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At least 1,045 records · Page 58
All sourcesNASA NTRSDOE OSTI
NASA NTRS1989

Solar dynamic power for Space Station Freedom

The Space Station Freedom Program is presently planned to consist of two phases. At the completion of Phase 1, Freedom's manned base will consist of a transverse boom with attached manned modules and 75 kW of available electric power supplied by photovoltaic (PV) power sources. In Phase 2, electric power available to the manned base will be increased to 125 kW by the addition of two solar dynamic (SD) power modules, one at each end of the transverse boom. Power for manned base growth beyond Phase 2 will be supplied by additional SD modules. Studies show that SD power for the growth eras will result in life cycle cost savings of $3 to $4 billion when compared to PV-supplied power. In the SD power modules for Space Station Freedom, an offset parabolic concentrator collects and focuses solar energy into a heat receiver. To allow full power operation over the entire orbit, the receiver includes integral thermal energy storage by means of the heat of fusion of a salt mixture. Thermal energy is removed from the receiver and converted to electrical energy by a power conversion unit (PCU) which includes a closed brayton cycle (CBC) heat engine and an alternator. The receiver/PCU/radiator combination will be completely assembled and charged with gas and cooling fluid on earth before launch to orbit. The concentrator subassemblies will be pre-aligned and stowed in the orbiter bay before launch. On orbit, the receiver/PCU/radiator assembly will be installed as a unit. The pre-aligned concentrator panels will then be latched together and the total concentrator attached to the receiver/PCU/radiator by the astronauts. After final electric connections are made and checkout is complete, the SD power module will be ready for operation.

Labus, Thomas L.↗
NASA NTRS1989

An RF heated tandem mirror plasma propulsion study

Experimental results on a tandem mirror hybrid plume rocket involving a three-stage system of plasma injection, heating, and subsequent injection through a magnetic nozzle are presented. In the experiments, a plasma is created by breaking down the gas with electron cyclotron resonance heating at 2 kW in the central cell, and the ion species is then heated to high temperatures with ion cyclotron resonance heating at 10 kW in the end cell. A Langmuir probe measured an electron density of 2.5 x 10 to the 16th/cu m and a temperature of 100 eV in the central cell and an ion density of 1.25 x 10 to the 17th/cu m and a temperature of 500 eV in the end cell.

Yang, T. F.↗
NASA NTRS1989

Laser-powered Martian rover

Two rover concepts were considered: an unpressurized skeleton vehicle having available 4.5 kW of electrical power and limited to a range of about 10 km from a temporary Martian base and a much larger surface exploration vehicle (SEV) operating on a maximum 75-kW power level and essentially unrestricted in range or mission. The only baseline reference system was a battery-operated skeleton vehicle with very limited mission capability and range and which would repeatedly return to its temporary base for battery recharging. It was quickly concluded that laser powering would be an uneconomical overkill for this concept. The SEV, on the other hand, is a new rover concept that is especially suited for powering by orbiting solar or electrically pumped lasers. Such vehicles are visualized as mobile habitats with full life-support systems onboard, having unlimited range over the Martian surface, and having extensive mission capability (e.g., core drilling and sampling, construction of shelters for protection from solar flares and dust storms, etc.). Laser power beaming to SEV's was shown to have the following advantages: (1) continuous energy supply by three orbiting lasers at 2000 km (no storage requirements as during Martian night with direct solar powering); (2) long-term supply without replacement; (3) very high power available (MW level possible); and (4) greatly enhanced mission enabling capability beyond anything currently conceived.

