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  • Other Sources  (816)
  • SPACECRAFT PROPULSION AND POWER  (816)
  • 1980-1984  (816)
  • 1
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    Space aging of solid rocket materials (P0005) (1984)
    In:  CASI
    Details
    Publication Date: 2006-02-14
    Description: The objective of this experiment is to determine the effects of long-term orbital exposure on the materials used in solid-rocket space motors. Specifically, structural materials and propellants from the STAR/PAM-D series motors and the PAM DII/IPSM-II motors will be tested, as well as advanced composite case and nozzle materials planned for future use. The experiment approach is to expose samples of solid-rocket propellant, liner, insulation, case, and nozzle specimens to the space environment and to compare preflight and postflight measurements of various mechanical, chemical, and ballistic properties.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Langley Research Center Long Duration Exposure Facility (LDEF); p 94-96
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  • 2
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    The potential impact of new power system technology on the design of a manned Space Station (1984)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Description: Larger, more complex spacecraft of the future such as a manned Space Station will require electric power systems of 100 kW and more, orders of magnitude greater than the present state of the art. Power systems at this level will have a significant impact on the spacecraft design. Historically, long-lived spacecraft have relied on silicon solar cell arrays, a nickel-cadmium storage battery and operation at 28 V dc. These technologies lead to large array areas and heavy batteries for a Space Station application. This, in turn, presents orbit altitude maintenance, attitude control, energy management and launch weight and volume constraints. Size (area) and weight of such a power system can be reduced if new higher efficiency conversion and lighter weight storage technologies are used. Several promising technology options including concentrator solar photovoltaic arrays, solar thermal dynamic and ultimately nuclear dynamic systems to reduce area are discussed. Also, higher energy storage systems such as nickel-hydrogen and the regenerative fuel cell (RFC) and higher voltage power distribution which add system flexibility, simplicity and reduce weight are examined. Emphasis placed on the attributes and development status of emerging technologies that are sufficiently developed so that they could be available for flight use in the early to mid 1990's.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 3
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    Simplification of power electronics for ion thruster neutralizers (1984)
    In:  Other Sources
    Details
    Publication Date: 2011-08-18
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Spacecraft and Rockets (ISSN 0022-4650); 21; 473-480
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  • 4
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    Improved performance design of gallium arsenide solar cells for space (1984)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Description: An improved design, shallow junction heteroface, n-p, gallium arsenide solar cell for space applications is reported, with a predicted AM0 efficiency in the 21.9 to 23.0 percent range. The optimized n-p structure, while slightly more efficient, has the added advantage of being less susceptible to radiation-induced degradation by virtue of this thin top junction layer. Detailed spectral response curves and an analysis of the loss mechanisms are reported. The details of the design are readily measurable. The optimized designs were reached by quantifying the dominant loss mechanisms and then minimizing them by using computer simulations.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: IEEE Transactions on Electron Devices (ISSN 0018-9383); ED-31; 689-695
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  • 5
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    Nuclear-electric power in space (1984)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: Prospective missions requiring large power supplies that might be satisfied with space nuclear reactors (SNR) are discussed, along with design concepts and problems and other potential high-power space systems. Having a minimum economic output of 10 kWe, SNR seem well-suited as the power sources for DBS systems, space-based ATC systems manned planetary missions, an expanding Space Station, materials processing, and outer planets missions. SNR avoid the large area problems of solar cell arrays, short lifetimes of thermionic converters, and vibration and heat control in Stirling engines. Design problems exist for SNR in the heat transfer and rejection systems, radioactive emissions and degradation of reactor materials, and size. The latter is a function of Shuttle payload constaints and raises the possibility of having to load the fuel while in orbit. The earliest operational date of SNRs is projected for the early 1990s, if progress is good in the current SP-100 program.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: IEEE Spectrum (ISSN 0018-9235); 21; 58-65
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  • 6
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    100 kWe magnetoplasmadynamic thruster system design (1984)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Spacecraft and Rockets (ISSN 0022-4650); 21; 563-572
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  • 7
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    J-series ion thruster isolator failure analysis (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Spacecraft and Rockets (ISSN 0022-4650); 21; 321
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  • 8
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    Numerical modeling of laser thermal propulsion flows (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: An review of the problems associated with modeling laser thermal propulsion flows, a synopsis of the status of such models, and the attributes of a successful model are presented. The continuous gaseous hydrogen laser-supported combustion wave (LSCW) thruster, for which a high-energy laser system (preferably space-based) should exist by the time the propulsion technology is developed, is considered in particular. The model proposed by Raizer (1970) is based on the assumptions of one-dimensional flow at constant pressure with heat conduction as the principal heat transfer mechanism. Consideration is given to subsequent models which account for radiative transfer into the ambient gas; provide a two-dimensional generalization of Raizer's analysis for the subsonic propagation of laser sparks in air; include the effect of forward plasma radiation in a one-dimensional model; and attempt a time-dependent (elliptic) solution of the full Navier-Stokes equations for the flow in a simple axisymmetric thruster. Attention is also given to thruster and nozzle flow models and thermodynamic and transport properties.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 9
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    Surface chloride salt formation on Space Shuttle exhaust alumina (1984)
    In:  Other Sources
    Details
    Publication Date: 2011-08-18
    Description: Aluminum oxide samples from the exhaust of Space Shuttle launches STS-1, STS-4, STS-5, and STS-6 were collected from surfaces on or around the launch pad complex and chemically analyzed. The results indicate that the particulate solid-propellant rocket motor (SRM) alumina was heavily chlorided. Concentrations of water-soluble aluminum (III) ion were large, suggesting that the surface of the SRM alumina particles was rendered soluble by prior reactions with HCl and H2O in the SRM exhaust cloud. These results suggest that Space Shuttle exhaust alumina particles are good sites for nucleation and condensation of atmospheric water. Laboratory experiments conducted at 220 C suggest that partial surface chloriding of alumina may occur in hot Space Shuttle exhaust plumes.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Geophysical Research (ISSN 0148-0227); 89; 2535-254
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  • 10
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    Comparison of potential electric propulsion systems for orbit transfer (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Previously cited in issue 03, p. 283, Accession no. A83-14375
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Spacecraft and Rockets (ISSN 0022-4650); 21; 88-95
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  • 11
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    Advanced OTV engine concepts (1984)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The results and status of engine technology efforts to date and related company funded activities are presented. Advanced concepts in combustors and injectors, high speed turbomachinery, controls, and high-area-ratio nozzles that package within a short length result is engines with specific impulse values 35 to 46 seconds higher than those now realized by operational systems. The improvement in life, reliability, and maintainability of OTV engines are important.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Research Center OTV Propulsion Issues; p 135-148
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  • 12
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    Advanced OTV engines and issues (1984)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The propulsion system requirements of orbit transfer vehicles (OTV). A baseline expander cycle engine which will meet those requirements was defined. The principal characteristics of a baseline engine and some options which are available to accommodate OTV system optimization studies was discussed. Engine program issues which are dependent on the mission scenario and the vehicle system configuration are shown. The rationale for a new cryogenic OTV engine is summarized.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Research Center OTV Propulsion Issues; p 127-134
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  • 13
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    Aerojet advanced engine concept (1984)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The future orbit transfer vehicle (OTV) requirements which dictate the need for a highly versatile, highly reliable, reusable propulsion module are discussed. To attain maximum operational economy, space-basing is essential. This requires a reusable, maintenance free engine. The design features of this space based engine are defined. A new engine cycle and its advantages allow all the maintenance goals to be attained. Rubbing contact and interpropellant seals and purges are eliminated when GO2 is used to drive the LO2 pump. The TPA design has only one moving part. The use of both GH2 and GO2 to drive the turbines lowers the turbine temperatures in addition lower GH2 temperatures and pressures improve chamber cooling and longer life. The use of GO2 as a turbine drive fluid is addressed. Space based engines require an integrated control and health monitoring system to improve system reliability and eliminate all scheduled maintenance. It is concluded that all OTV propulsion requirements can be fulfilled with a single engine. The technological developments required to demonstrate that engine are outlined.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Research Center OTV Propulsion Issues; p 113-125
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  • 14
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    Developments in space power components for power management and distribution (1984)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Advanced power electronic components development for space applications is discussed. The components described include transformers, inductors, semiconductor devices such as transistors and diodes, remote power controllers, and transmission lines.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Langley Research Center An Assessment of Integrated Flywheel System Technol.; p 369-387
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  • 15
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    Magnetically suspended flywheel system study (1984)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: A program to study the application of a graphite/epoxy, magnetically suspended, pierced disk flywheel for the combined function of spacecraft attitude control and energy storage (ACES) is described. Past achievements of the program include design and analysis computer codes for the flywheel rotor, a magnetically suspended flywheel model, and graphite/epoxy rotor rings that were successfully prestressed via interference assembly. All hardware successfully demonstrated operation of the necessary subsystems which form a complete ACES design. Areas of future work include additional rotor design research, system definition and control strategies, prototype development, and design/construction of a UM/GSFC spin test facility. The results of applying design and analysis computer codes to a magnetically suspended interference assembled rotor show specific energy densities of 42 Wh/lb (92.4 Wh/kg) are obtained for a 1.6 kWh system.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Langley Research Center An Assessment of Integrated Flywheel System Technol.; p 307-328
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  • 16
