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At least 955 records · Page 53

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NASA NTRS1972

A ten-stage electrostatic depressed collector for improving klystron efficiency.

A ten-stage electrostatic depressed collector, designed with the aid of an analog computer, was tested on a 1-kW CW 750-MHz klystron. Excellent correlation was achieved between computed and measured performance under varying conditions of RF drive. At full RF power output approximately 60 per cent of the spent beam energy was recovered by use of the depressed collector. The net power conversion efficiency of the tube was raised from its undepressed value of 54.3 per cent to approximately 70.9 per cent. At one-half full power output, a collector efficiency of 70 per cent was measured. At zero RF power output collector efficiency was 80 per cent. To achieve these results it was necessary to install a small focusing coil between the final drift tube and the collector. No spurious oscillations or instabilities were detected when collector depression was employed, nor was electron backstreaming increased significantly. Intentional short circuiting of adjacent collector electrode pairs was shown to cause only minor degradation in collector performance.

Neugebauer, W.↗

NASA NTRS1972

Mission analysis for design of a SEP multi-mission spacecraft.

Study of the multimission capability of a solar electric propulsion spacecraft. The study includes such factors as spacecraft constraints, payload variations with launch opportunity, propellant on-loading and off-loading requirements, and limitations in thrusting time, throttling range and steering. The study determined that a single spacecraft design, with P sub zero = 12 kW and I sub s = 3000 sec launched by a Titan IIIC or Titan IIID/Centaur can be used to conduct at least five diverse missions: Eros rendezvous, 1 AU out-of-ecliptic, Mercury orbiter, Jupiter flyby/probe, and Jupiter orbiter. Flight time and injection velocity vary as needed to achieve the desired payload capability. The mission analysis approach and results are presented.

Driver, J. M.↗

NASA NTRS1972

Solar electric low-thrust Mercury orbiter missions.

Results of a preliminary study of unmanned Mercury orbiter missions which might be flown in the early 1980s. Five different levels of orbiter size and complexity are considered, ranging from a small orbiter carrying 22 kg of low data-rate instruments to a dual satellite mission employing 110 kg of science instruments emphasizing regional and local scale surface imagery. An Intermediate-20 class launch vehicle is required to deliver a minimal surface-imaging science package into Mercury orbit using a ballistic mission mode. Employment of a 15 kW solar electric low-thrust stage permits delivery of the more capable dual satellite package with a Titan IIID (5)/Centaur.

Klopp, D. A.↗

NASA NTRS1972

Optimal Quasi-steady Plasma Thruster system characteristics.

The overall characteristics of a generalized Quasi-steady Plasma Thruster (QPT) system consisting of thruster head, power conditioning network, propellant supply subsystem are studied. Energy balance equations for the system are coupled with component mass relationships in order to determine overall system mass and performance. Power supply power levels varying from 100 to 10,000 watts with thruster power levels ranging from 300 kw to 30 Mw employing argon as the propellant are considered. The manner in which overall system mass, average thrust, and burn time vary as a function power supply power level, quasi-steady power level, and pulse time are studied. Results indicate the existence of optimum pulse times when system mass is employed as an optimization criterion.

Ludwig, D. E.↗

NASA NTRS1972

A common solar electric upper stage for multi-mission applications.

The concept of using a solar-electric upper stage for multi-mission application is practical, economical, and versatile approach to space expMission capabilities include a Mercury orbiter, asteroid and comet rendezvous, a solar probe, and extra-ecliptic probe, and outer planet flybys and orbiters. The 17.5-kw vehicle has an injected mass of 1500 to 2500 kg and carries up to 500 kg of attached or separable payload packages. Mission characteristics, performance tradeoffs and a concepegn configuration are presented, and program plans, cost estimates and advanced technology requirements are identified.

Meissinger, H. F.↗

NASA NTRS1972

A high-power communications technology satellite for the 12 and 14 GHz bands.

Description of the mission, subsystems and communication capabilities of a joint Communications Technology Satellite (CTS) scheduled for launch in 1975 by the Canadian Department of Communications in cooperation with NASA. The principal objectives of the mission are TV transmission at 12 GHz to low-cost ground terminals, up-link TV transmission at 14 GHz transportable terminals, and flight tests of spacecraft subsystems and components for future communications satellites. The major advanced spacecraft subsystems are a novel superefficiency TWT design, a 0.4 mlb Mercury Bombardment ion engine for north-south station keeping, a 3-axis stabilization system to maintain a high antenna boresight pointing accuracy, a liquid metal slip ring experiment, and a lightweight extendible solar array with an initial power output greater than 1 kW.

Franklin, C. A.↗

NASA NTRS1972

Propulsion requirements for communications satellites.