Harries, W. L.↗
NASA NTRS1989

Electrical performance characteristics of high power converters for space power applications

The first goal of this project was to investigate various converters that would be suitable for processing electric power derived from a nuclear reactor. The implementation is indicated of a 20 kHz system that includes a source converter, a ballast converter, and a fixed frequency converter for generating the 20 kHz output. This system can be converted to dc simply by removing the fixed frequency converter. This present study emphasized the design and testing of the source and ballast converters. A push-pull current-fed (PPCF) design was selected for the source converter, and a 2.7 kW version of this was implemented using three 900 watt modules in parallel. The characteristic equation for two converters in parallel was derived, but this analysis did not yield any experimental methods for measuring relative stability. The three source modules were first tested individually and then in parallel as a 2.7 kW system. All tests proved to be satisfactory; the system was stable; efficiency and regulation were acceptable; and the system was fault tolerant. The design of a ballast-load converter, which was operated as a shunt regulator, was investigated. The proposed power circuit is suitable for use with BJTs because proportional base drive is easily implemented. A control circuit which minimizes switching frequency ripple and automatically bypasses a faulty shunt section was developed. A nonlinear state-space-averaged model of the shunt regulator was developed and shown to produce an accurate incremental (small-signal) dynamic model, even though the usual state-space-averaging assumptions were not met. The nonlinear model was also shown to be useful for large-signal dynamic simulation using PSpice.

Stuart, Thomas A.↗
NASA NTRS1989

Design of a thrust stand for high power electric propulsion devices

A thrust stand for use with high power electric propulsion devices was designed and tested. The thrust stand was specifically tailored to the needs of a 100 to 250 kW magnetoplasmadynamic (MPD) thruster program currently in progress at the NASA Lewis Research Center. The thrust stand structure was built as an inverted pendulum arrangement, supported at the base by water-cooled electrical power flexures. Thrust stand tares due to thruster discharge current were demonstrated to be negligible. Tares due to an applied field magnet current, after considerable effort, were reduced to less than 3.0 percent of measured thrust. These tares, however, could be determined independently and subtracted from the indicated thrust measurement. A detailed description is given for the thrust stand design and operation with a 100 kW class MPD device. Other thrust stand tares due to vibration and thermal effects are discussed, along with issues of accuracy and repeatability.

Haag, Thomas W.↗
NASA NTRS1989

The advanced photovoltaic solar array program

The background and development status of an ultralightweight flexible-blanket flatpack, fold-out solar array is presented. It is scheduled for prototype demonstration in late 1989. The Advanced Photovoltaic Solar Array (APSA) design represents a critical intermediate milestone of the goal of 300 W/kg at beginning-of-life (BOL) with specific performance characteristics of 130 W/kg (BOL) and 100 W/kg at end-of-life (EOL) for a 10-year geosynchronous geostationary earth orbit 10-kW (BOL) space power system. The APSA wing design is scalable over a power range of 2 to 15 kW and is suitable for a full range of missions including Low Earth Orbit (LEO), orbital transfer from LEO to geostationary earth orbit and interplanetary flight.

Kurland, R. M.↗
NASA NTRS1989

Arcjet Cathode Phenomena

Cathode tips made from a number of different materials were tested in a modular arcjet thruster in order to examine cathode phenomena. Periodic disassembly and examination, along with the data collected during testing, indicated that all of the tungsten-based materials behaved similarly despite the fact that in one of these samples the percentage of thorium oxide was doubled and another was 25 percent rhenium. The mass loss rate from a 2 percent thoriated rhenium cathode was found to be an order of magnitude greater than that observed using 2 percent thoriated tungsten. Detailed analysis of one of these cathode tips showed that the molten crater contained pure tungsten to a depth of about 150 microns. Problems with thermal stress cracking were encountered in the testing of a hafnium carbide tip. Post test analysis showed that the active area of the tip had chemically reacted with the propellant. A 100 hour continuous test was run at about 1 kW. Post test analysis revealed no dendrite formation, such as observed in a 30 kW arcjet lifetest, near the cathode crater. The cathodes from both this test and a previously run 1000 hour cycled test displayed nearly identical arc craters. Data and calculations indicate that the mass losses observed in testing can be explained by evaporation.