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    Overview of magnetic bearing control and linearization approaches for annular magnetically suspended devices (1984)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: An overview of magnetic bearing control and linearization approaches which have been considered for annular magnetically suspended devices is presented. These devices include the Annular Momentum Control Device and the Annular Suspension and Pointing System. Two approaches were investigated for controlling the magnetic actuator. One approach involves controlling the upper and lower electromagnets differentially about a bias flux. The bias flux can either be supplied by permanent magnets in the magnetic circuit or by bias currents. In the other approach, either the upper electromagnet or the lower electromagnet is controlled depending on the direction of force required. One advantage of the bias flux is that for small gap perturbations about a fixed operating point, the force-current characteristic is linear. Linearization approaches investigated for individual element control include an analog solution of the nonlinear electromagnet force equation and a microprocessor-based table lookup method.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: An Assessment of Integrated Flywheel System Technol.; p 297-306
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  • 17
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    Magnetic suspension options for spacecraft inertia-wheel applications (1984)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Design criteria for spacecraft inertia-wheel suspensions are listed. The advantages of magnetic suspensions over other suspension types for spacecraft inertia-wheel applications are cited along with the functions performed by magnetic suspension. The common designs for magnetic suspensions are enumerated. Materials selection of permanent magnets and core materials is considered.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Langley Research Center An Assessment of Integrated Flywheel System Technol.; p 281-296
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  • 18
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    Description of a laboratory model Annular Momentum Control Device (AMCD) (1984)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The basic concept of the Annular Momentum Control Device (AMCD) is that of a rotating annular rim suspended by noncontacting magnetic bearings and driven by a noncontacting electromagnetic spin motor. The purpose of this paper is to highlight some of the design requirements for AMCD's in general and describe how these requirements were met in the implementation of laboratory test model AMCD. An AMCD background summary is presented.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: An Assessment of Integrated Flywheel System Technol.; p 157-167
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  • 19
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    Perspectives on energy storage wheels for space station application (1984)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Several of the issues of the workshop are addressed from the perspective of a potential Space Station developer and energy wheel user. Systems' considerations are emphasized rather than component technology. The potential of energy storage wheel (ESW) concept is discussed. The current status of the technology base is described. Justification for advanced technology development is also discussed. The study concludes that energy storage in wheels is an attractive concept for immediate technology development and future Space Station application.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Langley Research Center An Assessment of Integrated Flywheel System Technol.; p 117-127
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  • 20
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    Inertial storage for satellites (1984)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: A new system is being developed that performs satellite attitude control, attitude reference, and energy storage utilizing inertia wheels. The baseline approach consists of two counter rotating flywheels suspended in specially designed magnetic bearings, spin axis motor/generators, and a control system. The control system regulates the magnetic bearings and spin axis motor/generators and interacts with other satellite subsystems (photovoltaic array, star trackers, Sun sensors, magnetic torquers, etc.) to perform the three functions. Existing satellites utilize separate subsystems to perform attitude control, provide attitude reference, and store energy. These functions are currently performed using reaction or momentum wheels, gyros, batteries, and devices that provide an absolute reference (Sun sensors and star trackers). A Combined Attitude, Reference, and Energy Storage (CARES) system based on high energy density inertial energy storage wheels (flywheels) has potential advantages over existing technologies. Even when used only for energy storage, this system offers the potential for substantial improvements in life, energy efficiency, and weight over existing battery technologies. Utilizing this same device for both attitude control and attitude reference would result in significant additional savings in overall satellite weight and complexity.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Langley Research Center An Assessment of Integrated Flywheel System Technol.; p 101-116
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  • 21
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    Space station power system (1984)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: It is pointed out that space station planning at NASA began when NASA was created in 1958. However, the initiation of the program for a lunar landing delayed the implementation of plans for a space station. The utility of a space station was finally demonstrated with Skylab, which was launched in 1972. In May 1982, the Space Station Task Force was established to provide focus and direction for space station planning activities. The present paper provides a description of the planning activities, giving particular attention to the power system. The initial space station will be required to supply 75 kW of continuous electrical power, 60 kW for the customer and 15 kW for space station needs. Possible alternative energy sources for the space station include solar planar or concentrator arrays of either silicon or gallium arsenide.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: IEEE Transactions on Aerospace and Electronic Systems (ISSN 0018-9251); AES-20; 666-671
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  • 22
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    High-temperature gaseous oxygen/hydrogen thrusters for space station (1984)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: Efforts to determine the critical technologies necessary for the development of a high-temperature GO2/GH2 thruster are reviewed, for space station applications. Two types of thrust chambers are evaluated: one operable at high temperature and the other incorporating regenerative cooling. The high temperature chamber made of rhenium requires minimum cooling of the chamber wall, however, an oxidation barrier should be incorporated to prevent the rhenium thruster from readily oxidizing. The use of rhenium brought about a lower cost, lower weight, and simplicity of fabrication. The igniter-injector, the chamber, the test facility, and the test program are discussed. It is found that an obtained flow split of 58/42 increased the thruster performance from 3893 N-s/Kg to 4030 N-s/Kg for the same chamber pressure and overall mixture ratio. An increase in performance is also observed when the core mixture ratio is lowered. The thruster was tested for 1.9 hrs with no well degradation.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 23
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    Space Shuttle Orbiter improved auxiliary power unit (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: The Space Shuttle Orbiter Auxiliary Power Unit subsystem has operated successfully on three vehicles by meeting mission requirements and has proven the design for space operation. The current Auxiliary Power Unit (APU) operational life is limited to 12 missions and the APU turnaround between flights is longer than originally anticipated. The Improved APU objective is to increase life to 50 missions, reduce the three - APU subsystem vehicle weight by 140 lbs., and reduce turnaround time. The design changes incorporated into the Improved APU and the associated development testing are described.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: APL The 1984 JANNAF Propulsion Meeting, Vol. 2; p 315-326
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  • 24
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    Improved ion containment using a ring-cusp ion thruster (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Spacecraft and Rockets (ISSN 0022-4650); 21; 488-495
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  • 25
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    Integration becomes the name of the OTV game (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Since 1981, NASA has been supporting an Orbit Transfer Vehicle (OTV) propulsion technology development program which encompasses the efforts of three major engine manufacturers. A thrust variability ratio of 30:1 has been stipulated for the baseline 520 lbf-sec/lbm engines, in order to yield the versatility needed for low acceleration missions, orbit-transfer missions, and aeromaneuvering tasks. These goals may not be reachable either individually or collectively; the program supports generic technology benefitting all three engine concepts, as well as concept-specific technology. Engine/vehicle integration will determine the final configuration. The three engines under study differ as to turbomachinery types and operating speeds.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Aerospace America (ISSN 0740-722X); 22; 70-73
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  • 26
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    Hydrogen chloride and aerosol ground cloud characteristics resulting from Space Shuttle launches (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Airborne measurements of gaseous HCl, gaseous and aerosol HCl, particulates, relative humidity and temperature were obtained in ground clouds produced during three Space Shuttle launches. Partitioning of HCl between HCl aerosol and gaseous HCl was investigated as the solid rocket exhaust cloud diluted with ambient air to evaluate the conditions under which aerosol formation occurs in the troposphere in the presence of hygroscopic HCl vapor. Equilibrium predictions for aqueous HCl aerosol formation generally agree with the measured HCl partitioning over HCl concentrations from 0.5 to 36 ppm. HCl concentration dispersion within four cloud segments at time t (min) was evaluated using the expression C = C(0) (t to the alpha power) where C(0) varied from 145 to 2250 ppm and alpha varied from -1.14 to -1.73. Aerosol fallout from the exhaust clouds was measured with time by monitoring HCl concentrations and aerosol distributions 100 m below the cloud as it drifted away from the launch site. Significant amounts of HCl were found to be removed by fallout of particles in the 80-220 micron diameter range up to 30 min after launch.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Atmospheric Environment (ISSN 0004-6981); 18; 4, 19
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  • 27
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    Design, development, performance, and reconditioning of Ni-Cd batteries using polyropylene separators (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: A performance assessment is made for the Viking Mars Lander Program's sealed, sterilizable, 8-ampere hour Ni-Cd batteries, which use a nonwoven polypropylene separator material. Attention is given to separator wettability, the optimization of electrolyte quantity, and the reduction of plate carbonate, in view of thermal considerations and other environmental design and test requirements generated by mission characteristics. Life data based on mission experience identify, in addition to performance and degradation behavior, a series of shallow discharge reconditioning cycles and an intensive program of deep discharge reconditioning.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Power Sources (ISSN 0378-7753); 12; 305-316
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  • 28
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    LeRC rail accelerators - Test designs and diagnostic techniques (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: The feasibility of using rail accelerators for various in-space and to-space propulsion applications was investigated. A 1 meter, 24 sq mm bore accelerator was designed with the goal of demonstrating projectile velocities of 15 km/sec using a peak current of 200 kA. A second rail accelerator, 1 meter long with a 156.25 sq mm bore, was designed with clear polycarbonate sidewalls to permit visual observation of the plasma arc. A study of available diagnostic techniques and their application to the rail accelerator is presented. Specific topics of discussion include the use of interferometry and spectroscopy to examine the plasma armature as well as the use of optical sensors to measure rail displacement during acceleration. Standard diagnostics such as current and voltage measurements are also discussed. Previously announced in STAR as N83-35053
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: IEEE Transactions on Magnetics (ISSN 0018-9464); MAG-20; 324-327
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  • 29
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    Nuclear electric propulsion (NEP) spacecraft configuration study (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: The NEP spacecraft design is for the 100 kWe SP-100 nuclear reactor thermoelectric power supply. Thermoelectric conversion, one of several conversion options being considered for the SP-100, is discussed. Whereas many of the earlier NEP spacecraft designs have been based on highly integrated power and thruster subsystems and have resulted in subsystem requirements that depended largely on the design mission, the SP-100 nuclear power system is a general-purpose power supply. The configuration design must be compatible with this power subsystem design philosophy. The study also detects power and propulsion subsystem integration issues and related technology requirements. It is noted that the new NEP spacecraft design satisfies the most important subsystem integration requirements at the expense of some increase in vehicle inert mass relative to that used in previous studies.
