources, emphasis is being placed on such topics as: thruster lifetime, spacecraft interactions, electromagnetic interference, and erosion product deposition. Facilities in U.S. laboratories are being specially configured to achieve these goals.

Caveny, Leonard H.↗
NASA NTRS1988

Summary of aerospace and nuclear engineering activities

The Texas A&M Nuclear and Aerospace engineering departments have worked on five different projects for the NASA/USRA Advanced Design Program during the 1987/88 year. The aerospace department worked on two types of lunar tunnelers that would create habitable space. The first design used a heated cone to melt the lunar regolith, and the second used a conventional drill to bore its way through the crust. Both used a dump truck to get rid of waste heat from the reactor as well as excess regolith from the tunneling operation. The nuclear engineering department worked on three separate projects. The NEPTUNE system is a manned, outer-planetary explorer designed with Jupiter exploration as the baseline mission. The lifetime requirement for both reactor and power-conversion systems was twenty years. The second project undertaken for the power supply was a Mars Sample Return Mission power supply. This was designed to produce 2 kW of electrical power for seven years. The design consisted of a General Purpose Heat Source (GPHS) utilizing a Stirling engine as the power conversion unit. A mass optimization was performed to aid in overall design. The last design was a reactor to provide power for propulsion to Mars and power on the surface. The requirements of 300 kW of electrical power output and a mass of less than 10,000 Rg were set. This allowed the reactor and power conversion unit to fit within the Space Shuttle cargo bay.

Source record↗
NASA NTRS1990

Space Station thermal control system evolution

The thermal control system (TCS) for the space station assembly complete configuration includes a two-phase central thermal bus with a supplemental body mounted radiator system. Evolution of the space station from a heat rejection capacity of 75 kW to 300 kW will require scars to expand the thermal fluid distribution network, equipment replacement to enable greater thermal transport capacity, and enlargement of the heat rejection subsystem for increased heat rejection. The TCS requirements for assembly completion and growth are presented along with a review of the basic structure of the active and passive thermal control systems which include provisions for growth.

Bullock, Richard L.↗
NASA NTRS1993

SPDE/SPRE final summary report

Mechanical Technology Incorporated (MTI) performed acceptance testing on the Space Power Research Engine (SPRE), which demonstrated satisfactory operation and sufficient reliability for delivery to NASA Lewis Research Center. The unit produced 13.5 kW PV power with an efficiency of 22 percent versus design goals of 28.8 kW PV power and efficiency of 28 percent. Maximum electric power was only 8 kWe due to lower alternator efficiency. One of the major shortcomings of the SPRE was linear alternator efficiency, which was only 70 percent compared to a design value of 90 percent. It was determined from static tests that the major cause for the efficiency shortfall was the location of the magnetic structure surrounding the linear alternator. Testing of an alternator configuration without a surrounding magnetic structure on a linear dynamometer confirmed earlier static test results. Linear alternator efficiency improved from 70 percent to over 90 percent. Testing of the MTI SPRE was also performed with hydrodynamic bearings and achieved full-stroke, stable operation. This testing indicated that hydrodynamic bearings may be useful in free piston Stirling engines. An important factor in achieving stable operation at design stroke was isolating a portion of the bearing length from the engine pressure variations. In addition, the heat pipe heater head design indicates that integration of a Stirling engine with a heat source can be performed via heat pipes. This design provides a baseline against which alternative designs can be measured.

Dochat, George↗
NASA NTRS1993

Segmented ion thruster

Apparatus and methods for large-area, high-power ion engines comprise dividing a single engine into a combination of smaller discharge chambers (or segments) configured to operate as a single large-area engine. This segmented ion thruster (SIT) approach enables the development of 100-kW class argon ion engines for operation at a specific impulse of 10,000 s. A combination of six 30-cm diameter ion chambers operating as a single engine can process over 100 kW. Such a segmented ion engine can be operated from a single power processor unit.

Brophy, John R.↗
NASA NTRS1993

Power console development for NASA's electric propulsion outreach program

NASA LeRC is developing a 30 cm diameter xenon ion thruster for auxiliary and primary propulsion applications. To maximize expectations for user-acceptance of ion propulsion technology, NASA LeRC, through their Electric Propulsion Outreach Program, is providing sectors of industry with portable power consoles for operation of 5 KW-class xenon ion thrusters. This power console provides all necessary functions to permit thruster operations over a 0.5-5 KW envelope under both manual and automated control. These functions include the following: discharge, cathode heater, neutralizer keeper, and neutralizer heater currents, screen and accelerator voltages, and a gas feed system to regulate and control propellant flow to the thruster. An electronic circuit monitors screen and accelerator currents and controls arcing events. The power console was successfully integrated with the NASA 30 cm thruster.