Curran, Francis M.↗
NASA NTRS1990

Arcjet load characteristics

Experiments were conducted to define the interface characteristics and constraints of 1 kW class arcjets run on simulated decomposition products of hydrazine and power processors. The impacts of power supply output current ripple on arcjet performance were assessed by variation of the ripple frequency from 100 Hz to 100 kHz with 10 percent peak-to-peak ripple amplitude at 1.2 kW. Ripple had no significant effects on thrust, specific impulse or efficiency. The impact of output ripple on thruster lifetime was not assessed. The static and dynamic impedances of the arcjet were quantified with two thrusters of nearly identical configuration. Superposition of an AC component on the DC arc current was used to characterize the dynamic impedance as a function of flow rate and DC current level. A mathematical model was formulated from these data. Both the static and dynamic impedance magnitude were found to be dependent on mass flow rate. The amplitude of the AC component was found to have little effect on the dynamic impedance. Reducing the DC level from 10 to 8 amps led to a large change in the magnitude of the dynamic impedance with no observable phase change. The impedance data compared favorably between the two thrusters.

Hamley, John A.↗
NASA NTRS1989

Designing Space Station Freedom for evolution

The requirements for using the Space Station for research and development and to support human missions to the Moon and Mars are examined. Space Station configurations are presented for these two cases. It is suggested that, by the year 2010, the Space Station requirements for research and development should include 275 kW of average power, a crew of 24, 5 full laboratory modules, and 3 pocket labs. For supporting the Lunar/Mars program as a transportation node facility, the Space Station should require 175 kw of average power, 16 crew members, 3 full laboratory modules, and 1 pocket lab.

Huckins, Earle K., III↗
NASA NTRS1989

Effect of advanced component technology on helicopter transmissions

Experimental tests were performed on the NASA/Bell Helicopter Textron (BHT) 500 hp advanced technology transmission (ATT) at the NASA Lewis Research Center. The ATT was a retrofit of the OH-58C helicopter 236 kW (317 hp) main rotor transmission, upgraded to 373 kW (500 hp), with a design goal of retaining long life with a minimum increase in cost, weight, and size. Vibration, strain, efficiency, deflection, and temperature experiments were performed and the results were compared to previous experiments on the OH-58A, OH-58C, and UH-60A transmissions. The high-contact-ratio gears and the cantilevered-mounted, flexible ring gear of the ATT reduced vibration compared to that of the OH-58C. The ATT flexible ring gear improved planetary load sharing compared to that of the rigid ring gear of the UH-60A transmission. The ATT mechanical efficiency was lower than that of the OH-58A transmission, probably due to the high-contact-ratio planetary gears.

Lewicki, D. G.↗
NASA NTRS1989

Space Station Freedom photovoltaic power module design status

Electric power generation for the Space Station Freedom will be provided by four photovoltaic (PV) power modules using silicon solar cells during phase I operation. Each PV power module requires two solar arrays with 32,800 solar cells generating 18.75 kW of dc power for a total of 75 kW. A portion of this power will be stored in nickel-hydrogen batteries for use during eclipse, and the balance will be processed and converted to 20 kHz ac power for distribution to end users through the power management and distribution system. The design incorporates an optimized thermal control system, pointing and tracking provision with the application of gimbals, and the use of orbital replacement units to achieve modularization. The design status of the PV power module, as derived from major trade studies, is discussed at hardware levels ranging from component to system. Details of the design are presented where appropriate.

Jimenez, Amador P.↗
NASA NTRS1989

Space Station Freedom solar array design development

The Space Station Freedom solar array program is required to provide a 75-kW power module that uses eight solar array (SA) wings over a four-year period in low earth orbit (LEO). Each wing will be capable of providing 23.4 kW at the 4-yr design point. The design of flexible-substrate SAs that must survive exposure to the space environment, including atomic oxygen, for an operating life of fifteen years is discussed. The tradeoff study and development areas being investigated include solar cell module size, solar cell weld pads, panel stiffener frames, materials inherently resistant to atomic oxygen, and weight reduction design alternatives.

Winslow, Cindy↗
NASA NTRS1989

Advanced Photovoltaic Solar Array program status

The Advanced Photolvoltaic Solar Array (APSA) Program is discussed. The objective of the program is to demonstrate a producible array system by the end of this decade with a beginning-of-life (BOL) specific power of 130 W/kg at 10 kW as an intermediate milestone toward the ultimate goal of 300 W/kg at 25 kW by the year 2000. The near-term goal represents a significant improvement over existing rigid panel flight arrays (25 to 45 W/kg) and the first-generation flexible blanket NASA/OAST SAFE I array of the early 1980s, which was projected to provide about 60 W/kg BOL. The prototype wing hardware is in the last stages of fabrication and integration. The current status of the program is reported. The array configuration and key design details are shown. Projections are shown for future performance enhancements that may be expected through the use of advanced structural components and solar cells.