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  • 30
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    Deflagration plasma thruster (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: This paper introduces the application of the magnetized plasma deflagration process to space propulsion. The deflagration process has the unique capability of efficiently converting input energy into kinetic energy in the accelerating direction. To illustrate the totally divergent characters of 'snowplow' detonation and deflagration discharges, examples of the differences between deflagration and detonation 'snowplow' discharges are expressed in terms of current densities, temperature, and particle velocities. Magnetic field profiles of the deflagration mode of discharges are measured. Typical attainable plasma characteristics are described in terms of velocity, electron temperature, and density, as well as measurement techniques. Specific impulses measured by piezo-electric probe and pendulum methods are presented. The influence of the transmission line in the discharge circuits on plasma velocity is measured by means of a microwave time-of-flight method. The results for the deflagration thruster are compared with other space thrusters. Further research areas are identified.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 31
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    Equivalent electron fluence for space qualification of shallow junction heteroface GaAs solar cells (1984)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Description: It is desirable to perform qualification tests prior to deployment of solar cells in space power applications. Such test procedures are complicated by the complex mixture of differing radiation components in space which are difficult to simulate in ground test facilities. Although it has been shown that an equivalent electron fluence ratio cannot be uniquely defined for monoenergetic proton exposure of GaAs shallow junction cells, an equivalent electron fluence test can be defined for common spectral components of protons found in space. Equivalent electron fluence levels for the geosynchronous environment are presented.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: IEEE Transactions on Electron Devices (ISSN 0018-9383); ED-31; 622-625
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    Thrust for interorbital propulsion - A question of lifetime (1984)
    In:  Other Sources
    Details
    Publication Date: 2011-08-18
    Description: The most promising operational window for the use of magnetoplasmadynamic (MPD) thrusters is identified to be at megawatt power levels for orbital maneuvering. For such applications, the operation of a steady-state MPD thruster system imposes stringent requirements on the lifetime of thruster surfaces which will erode due to interactions with the plasma working fluid. Basic erosion mechanisms are presented and the problems associated with measuring the erosion rates are discussed. An experimental approach that would allow the development of a phenomenological model for erosion is proposed.
    Keywords: SPACECRAFT PROPULSION AND POWER
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    Propulsion options for orbit-on-demand vehicles (1984)
    In:  Other Sources
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    Publication Date: 2018-12-01
    Description: Vehicles which can achieve orbit within minutes of a command decision may be needed in the future for a variety of missions. Such orbit-on-demand vehicles may have propulsion requirements that are somewhat different from vehicles designed for routine transportation, but the propulsion evaluation studies of the past need to be considered as a starting point for orbit-on-demand vehicle studies. This paper surveys airbreathing propulsion studies including composite, airturbo-rocket, and scramjet systems and rocket propulsion studies including composite, airturborocket, and scramjet and rocket propulsion studies including dual-fuel and pure hydrocarbon systems. One indication from the results is that a horizontal takeoff airbreathing system with supersonic staging will have a higher development cost than rocket systems primarily because of the cost of the airbreathing engine development. Another indication is that pure hydrocarbon rocket propulsion for a vertical takeoff system may be feasible. Eliminating the requirement for hydrogen fuel may be worthwhile for orbit-on-demand vehicles.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 34
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    Orbital transfer vehicle engine integration study (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: NASA-LeRC is sponsoring industry studies to establish the technology base for an advanced engine for orbital transfer vehicles for mid-1990s IOC. Engine contractors are being assisted by vehicle contractors to define the requirements, interface conditions, and operational design criteria for new LO2-LH2 propulsion systems applicable to future orbit transfer vehicles and to assess the impacts on space basing, man rating, and low-G transfer missions on propulsion system design requirements. The results of a study is presented. The primary study emphasis was to determine what the OTV engine thrust level should be, how many engines are required on the OTV, and how the OTV engine should be designed. This was accomplished by evaluating planned OTV missions and concepts to determine the requirements for the OTV propulsion system, conducting tradeoffs and comparisons to optimize OTV capability, and evaluating reliability and maintenance to determine the recommended OTV engine design for future development.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-174842 , NAS 1.26:174842 , GDC-SP-84-050
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  • 35
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    Dynamic Conversion Working Group Summary (1984)
    In:  CASI
    Details
    Publication Date: 2014-09-13
    Description: The potential of dynamic conversion devices for use in solar and nuclear dynamic space power systems was addressed. Conversion systems considered were based on the use of Brayton, Stirling and Rankine cycles. Both organic and liquid metal Rankine cycles were included. The basic system considerations were: mission requirements, system attributes, system options, technology issues and constraints, and priorities of needed technology development. Mission requirements, where dynamic conversion was considered enabling technology, were identified along with the associated power levels and potential energy sources. When considering the system options special attention was given to recommend operating temperatures and other significant discriminators. A list of prioritized tasks considered important for the successful development of dynamic conversion systems for 1995 and beyond was compiled.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Research Center Space Power; p 297-299
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  • 36
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    Advanced and Nontraditional Concepts Working Group Report (1984)
    In:  CASI
    Details
    Publication Date: 2014-09-13
    Description: Advanced and nontraditional concepts relating to future space power requirements were examined with special emphasis on the requirements for the space station. Key findings in the following ares are outlined: dynamic radiation concepts, space nuclear reactors, energy conversion cycles, beam power transmission, electrodynamic tethers and advanced photovoltaics.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Research Center Space Power; p 327-335
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    Electrical System Technology Working Group (WG) Report (1984)
    In:  CASI
    Details
    Publication Date: 2014-09-13
    Description: The technology needs for space power systems (military, public, commercial) were assessed for the period 1995 to 2005 in the area of power management and distribution, components, circuits, subsystems, controls and autonomy, modeling and simulation. There was general agreement that the military requirements for pulse power would be the dominant factor in the growth of power systems. However, the growth of conventional power to the 100 to 250kw range would be in the public sector, with low Earth orbit needs being the driver toward large 100kw systems. An overall philosophy for large power system development is also described.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Research Center Space Power; p 317-321
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  • 38
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    Electrodynamic Tethers. 1: Power Generator in LEO. 2: Thrust for Propulsion and Power Storage (1984)
    In:  CASI
    Details
    Publication Date: 2014-09-13
    Description: An electrodynamic tether consists of a long insulated wire in space whose orbital motion cuts across lines of magnetic flux to produce an induce voltage that in typical low orbits averages about 200 v/km. Such a system should be capable of generating substantial electrical power, at the expense of IXB drag acting on its orbital energy. If a reverse current is driven against the induced voltage, the system should act as a motor producing IXB thrust. A reference system was designed, capable of generating 20 KW of power into an electrical load located anywhere along the wire at the expense of 2.6N (20,000 J/sec) drag on the wire. In an ideal system, the conversion between mechanical and electrical energy would reach 100% efficiency. In the actual system part of the 20 KW is lost to internal resistance of the wire, plasma and ionosphere, while the drag force is increased by residual air drag. The 20 KW PMG system as designed is estimated to provide 18.7 KW net power to the load at total drag loss of 20.4 KJ/sec, or an overall efficiency of 92%. Similar systems using heavier wire appear capable of producing power levels in excess of 1 Megawatt at voltages of 2-4 KV, with conversion efficiency between mechanical and electrical power better than 95%. The hollow cathode based system should be readily reversible from generator to motor operation by driving a reverse current using onboard power.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Research Center Space Power; p 275-284
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  • 39
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    Heat Rejection Working Group Synopsis (1984)
    In:  CASI
    Details
    Publication Date: 2014-09-13
    Description: Heat rejection system requirements of specific mission types (space station, planetary exploration, commercial, very high power, and military missions) are discussed. Heat pipe radiators, weight and volume reduction, stable coatings, and working fluids are addressed.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Research Center Space Power; p 309-316
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  • 40
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    The Liquid Droplet Radiator - an Ultralightweight Heat Rejection System for Efficient Energy Conversion in Space (1984)
    In:  CASI
    Details
    Publication Date: 2014-09-13