The concept of characteristics thrust is introduced herein as a means of classifying propulsion system tasks related particularly to geosynchronous communications spacecraft. Approximate analytical models are developed to permit estimation of characteristic thrust for injection error corrections, orbit angle re-location, north-south station keeping, east-west station keeping, spin axis precession control, attitude rate damping, and orbit raising applications. Performance assessment factors are then outlined in terms of characteristic power, characteristic weight, and characteristic volume envelope, which are related to the characteristic thrust. Finally, selected performance curves are shown for power as a function of spacecraft weight, including the influence of duty cycle on north-south station keeping, a 90 degree orbit angle re-location in 14 days, and finally comparison of orbit raising tasks from low and intermediate orbits to a final geosynchronous station. Power requirements range from less than 75 watts for north-south station keeping on small payloads up to greater than 15 KW for a 180 day orbit raising mission including a 28.5 degree plane change.

Isley, W. C.↗

NASA NTRS1972

Reflectance of evaporated rhenium and tungsten films in the vacuum ultraviolet from 300 to 2000 A.

Discussion of the dependence of the reflectance of Re and W on the substrate temperature during deposition, film thickness, and aging during exposure to air. Re and W of 99.99% purity were evaporated with a 6-kW fine-focused electron gun and deposited on glass and fused quartz plates of various temperatures ranging from 40 to 500 C. With Re, films of highest reflectance were obtained by evaporation onto unheated substrates, whereas with W, heating of the substrate greatly increased the reflectance of the deposited films. For both metals, the reflectance losses during extended exposure to air remained rather small, indicating that the oxide films formed on both film materials at room temperature were very thin.

Cox, J. T.↗

NASA NTRS1972

An out-of-core thermionic-converter system for nuclear space power.

Reexamination of designs of nuclear thermionic space power systems with the converter outside the reactor in the perspective of recent advances in heat-transfer methods, materials, converter performance, and radiation design. The 40- to 70-kW(e) power range is treated. The configuration is found to meet the constraints of readily available launch vehicles. It allows for off-design operation including startup, shutdown, and possible emergency conditions; provides tolerance of failure by extensive use of modular, redundant elements; incorporates and uses heat pipes in a fashion that reduces the need for extensive in-pile testing of system components; and uses thermionic converters, nuclear fuel elements, and heat-transfer devices in a geometrical form adapted from existing incore thermionic system designs.

Breitwieser, R.↗

NASA NTRS1972

Solar radiation calibration facility.

A recently completed calibration facility for the NASA Mississippi Test Site is described. It combines the spectrum of water cooled long arc xenon radiation sources with the integration sphere concept covered by a special barium sulfate diffusely reflecting paint. The sphere is six feet in diameter and completely encompasses the four xenon sources (2.5 kW each) and the five pyranometers under calibration. The sources are contained within the base of the central pedestal and allow the radiation to be emitted into the lower hemisphere. Only through diffuse reflection and multiple scattering is the radiation finally incident upon the test zone, which is in the equatorial plane of the sphere on the top of the pedestal. The sphere is temperature controlled.

Hershey, T. L.↗

NASA NTRS1972

Study of radioisotope safety devices for electric propulsion system, Volume 1: Summary report

A new reference design was prepared for the 5 kW(e) thermionic power supply. The safety equipment in this design is a passive containment system which does not rely on the operation of any mechanisms such as a launch escape rocket or deployment of parachutes. It includes: (1) a blast shield to protect against the explosion of the launch vehicle; (2) a combination of refractory thermal insulation and heat storage material to protect against a sustained launch pad fire; (3) a reentry body with a spherical nose and a large conical flare at the aft end to stabilize the reentry attitude and lower the terminal velocity in air; (4) composite graphite thermal protection to sustain the reentry heat pulse; (5) crushable honeycomb behind the nose to limit the deceleration of the radioisotope source due to impact on land at terminal velocity; (6) a double-walled secondary containment vessel surrounding the isotopic capsules; (7) neutron shielding to reduce external dose rates; (8) an auxiliary cooling system employing redundant heat pipes to remove the radioactive decay heat from the heat source and reject it to the surroundings or to a forced convection loop.

Bradshaw, G. B.↗

NASA NTRS1972

Study to adapt solar electric propulsion to the Pioneer F and G spacecraft

The addition of an electric thrust subsystem to the spin-stabilized Pioneer F and G spacecraft to improve performance capability for certain missions is discussed. The evaluation was performed for the Atlas and Titan launch vehicles with Centaur and TE-364-4 stages and for electric thrust stages of 8- and 5-kw with three 30- and five 15-cm thrusters respectively. The combination of a spinning spacecraft with electric propulsion is a concept only recently evaluated and the penalty from spinning over three-axis stabilized is not as significant as might initally be thought. There are major gains in weight, cost, and reliability, the disadvantages being lower data rate during the thrust phase and less efficient pointing. A variety of missions were evaluated from a solar approach mission into 0.14 AU to a flyby mission of Neptune at approximately 30 AU. Performance improvements were present for all missions evaluated.