Pinero, Luis R.↗
NASA NTRS1993

Investigation of a subsonic-arc-attachment thruster using segmented anodes

To investigate high frequency arc instabilities observed in subsonic-arc-attachment thrusters, a 3 kW, segmented-anode arc jet was designed and tested using hydrogen as the propellant. The thruster nozzle geometry was scaled from a 30 kW design previously tested in the 1960's. By observing the current to each segment and the arc voltage, it was determined that the 75-200 kHz instabilities were results of axial movements of the arc anode attachment point. The arc attachment point was fully contained in the subsonic portion of the nozzle for nearly all flow rates. The effects of isolating selected segments were investigated. In some cases, forcing the arc downstream caused the restrike to cease. Finally, decreasing the background pressure from 18 to 0.05 Pa affected the pressure distribution in the nozzle including the pressure in the subsonic arc chamber.

Berns, Darren H.↗
NASA NTRS1992

Pressurized Lunar Rover (PLR)

The objective of this project was to design a manned pressurized lunar rover (PLR) for long-range transportation and for exploration of the lunar surface. The vehicle must be capable of operating on a 14-day mission, traveling within a radius of 500 km during a lunar day or within a 50-km radius during a lunar night. The vehicle must accommodate a nominal crew of four, support two 28-hour EVA's, and in case of emergency, support a crew of six when near the lunar base. A nominal speed of ten km/hr and capability of towing a trailer with a mass of two mt are required. Two preliminary designs have been developed by two independent student teams. The PLR 1 design proposes a seven meter long cylindrical main vehicle and a trailer which houses the power and heat rejection systems. The main vehicle carries the astronauts, life support systems, navigation and communication systems, lighting, robotic arms, tools, and equipment for exploratory experiments. The rover uses a simple mobility system with six wheels on the main vehicle and two on the trailer. The nonpressurized trailer contains a modular radioisotope thermoelectric generator (RTG) supplying 6.5 kW continuous power. A secondary energy storage for short-term peak power needs is provided by a bank of lithium-sulfur dioxide batteries. The life support system is partly a regenerative system with air and hygiene water being recycled. A layer of water inside the composite shell surrounds the command center allowing the center to be used as a safe haven during solar flares. The PLR 1 has a total mass of 6197 kg. It has a top speed of 18 km/hr and is capable of towing three metric tons, in addition to the RTG trailer. The PLR 2 configuration consists of two four-meter diameter, cylindrical hulls which are passively connected by a flexible passageway, resulting in the overall vehicle length of 11 m. The vehicle is driven by eight independently suspended wheels. The dual-cylinder concept allows articulated as well as double Ackermann steering. The primary power of 8 kW is supplied by a dynamic isotope system using a closed Brayton cycle with a xenon-hydrogen mixture as the working fluid. A sodium-sulfur battery serves as the secondary power source. Excess heat produced by the primary power system and other rover systems is rejected by radiators located on the top of the rear cylinder. The total mass of the PLR 2 is 7015 kg. Simplicity and low total weight have been the driving principles behind the design of PLR 1. The overall configuration consists of a 7-m-long, 3-m-diameter cylindrical main vehicle and a two-wheeled trailer. The cylinder of the main body is capped by eight-section, faceted, semi-hemispherical ends. The trailer contains the RTG power source and is not pressurized. The shell of the main body is constructed of a layered carbon fiber/foam/Kevlar sandwich structure. Included in the shell is a layer of water for radiation protection. The layer of water extends from the front of the rover over the crew compartment and creates a safe haven for the crew during a solar flare-up. The carbon fiber provides the majority of the strength and stiffness and the Kevlar provides protection from micrometeoroids. The Kevlar is covered with a gold foil and multi-layer insulation (MLI) to reduce radiation degradation and heat transfer through the wall. A thin thermoplastic layer seals the fiber and provides additional strength.

Creel, Kenneth↗
NASA NTRS1994

Development Status of the NASA 30-cm Ion Thruster and Power Processor

Xenon ion propulsion systems are being developed by NASA Lewis Research Center and the Jet Propulsion Laboratory to provide flight qualification and validation for planetary and earth-orbital missions. In the ground-test element of this program, light-weight (less than 7 kg), 30 cm diameter ion thrusters have been fabricated, and preliminary design verification tests have been conducted. At 2.3 kW, the thrust, specific impulse, and efficiency were 91 mN, 3300 s, and 0.65, respectively. An engineering model thruster is now undergoing a 2000 h wear-test. A breadboard power processor is being developed to operate from an 80 V to 120 V power bus with inverter switching frequencies of 50 kHz. The power processor design is a pathfinder and uses only three power supplies. The projected specific mass of a flight unit is about 5 kg/kW with an efficiency of 0.92 at the full-power of 2.5 kW. Preliminary integration tests of the neutralizer power supply and the ion thruster have been completed. Fabrication and test of the discharge and beam/accelerator power stages are underway.

Sovey, James S.↗