    Description: A heat rejection system for space is described which uses a recirculating free stream of liquid droplets in place of a solid surface to radiate waste heat. By using sufficiently small droplets ( 100 micron diameter) of low vapor pressure liquids the radiating droplet sheet can be made many times lighter than the lightest solid surface radiators (heat pipes). The liquid droplet radiator (LDR) is less vulnerable to damage by micrometeoroids than solid surface radiators, and may be transported into space far more efficiently. Analyses are presented of LDR applications in thermal and photovoltaic energy conversion which indicate that fluid handling components (droplet generator, droplet collector, heat exchanger, and pump) may comprise most of the radiator system mass. Even the unoptimized models employed yield LDR system masses less than heat pipe radiator system masses, and significant improvement is expected using design approaches that incorporate fluid handling components more efficiently. Technical problems (e.g., spacecraft contamination and electrostatic deflection of droplets) unique to this method of heat rejectioon are discussed and solutions are suggested.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Research Center Space Power; p 261-274
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    Effects of the Solar and Extraterrestrial Environments on Space Power Systems (1984)
    In:  CASI
    Details
    Publication Date: 2014-09-12
    Description: Environments surrounding the major extraterrestrial bodies in the solar system and their interactions with spacecraft power systems are summarized. The environments associated with neutrals/dust, low energy plasma, and where applicable, magnetospheres are discussed for a wide variety of cases. The impact of these environments on power systems - in particular, radiation effects, spacecraft charging, plasma interactions, surface sputtering/erosion, and induced currents - are presented. As power systems must be designed to survive in these hostile environments, it is important that they be taken into account in planning future power systems.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Research Center Space Power; p 225-249
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  • 42
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    Analytical Modeling of Spacecraft Power Systems (1984)
    In:  CASI
    Details
    Publication Date: 2014-09-12
    Description: During the 25 years of space flight with unmanned Earth orbiting satellites, there has been an evolution of power systems in three general areas. The size of power system in terms of power demand at the bus has increased frorm a few watts in the early 1960s to a few hundred watts during the 1970s. Today, the bus power requirements are typically in the .5 to 1 kw range with some mission requirements exceeding the 1 kw size. Power grounding and user isolation are major design considerations when high fidelity power is required by a user.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Research Center Space Power; p 219-223
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    Space Power Management and Distribution Status and Trends (1984)
    In:  CASI
    Details
    Publication Date: 2014-09-12
    Description: An overview of space power management and distribution (PMAD) is provided which encompasses historical and current technology trends. The PMAD components discussed include power source control, energy storage control, and load power processing electronic equipment. The status of distribution equipment comprised of rotary joints and power switchgear is evaluated based on power level trends in the public, military, and commercial sectors. Component level technology thrusts, as driven by perceived system level trends, are compared to technology status of piece-parts such as power semiconductors, capacitors, and magnetics to determine critical barriers.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Research Center Space Power; p 205-218
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  • 44
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    Power for Commercial Space Applications (1984)
    In:  CASI
    Details
    Publication Date: 2014-09-12
    Description: Since the proposed space station is intended to be a permanent installation, used in part by commercial organizations, its design requirements differ fundamentally from those in previous manned spacecraft. In particular, commercialization on a significant scale depends critically on the ability to control capital and operational costs, including the cost of energy, and this demands new approaches to systems such as the power supply for the space station. These considerations suggest guidelines for power plant development. The cost of energy in space, guidelines for power supply development, quality standards, and the role of government are discussed.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Research Center Space Power; p 37-44
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  • 45
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    Space Power for Communication Satellites Beyond 1995 (1984)
    In:  CASI
    Details
    Publication Date: 2014-09-12
    Description: The space power trends for communication satellites beginning in the mid-70's are reviewed. Predictions of technology advancements and requirements were compared with actual growth patterns. The conclusions derived suggest that the spacecraft power system technology base and present rate of advancement will not be able to meet the power demands of the early to mid-90's. It is recommended that an emphasis on accelerating the technology development be made to minimize the technology gap.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Research Center Space Power; p 31-36
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  • 46
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    Emerging Space Nuclear Power Needs (1984)
    In:  CASI
    Details
    Publication Date: 2014-09-12
    Description: Growing interest in new classes of military and civil space systems which demand substantial increases in power over current satellites is generating a renewed interest in space qualified nuclear power systems. Indeed, one can say that power is a limiting technology to the achievement of many future goals in space. The speed of nuclear power system development is currently limited by the lack of a clear distinct definition of system requirements. Emerging system requirements are discussed for the following fields: robust surveillance systems, survivable communication systems with anti-jam capabilities, electric propulsion systems, and weapons applications.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Research Center Space power; p 27-30
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    Advanced Concepts Liquid Droplet Radiator (1984)
    In:  CASI
    Details
    Publication Date: 2014-09-13
    Description: An exploratory point design study was carried out on a shuttle-launchable megawatt nuclear orbit transfer vehicle (OTV) with a 5000 kg payload capacity. The system, which consists of a fixed bed reactor, a Brayton cycle power conversion system, and a liquid droplet radiator (LDR) to reject heat, is deployable from a small package. The methods and technologies of this design are discussed, as well as critical design problems. While this is a preliminary study, it indicates that a space-nuclear reactor, combined with the LDR, make possible a continuous 10 MW(e) power station on orbit with a single shuttle launch.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Research Center Space Power; p 251-260
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  • 48
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    Atomization data requirements for rocket combustor modeling (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The complex computer codes, which model liquid rocket combustors, require information about the distribution and atomization of these liquid reactants. The available information is, in general, of questionable validity and applicability. Authors and users of combustion codes are often unaware of, or underestimate the importance of, these deficiencies in atomization data. These deficiencies and their importance are examined. Results of analyses performed with a state-of-the-art rocket combustion code are presented which demonstrate the important effects of such atomization information as initial droplet sizes and size distribution on vaporization rate and losses. Also, the questionable aspects and inapplicability of the available atomization data are discussed. One important and often neglected or misunderstood aspect of atomization data is the differences between spatial (concentration) and flux (often called temporal) droplet size distributions. These are described, and a computer model constructed to assess the difference between concentration and flux droplet size distributions is described and results presented. Experimental data are also given to demonstrate this difference. Finally, experimental results are presented that demonstrate the very great, and often neglected effect, of the local gas velocity field on atomization.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: APL 21st JANNAF Combust. Meeting, Vol. 1; p 369-377
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  • 49
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    Ring-cusp ion thruster with shell anode (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An improved ion thruster for low specific impulse operation in the 1500 sec to 6000 sec range has a multicusp boundary field provided by high strength magnets on an iron anode shell which lengthens the paths of electrons from a hollow cathode assembly. A downstream anode pole piece in the form of an iron ring supports a ring of magnets to provide a more uniform beam profile. A cylindrical cathode magnet can be moved selectively in an axial direction along a feed tube to produce the desired magnetic field at the cathode tip.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NAS 1.71:LEW-13881-1
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    Analysis of large solid propellant rocket engine exhaust plumes using the direct simulation Monte Carlo method (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A new solution procedure has been developed to analyze the flowfield properties in the vicinity of the Inertial Upper Stage/Spacecraft during the 1st stage (SRMI) burn. Continuum methods are used to compute the nozzle flow and the exhaust plume flowfield as far as the boundary where the breakdown of translational equilibrium leaves these methods invalid. The Direct Simulation Monte Carlo (DSMC) method is applied everywhere beyond this breakdown boundary. The flowfield distributions of density, velocity, temperature, relative abundance, surface flux density, and pressure are discussed for each species for 2 sets of boundary conditions: vacuum and freestream. The interaction of the exhaust plume and the freestream with the spacecraft and the 2-stream direct interaction are discussed. The results show that the low density, high velocity, counter flowing free-stream substantially modifies the flowfield properties and the flux density incident on the spacecraft. A freestream bow shock is observed in the data, located forward of the high density region of the exhaust plume into which the freestream gas does not penetrate. The total flux density incident on the spacecraft, integrated over the SRM1 burn interval is estimated to be of the order of 10 to the 22nd per sq m (about 1000 atomic layers).