Kurland, Richard M.↗
NASA NTRS1990

Temperature measurements of a high-power microwave feedhorn window

Temperature measurements of a high-power microwave feedhorn window, obtained using an imaging IR radiometer during transmitter operation at 365 kW CW and 8.5 GHz, are discussed. The window under investigation was constructed of HTP-6, a high-thermal-performance material developed to shield the Space Shuttle Orbiter from the heat of reentry. The measurement technique is described, and experimental results are presented. The window performed adequately at 365 kW CW with a center temperature of 475 C. The tests verify that HTP-6 can be used as a window material or a support structure in high-power waveguides at power densities of 1.47 kW/sq cm for extended periods of time, with no change in its mechanical characteristics.

Hoppe, Daniel J.↗
NASA NTRS1990

Arcjet load characteristics

Experiments were conducted to define the interface characteristics and constraints of 1 kW class arcjets run on simulated decomposition products of hydrazine and power processors. The impacts of power supply output current ripple on arcjet performance were assessed by variation of the ripple frequency from 100 Hz to 100 kHz with 10 percent peak-to-peak ripple amplitude at 1.2 kW. Ripple had no significant effects on thrust, specific impulse or efficiency. The impact of output ripple on thruster lifetime was not assessed. The static and dynamic impedances of the arcjet were quantitied with two thrusters of nearly identical configuration. Superposition of an AC component on the DC arc current was used to characterize the dynamic impedance as a function of flow rate and DC current level. A mathematical model was formulated from these data. Both the static and dynamic impedance magnitude were found to be dependent on mass flow rate. The amplitude of the AC component was found to have little effect on the dynamic impedance. Reducing the DC level from 10 to 8 amps led to a large change in the magnitude of the dynamic impedance with no observable phase change. The impedance data compared favorably between the two thrusters.

Hamley, John A.↗
NASA NTRS1991

Solar dynamic heat receiver technology

A full-size, solar dynamic heat receiver was designed to meet the requirements specified for electrical power modules on the U.S. Space Station, Freedom. The heat receiver supplies thermal energy to power a heat engine in a closed Brayton cycle using a mixture of helium-xenon gas as the working fluid. The electrical power output of the engine, 25 kW, requires a 100 kW thermal input throughout a 90 minute orbit, including when the spacecraft is eclipsed for up to 36 minutes from the sun. The heat receiver employs an integral thermal energy storage system utilizing the latent heat available through the phase change of a high-temperature salt mixture. A near eutectic mixture of lithium fluoride and calcium difluoride is used as the phase change material. The salt is contained within a felt metal matrix which enhances heat transfer and controls the salt void distribution during solidification. Fabrication of the receiver is complete and it was delivered to NASA for verification testing in a simulated low-Earth-orbit environment. This document reviews the receiver design and describes its fabrication history. The major elements required to operate the receiver during testing are also described.

Sedgwick, Leigh M.↗
NASA NTRS1991

Sensible heat receiver for solar dynamic space power system

A sensible heat receiver considered in this study uses a vapor grown carbon fiber-carbon (VGCF/C) composite as the thermal storage media and was designed for a 7 kW Brayton engine. The proposed heat receiver stores the required energy to power the system during eclipse in the VGCF/C composite. The heat receiver thermal analysis was conducted through the Systems Improved Numerical Differencing Analyzer and Fluid Integrator (SINDA) software package. The sensible heat receiver compares well with other latent and advanced sensible heat receivers analyzed in other studies while avoiding the problems associated with latent heat storage salts and liquid metal heat pipes. The concept also satisfies the design requirements for a 7 kW Brayton engine system. The weight and size of the system can be optimized by changes in geometry and technology advances for this new material.

Perez-Davis, Marla E.↗