Source record↗

NASA NTRS1972

Applications technology satellite advanced missions study, volume 1

Four different spacecraft configurations were developed for geostationary service as a high power communications satellite. The first configuration is a Thor-Delta launch into a low orbit with a spiral ascent to synchronous altitude by ion engine propulsion. The spacecraft is earth oriented with rotating solar arrays. Configuration 2 is a direct injection Atlas/Centaur/Burner II vehicle which when in orbit is sun-oriented with a rotating transponder tower. Configurations 3 and 4 are Titan IIIC launches, and are therefore larger and heavier than Configuration 2. They are both sun-oriented, with rotating transponder towers and are directly injected into orbit. Technology discussed includes high power (up to 2 kW) transmitters with collectors radiating heat directly into space, and contoured antenna patterns designed to illuminate particular earth regions. There is also a review of potential users of the services which can be performed by this type satellite in such areas as information networking, public broadcasting and educational television.

Robinson, D. L.↗

NASA NTRS1972

Laser induced fluorescence in algae: A new technique for remote detection

Measurements of the absorption and fluorescence spectra were obtained for four various types of marine and fresh water algae using a pulsed N2/Ne dye laser as the source of excitation. The absorption maxima for the algae ranged from 420 to 675 nm, while their fluorescent spectra ranged from 580 to 685 nm. It appears feasible that various algal species can be identified by detection of their fluorescent signatures using a tunable laser as the excitation source. However, if one is concerned only with detection of chlorophyll a, the optimum excitation is approximately 600 + 50 nm while detection is at 685 nm. An analysis of both calculations and laboratory results indicates that it should be feasible to measure chlorophyll a in concentrations as low as 1.0 mg/m3 using a 100 kW peak pulsed laser from an altitude of 500 meters.

Friedman, E. J.↗

NASA NTRS1972

Development of RF plasma simulations of in-reactor tests of small models of the nuclear light bulb fuel region

Experiments were conducted to develop test configurations and technology necessary to simulate the thermal environment and fuel region expected to exist in in-reactor tests of small models of nuclear light bulb configurations. Particular emphasis was directed at rf plasma tests of approximately full-scale models of an in-reactor cell suitable for tests in Los Alamos Scientific Laboratory's Nuclear Furnace. The in-reactor tests will involve vortex-stabilized fissioning uranium plasmas of approximately 200-kW power, 500-atm pressure and equivalent black-body radiating temperatures between 3220 and 3510 K.

Roman, W. C.↗

NASA NTRS1972

Nuclear light bulb propellant heating simulation using a tungsten/argon aerosol and radiation from a dc arc surrounded by a segmented mirror cavity

Experiments were conducted to simulate radiant heating of the propellant stream of a nuclear light bulb engine to obtain high bulk exit temperatures in the flowing simulated propellant stream by absorption of large fractions of the incident thermal radiation. A high-power, vortex-stabilized dc arc within an uncooled fused silica tube was used as the radiant energy source. It was surrounded by a mirror system to increase the radiation incident on the simulated propellant. The 12.7-cm-long by 2.3-cm-wide, diverging-duct test section had a transparent front wall and a reflecting rear wall. The central stream of seeded gas, a tungsten-particle/argon aerosol, had unseeded argon buffer layers on both sides to prevent coating of the duct walls. Arc operating times were approximately 0.5 sec with power levels up to 780 kW. Bulk exit temperatures were measured using a calorimeter downstream of the duct. The maximum simulated propellant bulk exit temperature obtained was 4515 K, compared with 3860 K in previous tests, 3300 to 3700 K expected in in-reactor tests in the Nuclear Furnace, and 6660 K in the reference nuclear light bulb engine. The maximum temperature in these tests was limited primarily by the amount of radiation incident on the test section (determined by the arc operating characteristics and the effectiveness of the mirror cavity).

Klein, J. F.↗

NASA NTRS1973

A series-resonant silicon-controlled-rectifier power processor for ion thrusters

A program to develop a power processing system for ion thrusters is presented. Basic operation of the silicon controlled rectifier series inverter circuitry is examined. The approach for synthesizing such circuits into a system which limits the electrical stress levels on the power source, semiconductor switching elements, and the ion thruster load is described. Experimental results are presented for a 2.5-kW breadboard system designed to operate a 20-cm ion thruster.

Shumaker, H. A.↗

NASA NTRS1973

Improvements in and test results for the 2 to 15 kilowatt Brayton cycle electrical subsystem

The electrical subsystem of the 2- to 15-kW Brayton power conversion system consists of the auxiliary electrical equipment required for an integrated, self-contained system. For the last 2 years the electrical subsystem has been undergoing extensive tests. The first year of testing resulted in determining the performance characteristics of the electrical subsystem. During the second year several significant changes and improvements were investigated. An inverter designed for motor starting the alternator performed successfully. Some of the changes that have been made are a new alternator speed pickup, which is independent of the alternator output voltage; new, more efficient power supplies for the control system; and a volts-per-hertz reference for the alternator voltage regulator. Test data were taken on the temperature distribution of the electrical subsystem at startup conditions over a cold-plate temperature range of 25 to -50 C.

Vrancik, J. E.↗