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 84-0496
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    Aluminum chloride formation on Space Shuttle aluminum oxide (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Samples of particulates collected from the exhaust of Space Shuttle launches STS-1, -4, -5, -6, and -7 were analyzed. Scanning electron microscopy and X-ray diffractometry of these samples indicated that the particulates were spherical and predominantly composed of aluminum oxide. The water-soluble weight fraction, pH, acid-soluble weight fraction, and insoluble weight fraction were determined for each sample. Water-soluble weight fractions averaged about 7 percent of the total sample weight, were generally very acidic, and contained significantly elevated concentrations of chloride and aluminum (III) ion. The high concentrations of soluble aluminum (III) and chloride ions observed suggested that aluminum chlorides and/or oxychlorides had formed on the surface of the alumina particulates. More than 72 percent by weight of each sample was insoluble in either water or strong mineral acid, and was identified as alpha-Al2O3. The results from these analyses suggest that the surface of Space Shuttle exhaust alumina particulates will be highly acidic and heavily chlorided, and that a substantial amount of the surface chloride may be chemically associated with aluminum (III) ions rather than just physically adsorbed as HCl.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 84-0469
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    A comparison of the efficiency of numerical methods for integrating chemical kinetic rate equations (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The efficiency of several algorithms used for numerical integration of stiff ordinary differential equations was compared. The methods examined included two general purpose codes EPISODE and LSODE and three codes (CHEMEQ, CREK1D and GCKP84) developed specifically to integrate chemical kinetic rate equations. The codes were applied to two test problems drawn from combustion kinetics. The comparisons show that LSODE is the fastest code available for the integration of combustion kinetic rate equations. It is shown that an iterative solution of the algebraic energy conservation equation to compute the temperature can be more efficient then evaluating the temperature by integrating its time-derivative.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: APL Computational Methods. 1984 JPM Spec. Session; p 69-82
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  • 53
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    The annular flow electrothermal ramjet (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The annular flow, electrothermal, plug ramjet is examined as a possible means of achieving rapid projectile acceleration to velocities for such applications as direct launch of spacebound payloads. The performance of this ramjet operating with hydrogen propellant is examined for cases where this working fluid is treated: (1) as a perfect gas, and (2) as a gas that is allowed to dissociate and ionize and then recombine with finite reaction rates in the nozzle. Performance results for these cases are compared to the performance of a conventional ramjet operating with perfect gas hydrogen propellant. The performance of the conventional ramjet is superior to that of the annular flow, electrothermal ramjet. However, it is argued that the mechanical complexities associated with conventional ramjet operation are difficult to attain, and for this reason the annular flow, electrothermal ramjet is more desirable as a launch system. Models are presented which describe both electrothermal plug ramjet and conventional ramjet operation, and it is shown that for a given flight velocity there is a rate of heat addition per unit propellant mass for which ramjet operation is optimized.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-174704 , NAS 1.26:174704
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    Ion beamlet steering for two-grid electrostatic thrusters (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental study of ion beamlet steering in which the direction of beamlets emitted from a two grid aperture system is controlled by relative translation of the grids, is described. The results can be used to design electrostatic accelerating devices for which the direction and focus of emerging beamlets are important. Deflection and divergence angle data are presented for two grid systems as a function of the relative lateral displacement of the holes in these grids. At large displacements, accelerator grid impingements become excessive and this determines the maximum allowable displacement and as a result the useful range of beamlet deflection. Beamlet deflection is shown to vary linearly with grid offset angle over this range. The divergence of the beamlets is found to be unaffected by deflection over the useful range of beamlet deflection. The grids of a typical dished grid ion thruster are examined to determine the effects of thermally induced grid distortion and prescribed offsets of grid hole centerlines on the characteristics of the emerging beamlets. The results are used to determine the region on the grid surface where ion beamlet deflections exceed the useful range. Over this region high accelerator grid impingement currents and rapid grid erosion are predicted.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-174671 , NAS 1.26:174671
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    Preliminary engineering report for design of a subscale ejector/diffuser system for high expansion ratio space engine testing (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The design of a subscale jet engine driven ejector/diffuser system is examined. Analytical results and preliminary design drawings and plans are included. Previously developed performance prediction techniques are verified. A safety analysis is performed to determine the mechanism for detonation suppression.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-171137 , NAS 1.26:171137 , LMSC-HREC-TR-D951414
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    NASA Orbit Transfer Rocket Engine Technology Program (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The advanced expander cycle engine with a 15,000 lb thrust level and a 6:1 mixture ratio and optimized performance was used as the baseline for a design study of the hydrogen/oxgyen propulsion system for the orbit transfer vehicle. The critical components of this engine are the thrust chamber, the turbomachinery, the extendible nozzle system, and the engine throttling system. Turbomachinery technology is examined for gears, bearing, seals, and rapid solidification rate turbopump shafts. Continuous throttling concepts are discussed. Components of the OTV engine described include the thrust chamber/nozzle assembly design, nozzles, the hydrogen regenerator, the gaseous oxygen heat exchanger, turbopumps, and the engine control valves.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-168156 , NAS 1.26:168156
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    Analysis of SSME HPOTP rotordynamics subsynchronous whirl (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The causes and remedies of vibration and subsynchronous whirl problems encountered in the Shuttle Main Engine SSME turbomachinery are analyzed. Because the nonlinear and linearized models of the turbopumps play such an important role in the analysis process, the main emphasis is concentrated on the verification and improvement of these tools. It has been the goal of our work to validate the equations of motion used in the models are validated, including the assumptions upon which they are based. Verification of th SSME rotordynamics simulation and the developed enhancements, are emphasized.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-171094 , NAS 1.26:171094 , REPT-0340284FR-NA17-SSM04
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    Singlet oxygen generator for a solar powered chemically pumped iodine laser (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The potential of solid phase endoperoxides as a means to produce single-delta oxygen in the gas phase in concentrations useful to chemical oxygen-iodine lasers was investigated. The 1,4 - endoperoxide of ethyl 3- (4-methyl - 1-naphthyl) propanoate was deposited over an indium-oxide layer on a glass plate. Single-delta oxygen was released from the endoperoxide upon heating the organic film by means of an electrical discharge through the conductive indium oxide coating. The evolution of singlet-delta oxygen was determined by measuring the dimol emission signal at 634 nm. Comparison of the measured signal with an analytic model leads to two main conclusions: virtually all the oxygen being evolved is in the singlet-delta state and in the gas phase, and there is no significant quenching other than energy pooling on the time scale of the experiment (approximately 10 msec). The use of solid phase endoperoxide as a singlet-delta oxygen generator for an oxygen-iodine laser appears promising.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-172364 , NAS 1.26:172364
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    SSME interstage seal research (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Test results comprising direct and transverse force coefficients and leakage coefficients are reported for six seal configurations. All seals tested use the same smooth rotor and have the same constant minimum clearance. The following stator configurations were tested: (1) Smooth, (2) knurled pattern, (3) axially-grooved pattern with end seals, (4) diamond-grid roughened, (5) diamond-grid roughened with end seals, and (6) round-hole pattern. Comparison of the seals shows the Knurled-pattern stator to be the stiffest and the round-hole pattern stator to yield the largest net damping and the least leakage. The theory of reference is shown to substantially underestimate the stiffness and effective-added-mass coefficients, but do a reasonable job in predicting the net-damping-force coefficient.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-170995 , NAS 1.26:170995 , RD-1-84
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    Parametric analysis of hollow conductor parallel and coaxial transmission lines for high frequency space power distribution (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A parametric analysis was performed of transmission cables for transmitting electrical power at high voltage (up to 1000 V) and high frequency (10 to 30 kHz) for high power (100 kW or more) space missions. Large diameter (5 to 30 mm) hollow conductors were considered in closely spaced coaxial configurations and in parallel lines. Formulas were derived to calculate inductance and resistance for these conductors. Curves of cable conductance, mass, inductance, capacitance, resistance, power loss, and temperature were plotted for various conductor diameters, conductor thickness, and alternating current frequencies. An example 5 mm diameter coaxial cable with 0.5 mm conductor thickness was calculated to transmit 100 kW at 1000 Vac, 50 m with a power loss of 1900 W, an inductance of 1.45 micron and a capacitance of 0.07 micron-F. The computer programs written for this analysis are listed in the appendix.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-83601 , E-2013 , NAS 1.15:83601
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    Evaluation of propellant tank insulation concepts for low-thrust chemical propulsion systems (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An analytical evaluation of cryogenic propellant tank insulations for liquid oxygen/liquid hydrogen low-thrust 2224N (500 lbf) propulsion systems (LTPS) was conducted. The insulation studied consisted of combinations of N2-purged foam and multilayer insulation (MLI) as well as He-purged MLI-only. Heat leak and payload performance predictions were made for three Shuttle-launched LTPS designed for Shuttle bay packaged payload densities of 56 kg/cu m, 40 kg/cu m and 24 kg/cu m. Foam/MLI insulations were found to increase LTPS payload delivery capability when compared with He-purged MLI-only. An additional benefit of foam/MLI was reduced operational complexity because Orbiter cargo bay N2 purge gas could be used for MLI purging. Maximum payload mass benefit occurred when an enhanced convection, rather than natural convection, heat transfer was specified for the insulation purge enclosure. The enhanced convection environment allowed minimum insulation thickness to be used for the foam/MLI interface temperature selected to correspond to the moisture dew point in the N2 purge gas. Experimental verification of foam/MLI benefits was recommended. A conservative program cost estimate for testing a MLI-foam insulated tank was 2.1 million dollars. It was noted this cost could be reduced significantly without increasing program risk.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-168320 , NAS 1.26:168320 , BAC-D180-28273-1
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    Low loss injector for Space Shuttle main engine. Center director's discretionary fund (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An efficient propellant injection method to raise the Space Shuttle Main Engine (SSME) thrust and payload is discussed. Relatively large diameter injector elements with low pressure loss are recommended for the main combustion chamber and the pre-burners. Smaller losses admit more propellant flow which then raises thrust. Payload is not only gained by specific impulse but also by thrust. The chamber pressure is stabilized by selecting the proper cavity size for the injector elements while reducing the injection pressure loss which normally is kept high for stability. The rather large injector element recesses provide acoustic damping which makes baffles and acoustic absorbers unnecessary. A tenfold reduction of flow induced stresses which are rather high in the present design is shown. Relaxed tolerances, fewer elements, and better maintenance are offered. The study was conducted under a center director discretionary fund assignment.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-82567 , NAS 1.15:82567
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    Ion extraction capabilities of two-grid accelerator systems (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental investigation into the ion extraction capabilities of two-grid accelerator systems common to electrostatic ion thrusters is described. This work resulted in a large body of experimental data which facilitates the selection of the accelerator system geometries and operating parameters necessary to maximize the extracted ion current. Results suggest that the impingement-limited perveance is not dramatically affected by reductions in screen hole diameter to 0.5 mm. Impingement-limited performance is shown to depend most strongly on grid separation distance, accelerator hole diameter ratio, the discharge-to-total accelerating voltage ratio, and the net-to-total accelerating voltage ratio. Results obtained at small grid separation ratios suggest a new grid operating condition where high beam current per hole levels are achieved at a specified net accelerating voltage. It is shown that this operating condition is realized at an optimum ratio of net-to-total accelerating voltage ratio which is typically quite high. The apparatus developed for this study is also shown to be well suited measuring the electron backstreaming and electrical breakdown characteristics of two-grid accelerator systems.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-174621 , NAS 1.26:174621
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    Ion thruster performance model (1984)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: A model of ion thruster performance is developed for high flux density, cusped magnetic field thruster designs. This model is formulated in terms of the average energy required to produce an ion in the discharge chamber plasma and the fraction of these ions that are extracted to form the beam. The direct loss of high energy (primary) electrons from the plasma to the anode is shown to have a major effect on thruster performance. The model provides simple algebraic equations enabling one to calculate the beam ion energy cost, the average discharge chamber plasma ion energy cost, the primary electron density, the primary-to-Maxwellian electron density ratio and the Maxwellian electron temperature. Experiments indicate that the model correctly predicts the variation in plasma ion energy cost for changes in propellant gas (Ar, Kr and Xe), grid transparency to neutral atoms, beam extraction area, discharge voltage, and discharge chamber wall temperature. The model and experiments indicate that thruster performance may be described in terms of only four thruster configuration dependent parameters and two operating parameters. The model also suggests that improved performance should be exhibited by thruster designs which extract a large fraction of the ions produced in the discharge chamber, which have good primary electron and neutral atom containment and which operate at high propellant flow rates.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-174810 , NAS 1.26:174810
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    Investigation of a pulsed electrothermal thruster system (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The performance of an ablative wall Pulsed Electrothermal (PET) thruster is accurately characterized on a calibrated thrust stand, using polyethylene propellant. The thruster is tested for four configurations of capillary length and pulse length. The exhaust velocity is determined with twin time-of-flight photodiode stagnation probes, and the ablated mass is measured from the loss over ten shots. Based on the measured thrust impulse and the ablated mass, the specific impulse varies from 1000 to 1750 seconds. The thrust to power varies from .05 N/kW (quasi-steady mode) to .10 N/kW (unsteady mode). The thruster efficiency varies from .56 at 1000 seconds to .42 at 1750 seconds. A conceptual design is presented for a 40 kW PET propulsion system. The point design system performance is .62 system efficiency at 1000 seconds specific impulse. The system's reliability is enhanced by incorporating 20, 20 kW thruster modules which are fired in pairs. The thruster design is non-ablative, and uses water propellant, from a central storage tank, injected through the cathode.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-174768 , NAS 1.26:174768 , GTD-84-4
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    High altitude chemically reacting gas particle mixtures. Volume 2: Program manual for RAMP2 (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: All of the elements used in the Reacting and Multi-Phase (RAMP2) computer code are described in detail. The code can be used to model the dominant phenomena which affect the prediction of liquid and solid rocket nozzle and orbital plume flow fields.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-171827 , NAS 1.26:171827 , LMSC-HREC-TR-D867400-2-VOL-2
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    Preliminary evaluation of a liquid belt radiator for space applications (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The liquid belt radiator (LBR) is discussed. The LBR system operates either in the sensible heat mode or in the latent heat mode. Parametric analysis shows that the LBR may reduce the mass of heat pipe radiators by 70 to 90% when the LBR surface has a total emissivity in excess of 0.3. It is indicated that the diffusion pump oils easily meet this criteria with emissivities greater than 0.8. Measurements on gallium indicate that its emissivity is probably in excess of 0.3 in the solid state when small amounts of impurities are on the surface. The point design exhibits a characteristic mass of 3.1 kg/kW of power dissipation, a mass per unit prime radiating area of approximately 0.9 kg/sq ms and a total package volume of approximately 2.50 cubic m. This compares favorably with conventional technologies which have weights on the order of 4 kg/sq m.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-174807 , NAS 1.26:174807
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  • 68
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    High altitude chemically reacting gas particle mixtures. Volume 1: A theoretical analysis and development of the numerical solution (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The overall contractual effort and the theory and numerical solution for the Reacting and Multi-Phase (RAMP2) computer code are described. The code can be used to model the dominant phenomena which affect the prediction of liquid and solid rocket nozzle and orbital plume flow fields. Fundamental equations for steady flow of reacting gas-particle mixtures, method of characteristics, mesh point construction, and numerical integration of the conservation equations are considered herein.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-171826 , NAS 1.26:171826 , LMSC-HREC-TR-D867400-1-VOL-1
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    An analysis of return flux from the Space Shuttle Orbiter RCS engines (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The return flux from the Space Shuttle Orbiter reaction control system (RCS) engines to sensors in the open payload bay has been analyzed on the basis of Shuttle/Payload Contamination Evaluation (SPACE II) model predictions and orbital flight measurements. Model data are presented showing the variations of molecular return flux values with Orbiter orientation and instrument direction. The effects of multiple molecular collisions within RCS engine plumes and in their vicinity are discussed. These collisions significantly influence the amount of plume molecules returning to payload instruments and, therefore, the amount of contaminants received.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 84-0551
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  • 70
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    Computation of the Space Shuttle solid rocket booster nozzle start-up transient flow (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The first NASA Space Shuttle flight (STS-1) produced an overpressure wave that exceeded preflight predictions by as much as 5 to 1. This second overpressure wave occurred just after the solid rocket booster (SRB) igniter wave. To understand this overpressure phenomenon, a numerical simulation effort was undertaken. Both the SRB static firing test and STS-1 geometries were studied for two-dimensional, inviscid and viscous flow. The inviscid calculations did not produce significant second overpressure waves. However, the viscous calculations did produce second overpressure waves that qualitatively agree with experiment. These overpressure waves were present in both the static firing test and STS-1 geometries. This second overpressure wave is generated by the motion of the boundary layer separation point and the subsequent radial motion of the exhaust jet during the start-up of the SRB nozzle flow. The presence of the mobile launch platform exhaust hole wale amplifies this wave, but does not appear to be the source of any additional overpressure waves. The lack of good quantitative agreement between theory and experiment indicates that other overpressure sources, not accounted for by this simulation, may be present.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 84-0462
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  • 71
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    Rocket injector anomalies study. Volume 2: Results of parametric studies (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The employment of a existing computer program to simulate three dimensional two phase gas spray flows in liquid propellant rocket engines. This was accomplished by modification of an existing three dimensional computer program (REFLAN3D) with Euler/Lagrange approach for simulating two phase spray flow, evaporation and combustion. The modified code is referred to as REFLAN3D-SPRAY. Computational studies of the model rocket engine combustion chamber are presented. The parametric studies of the two phase flow and combustion shows qualitatively correct response for variations in geometrical and physical parameters. The injection nonuniformity test with blocked central fuel injector holes shows significant changes in the central flame core and minor influence on the wall heat transfer fluxes.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-174703 , NAS 1.26:174703 , CHAM-H3605/16-VOL-2
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    Electromagnetic thrusters for spacecraft prime propulsion (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The benefits of electromagnetic propulsion systems for the next generation of US spacecraft are discussed. Attention is given to magnetoplasmadynamic (MPD) and arc jet thrusters, which form a subset of a larger group of electromagnetic propulsion systems including pulsed plasma thrusters, Hall accelerators, and electromagnetic launchers. Mission/system study results acquired over the last twenty years suggest that for future prime propulsion applications high-power self-field MPD thrusters and low-power arc jets have the greatest potential of all electromagnetic thruster systems. Some of the benefits they are expected to provide include major reductions in required launch mass compared to chemical propulsion systems (particularly in geostationary orbit transfer) and lower life-cycle costs (almost 50 percent less). Detailed schematic drawings are provided which describe some possible configurations for the various systems.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 84-1446
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  • 73
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    Radiative resistojet performance characterization tests (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The test article, test approach, data analysis and results of a study undertaken to characterize performance of the augmentation section of the Rocket Research Company Augmented Catalytic Thruster as a gas resistojet using hydrogen, nitrogen and ammonia as propellants are described. This renewed interest in resistojets is a result of propulsion systems definition studies which indicate potential application to space station auxiliary propulsion.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-174763 , NAS 1.26:174763 , REPT-84-R-958
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  • 74
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    Research pressure instrumentation for NASA Space Shuttle main engine (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The development of prototype pressure transducers which are targeted to meet the Space Shuttle Main Engine SSME performance design goals is discussed. The fabrication, testing and delivery of 10 prototype units is examined. Silicon piezoresistive strain sensing technology is used to achieve the objectives of advanced state-of-the-art pressure sensors in terms of reliability, accuracy and ease of manufacture. Integration of multiple functions on a single chip is the key attribute of this technology.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-173932 , NAS 1.26:173932
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  • 75
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    Rocket injector anomalies study. Volume 1: Description of the mathematical model and solution procedure (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The capability of simulating three dimensional two phase reactive flows with combustion in the liquid fuelled rocket engines is demonstrated. This was accomplished by modifying an existing three dimensional computer program (REFLAN3D) with Eulerian Lagrangian approach to simulate two phase spray flow, evaporation and combustion. The modified code is referred as REFLAN3D-SPRAY. The mathematical formulation of the fluid flow, heat transfer, combustion and two phase flow interaction of the numerical solution procedure, boundary conditions and their treatment are described.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-174702 , NAS 1.26:174702 , CHAM-H3605/15-VOL-1
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  • 76
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    Environmentally-induced voltage limitations in large space power systems (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Large power systems proposed for future space missions imply higher operating voltage requirements which, in turn, will interact with the space plasma environment. The effects of these interactions can only be inferred because of the limited data base of ground simulations, small test samples, and two space flight experiments. This report evaluates floating potentials for a 100 kW power system operating at 300, 500, 750, and 1000 volts in relation to this data base. Of primary concern is the possibility of discharging to space. The implications of such discharges were studied at the 500 volt operational setting. It was found that discharging can shut down the power system if the discharge current exceeds the array short circuit current. Otherwise, a power oscillation can result that ranges from 2 to 20 percent, depending upon the solar array area involved in the discharge. Means of reducing the effect are discussed.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: IEEE Transactions on Nuclear Science (ISSN 0018-9499); NS-31; 1381-138
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  • 77
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    Oxygen-hydrogen thrusters for Space Station auxiliary propulsion systems (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The feasibility and technology requirements of a low-thrust, high-performance, long-life, gaseous oxygen (GO2)/gaseous hydrogen (GH2) thruster were examined. Candidate engine concepts for auxiliary propulsion systems for space station applications were identified. The low-thrust engine (5 to 100 lb sub f) requires significant departure from current applications of oxygen/hydrogen propulsion technology. Selection of the thrust chamber material and cooling method needed or long life poses a major challenge. The use of a chamber material requiring a minimum amount of cooling or the incorporation of regenerative cooling were the only choices available with the potential of achieving very high performance. The design selection for the injector/igniter, the design and fabrication of a regeneratively cooled copper chamber, and the design of a high-temperature rhenium chamber were documented and the performance and heat transfer results obtained from the test program conducted at JPL using the above engine components presented. Approximately 115 engine firings were conducted in the JPL vacuum test facility, using 100:1 expansion ratio nozzles. Engine mixture ratio and fuel-film cooling percentages were parametrically investigated for each test configuration.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-175691 , JPL-9950-974 , NAS 1.26:175691 , FR-956457-F-1
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    Alternate nozzle ablative materials program (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Four subscale solid rocket motor tests were conducted successfully to evaluate alternate nozzle liner, insulation, and exit cone structural overwrap components for possible application to the Space Shuttle Solid Rocket Motor (SRM) nozzle asasembly. The 10,000 lb propellant motor tests were simulated, as close as practical, the configuration and operational environment of the full scale SRM. Fifteen PAN based and three pitch based materials had no filler in the phenolic resin, four PAN based materials had carbon microballoons in the resin, and the rest of the materials had carbon powder in the resin. Three nozzle insulation materials were evaluated; an aluminum oxide silicon oxide ceramic fiber mat phenolic material with no resin filler and two E-glass fiber mat phenolic materials with no resin filler. It was concluded by MTI/WD (the fabricator and evaluator of the test nozzles) and NASA-MSFC that it was possible to design an alternate material full scale SRM nozzle assembly, which could provide an estimated 360 lb increased payload capability for Space Shuttle launches over that obtainable with the current qualified SRM design.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-175658 , JPL-PUB-84-58 , NAS 1.26:175658
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    Theory of ion acceleration with closed electron drift (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Spacecraft and Rockets (ISSN 0022-4650); 21; 558-562
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  • 80
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    Tethered propellant resupply technique for space stations (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: One of the primary functions of the space station is related to the propellant resupply of orbital transfer vehicles, orbital maneuvering vehicles, and satellites. Difficulties arise in the case of an acquisition of cryogenic propellants by means of a use of zero-gravity techniques. The use of the 'tethered propellant resupply technique' is, therefore, considered. A study is being conducted to determine the feasibility, design requirements, and operational limitations of this technique. Attention is given to aspects of gravity feed, transfer method selection, requirements related to the orbital transfer vehicle, hazard clearance, attitude control, depot operations, end mass velocity, the microgravity laboratory, and concept evaluation activities.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: IAF PAPER 84-442
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  • 81
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    Electrodynamic tethers. I - Power generation in LEO. II - Thrust for propulsion & power storage (1984)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The power gain and thrust for plasma engines available by unreeling 10 km of insulated Al wire from a spacecraft are investigated. The wire, unreeling in the vertical, would cut the earth's magnetic field lines, thereby generating 20 kW of power in the wire. A drag loss of 20.4 kJ/sec would reduce the power gain to 18.7 kW, an efficiency of 92 percent. Thicker wires could push the power gain to 1 MW at 95 percent efficiency. Conductive 'balloons' at the ends of the tether would function as ionospheric 'brushes' to complete the circuit. Reversing the IXB force by employing on-board stored power would drive the tether current against the induced voltage, providing a 1 N thrust for 8 kW of energy consumed, which could be supplied by solar panels during the day portion of orbit. The equivalent thrust by conventional stationkeeping means would consume 8000 kg of propellant/yr. Techniques for stabilizing the tether in the presence of variable magnetic fields are discussed.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: IAF PAPER 84-440
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    Development of the platelet micro-orifice injector (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: For some time to come, liquid rocket engines will continue to provide the primary means of propulsion for space transportation. The injector represents a key to the optimization of engine and system performance. The present investigation is concerned with a unique injector design and fabrication process which has demonstrated performance capabilities beyond that achieved with more conventional approaches. This process, which is called the 'platelet process', makes it feasible to fabricate injectors with a pattern an order of magnitude finer than that obtainable by drilling. The fine pattern leads to an achievement of high combustion efficiencies. Platelet injectors have been identified as one of the significant technology advances contributing to the feasibility of advanced dual-fuel booster engines. Platelet injectors are employed in the Space Shuttle Orbit Maneuvering System (OMS) engines. Attention is given to injector design theory as it relates to pattern fineness, a description of platelet injectors, and test data obtained with three different platelet injectors.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: IAF PAPER 84-299
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  • 83
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    Design and analysis report for the RL10-2B breadboard low thrust engine (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The breadboard low thrust RL10-2B engine is described. A summary of the analysis and design effort to define the multimode thrust concept applicable to the requirements for the upper stage vehicles is provided. Baseline requirements were established for operation of the RL10-2B engine under the following conditions: (1) tank head idle at low propellant tank pressures without vehicle propellant conditioning or settling thrust; (2) pumped idle at a ten percent thrust level for low G deployment and/or vehicle tank pressurization; and (3) full thrust (15,000 lb.). Several variations of the engine configuration were investigated and results of the analyses are included.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-174857 , NAS 1.26:174857 , FR-18046-3
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  • 84
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    Thrust chamber performance using Navier-Stokes solution (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The viscous, axisymmetric flow in the thrust chamber of the space shuttle main engine (SSME) was computed on the CRAY 205 computer using the general interpolants method (GIM) code. Results show that the Navier-Stokes codes can be used for these flows to study trends and viscous effects as well as determine flow patterns; but further research and development is needed before they can be used as production tools for nozzle performance calculations. The GIM formulation, numerical scheme, and computer code are described. The actual SSME nozzle computation showing grid points, flow contours, and flow parameter plots is discussed. The computer system and run times/costs are detailed.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-171346 , NAS 1.26:171346 , LMSC-HREC-TR-D951729
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    SSME thrust chamber simulation using Navier-Stokes equations (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The capability of the PHOENICS fluid dynamics code in predicting two-dimensional, compressible, and reacting flow in the combustion chamber and nozzle of the space shuttle main engine (SSME) was evaluated. A non-orthogonal body fitted coordinate system was used to represent the nozzle geometry. The Navier-Stokes equations were solved for the entire nozzle with a turbulence model. The wall boundary conditions were calculated based on the wall functions which account for pressure gradients. Results of the demonstration test case reveal all expected features of the transonic nozzle flows. Of particular interest are the locations of normal and barrel shocks, and regions of highest temperature gradients. Calculated performance (global) parameters such as thrust chamber flow rate, thrust, and specific impulse are also in good agreement with available data.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-171349 , NAS 1.26:171349 , CHAM-4070/3
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  • 86
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    Analytical and experimental investigation of rubbing interaction in labyrinth seals for a liquid hydrogen fuel pump (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Cracking of the titanium knife edges on the labyrinth seals of the liquid hydrogen fuel pump in the Space Shuttle main engine is considered. Finite element analysis of the thermal response of the knife edge in sliding contact with the wear ring surface shows that interfacial temperatures can be quite high and they are significantly influenced by the thermal conductivity of the surfaces in rubbing contact. Thermal shock experiments on a test specimen similar to the knife edge geometry demonstrate that cracking of the titanium alloy is possible in a situation involving repeated thermal cycles over a wide temperature range, as might be realized during a rub in the liquid hydrogen fuel pump. High-speed rub interaction tests were conducted using a representative knife edge and seal geometry over a broad range of interaction rates and alternate materials were experimentally evaluated. Plasma-sprayed aluminum-graphite was found to be significantly better than presently used aluminum alloy seals from the standpoint of rub performance. Ion nitriding the titanium alloy knife-edges also improved rub performance compared to the untreated baseline.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-174657 , NAS 1.26:174657 , CREARE-TN-371
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  • 87
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    Experimental investigation of the pulsed electrothermal (PET) thruster (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Burton et al. (1982) have discussed the theory of the Pulsed Electrothermal (PET) thruster, a device which in principle can operate with 70 percent efficiency at a specific impulse of 1000 seconds and higher. It is pointed out that this level of performance would be particularly attractive for orbit raising of large satellites and other near-earth missions, which cannot be easily accomplished by chemical propulsion. The present investigation is concerned with two PET thruster operating modes. A PET thruster was built and tested on a thrust stand. Exhaust velocities for polyethylene propellant vary from 20 to 27 km/sec. Single pulse specific impulse and efficiency measurements based on ablated mass show a thruster efficiency of 37-56 percent in the time range from 1000 to 1750 seconds. It is believed that an improved design with a thruster efficiency in the range from 70 to 80 percent might be possible.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 84-1386
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  • 88
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    Selection of burn-resistant materials for oxygen-driven turbopumps (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: NASA goals for reusable space-based, high-performance orbit transfer vehicle propulsion systems have resulted in a need for oxygen/hydrogen engines which include lightweight, highly reliable, liquid oxygen pumps. The selection of ignition- and burn-resistant materials is a major factor in the design of a compact 75,000-rpm turbopump which can deliver 6 lbM/sec of liquid oxygen at 5,000 psia. The potential operational hazards of rubbing friction and impact of foreign particles at high velocity were investigated experimentally for a wide range of candidate materials, i.e., nickel, copper, monel, 316 Stainless Steel, Hastelloy-X, Invar-36, and silicon carbide. Test parameters included oxygen pressure and temperature up to 5,000 psia and 800 F, respectively. The effect of increasing the O2 pressure from 1000 to 5000 psi is discussed. The applicability of the candidate materials to oxygen pump design was ranked by comparing the experimental results among themselves and with an analytically determined parameter, i.e., the burn factor. Nickel and copper demonstrated superior resistance to ignition and burn in the friction rubbing and particle impact tests relative to monel, stainless steel, and nickel-iron base superalloys.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 84-1287
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  • 89
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    Space power technology into the 21st century (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: This paper discusses the space power systems of the early 21st century. The focus is on those capabilities which are anticipated to evolve from today's state-of-the-art and the technology development programs presently in place or planned for the remainder of the century. The power system technologies considered include solar thermal, nuclear, radioisotope, photovoltaic, thermionic, thermoelectric, and dynamic conversion systems such as the Brayton and Stirling cycles. Energy storage technologies considered include nickel hydrogen biopolar batteries, advanced high energy rechargeable batteries, regenerative fuel cells, and advanced primary batteries. The present state-of-the-art of these space power and energy technologies is discussed along with their projections, trends and goals. A speculative future mission model is postulated which includes manned orbiting space stations, manned lunar bases, unmanned earth orbital and interplanetary spacecraft, manned interplanetary missions, military applications, and earth to space and space to space transportation systems. The various space power/energy system technologies anticipated to be operational by the early 21st century are matched to these missions.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 84-1136
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  • 90
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    A model for mercury orificed hollow cathodes - Theory and experiment (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA Journal (ISSN 0001-1452); 22; 1405-141
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  • 91
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    Laser propulsion (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An introduction to thermal laser propulsion is presented. This form of rocket propulsion uses a laser beam from a remotely-located laser to heat a propellant gas, which is then expanded in a conventional way to produce thrust. This propulsion scheme has the potential for producing high specific impulse (greater than 1000 s) at moderate to high thrust (1000 lbs). Laser propulsion can thus fill a niche in propulsion for spaceflight missions which can be filled by no other practical scheme. The system analyses and some of the experimental and theoretical studies which have been performed are briefly reviewed. Production of thrust by a pulsed laser has been demonstrated on a laboratory scale at an Isp of 1000 s in hydrogen. While more work is needed, it seems apparent that laser propulsion has an important and unique capability which should be pursued, and should be considered for space missions in the 1990's and beyond.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 84-1445
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  • 92
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    Development of a bipropellant propulsion module for the Galileo spacecraft (1984)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The bipropellant engine retropropulsion module (RPM) of the Jupiter-exploration Galileo spacecraft provides attitude control, trajectory correction, a deep space velocity increment, a spacecraft deflection maneuver for an entry capsule, a retromaneuver for Jupiter approach, and a 'perijove raise' maneuver for the Jovian system tour of the spacecraft. Attention is given to the design features and performance capabilities of the RPM, which uses a set of 12 10-N thrusters in conjunction with a 400-N engine. Tanks contain 935 kg of propellant to satisfy mission requirements.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 84-1233
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  • 93
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    Vapor flow into a capillary propellant-acquisition device (1984)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Previously cited in issue 16, p. 2321, Accession no. A83-36371
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Spacecraft and Rockets (ISSN 0022-4650); 21; 267-273
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  • 94
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    Proposed system design for a 20 kW pulsed electrothermal thruster (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A conceptual design is presented for a Pulsed Electrothermal (PET) propulsion system for the Air Force Space Based Radar satellite, which has a mass of 7000 kg. The proposed system boosts the SBR satellite from 150 n.m. to 600 n.m. with a 4 deg plane change, for a total mission Delta v of 1 km/sec. Satellite power available is 50 kW, and 45 kW are used to drive two water-injected 20 kW PET thrusters, delivering 5.6 N thrust to the SBR at 1000 seconds specific impulse. The predicted mission trip time is 15 days. The proposed system consumes 850 kg of water propellant, stored in a central tank and injected with pressurized helium. Component mass estimates based on space-qualified hardware are presented for the propellant handling, power conditioning and thruster subsystems. The estimated total mass is 400 kg and the propulsion system specific mass is alpha = 10 kg/kW. The proposed system efficiency of 0.62 at 1000 seconds specific impulse is supported by experimental performance measurements.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 84-1387
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  • 95
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    Demonstration of a new electrothermal thruster concept (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The design and test of a microwave electrothermal thruster are described. The device, which employs a coaxial microwave discharge, was tested in nitrogen gas with 200-600 W of 2.45-GHz input power. Experimental measurements of thrust, specific impulse, and energy efficiency are presented for different flow and discharge pressures. Measured energy efficiencies varied between 30-60 percent and the performance compared favorably with other electrothermal thrusters operating in nitrogen gas. The experimental performance demonstrated the feasibility of the concept.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Applied Physics Letters (ISSN 0003-6951); 44; 1014-101
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  • 96
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    Progress in space nuclear reactor power systems technology development - The SP-100 program (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Activities related to the development of high-temperature compact nuclear reactors for space applications had reached a comparatively high level in the U.S. during the mid-1950s and 1960s, although only one U.S. nuclear reactor-powered spacecraft was actually launched. After 1973, very little effort was devoted to space nuclear reactor and propulsion systems. In February 1983, significant activities toward the development of the technology for space nuclear reactor power systems were resumed with the SP-100 Program. Specific SP-100 Program objectives are partly related to the determination of the potential performance limits for space nuclear power systems in 100-kWe and 1- to 100-MW electrical classes. Attention is given to potential missions and applications, regimes of possible space power applicability, safety considerations, conceptual system designs, the establishment of technical feasibility, nuclear technology, materials technology, and prospects for the future.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 84-1132
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  • 97
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    Laser thermal propulsion (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Laser thermal propulsion (LTP) is studied for the case in which laser power is absorbed by a small very high-temperature plasma (about 20,000 K) and transferred to the remainder of the pure hydrogen propellant by radiation and mixing. This concept could lead to the realization of a lightweight orbital transfer vehicle propulsion system having a specific impulse in the range 1000-2000 s. Approximately 12 percent of the input power may be radiated to the thruster walls, and 15 percent of the total propellant flow must be heated to 20,000 K to provide a bulk temperature of 5000 K prior to expansion. Three principal research issues identified are: (1) conditions for hydrogen plasma ignition, (2) control of the plasma position within the laser beam, plasma stability, and plasma absorption efficiency, and (3) characterization of the mixing of the plasma and buffer flows.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 98
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    Injection module thrust vector control (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A three-axis controlled Injection Module propelled by a Star 48 solid rocket motor has been considered for use as a final 'kick' stage to supplement the Inertial Upper Stage in a proposed launch option for the Galileo and International Solar Polar missions. A flight control law for the Injection Module is developed. A position plus rate control law is considered, with integral, path guidance, and derived rate terms added for improved pointing accuracy and fuel efficiency. Selection of control gains is accomplished with the help of analytical limit cycle expressions and verified by computer simulation of the closed loop system. A computer simulation of the flight control system is built around a rigid spacecraft model with gyro dynamics and thruster delays included. Models for pitch/yaw/roll disturbance torques are included. Through simple gain changes the proposed flight control law is shown to accommodate the widely different mass properties of the Galileo and International Solar Polar spacecraft. Pointing accuracies of better than the desired 0.2 degrees are achieved.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AAS PAPER 83-324
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  • 99
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    Cryogenic Fluid Management Facility (1984)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The Cryogenic Fluid Management Facility (CFMF) is a reusable test bed which is designed to be carried into space in the Shuttle cargo bay to investigate systems and technologies required to efficiently and effectively manage cryogens in space. The facility hardware is configured to provide low-g verification of fluid and thermal models of cryogenic storage, transfer concepts and processes. Significant design data and criteria for future subcritical cryogenic storage and transfer systems will be obtained. Future applications include space-based and ground-based orbit transfer vehicles (OTV), space station life support, attitude control, power and fuel depot supply, resupply tankers, external tank (ET) propellant scavenging, space-based weapon systems and space-based orbit maneuvering vehicles (OMV). This paper describes the facility and discusses the cryogenic fluid management technology to be investigated. A brief discussion of the integration issues involved in loading and transporting liquid hydrogen within the Shuttle cargo bay is also included.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 84-1340
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    Space Station propulsion analysis study (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: This paper summarizes the impacts on the weight, volume and power usage of a manned space station and its 90-day resupply for three integrated, auxiliary propulsion subsystems. The study was performed in coordination with activities of the Space Station Concept Development Group (CDG). The study focused on three space station propulsion high-low thrust options that make use of fluids that will be available on the manned space station. Specific uses of carbon dioxide, water and cryogen boiloff were considered. For each of the options the increase in station hardware mass and volume to accommodate the dual thrust option is offset by the resupply savings, relative to the reference hydrazine system, after one to several resupplies. Over the life of the station the savings in cost of logistics could be substantial. The three options are examples of alternative technology paths that, because of the opportunity they provide for integration with the environmental control life support system (ECLSS) and OTV propellant storage systems, may reduce the scarring which is required on the early station to meet the increasing propulsion requirements of the growth station.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 84-1326
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