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  • Other Sources  (202)
  • SPACECRAFT PROPULSION AND POWER  (202)
  • 1985-1989  (202)
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  • 1986  (202)
  • 1
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    Automating the U.S. Space Station's electrical power system (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: NASA's Lewis Research Center is developing a highly automated system for the generation, storage and distribution of electrical power aboard the projected Space Station. This autonomous power system will employ conventional algorithms, enhanced by expert systems, to schedule power, allocate energy, diagnose causes of failure, propose goals, prepare plans for their implementation, evaluate their consequences, and select optimum plans for their execution. While crew-interactive expert systems will be ready for the initial Space Station, total system autonomy is expected to require additional development time.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Optical Engineering (ISSN 0091-3286); 25; 1181-118
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  • 2
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    Advanced solar thermal technologies for the 21st century (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: The paper considers the present status of solar thermal dynamic space power technologies and projects the various attributes of these systems into the future, to the years 2000 and 2010. By the year 2000, collector weights should decrease from 1.25 kg/sq m (1985 value) to about 1.0 kg/sq m. The specific weight is also expected to decrease from 6.0 kg/kw. By the year 2010, slight improvements in the free piston Stirling energy conversion system are postulated with efficiencies reaching 32 percent. In addition, advanced concentrator concepts should be operational.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 3
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    Alkaline RFC Space Station prototype - 'Next step Space Station' (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: The regenerative fuel cell, a candidate technology for the Space Station's energy storage system, is described. An advanced development program was initiated to design, manufacture, and integrate a regenerative fuel cell Space Station prototype (RFC SSP). The RFC SSP incorporates long-life fuel cell technology, increased cell area for the fuel cells, and high voltage cell stacks for both units. The RFC SSP's potential for integration with the Space Station's life support and propulsion systems is discussed.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 4
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    Space Station power system selection (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: The Space Station power system selection process is described with attention given to management organization and technical considerations. A hybrid power system was chosen because of the large life cycle cost savings. The power management and distribution system that was chosen was the 400 Hz system.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 5
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    A feasibility assessment of nuclear reactor power system concepts for the NASA Growth Space Station (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: A preliminary feasibility assessment of the integration of reactor power system concepts with a projected growth Space Station architecture was conducted to address a variety of installation, operational, disposition and safety issues. A previous NASA sponsored study, which showed the advantages of Space Station - attached concepts, served as the basis for this study. A study methodology was defined and implemented to assess compatible combinations of reactor power installation concepts, disposal destinations, and propulsion methods. Three installation concepts that met a set of integration criteria were characterized from a configuration and operational viewpoint, with end-of-life disposal mass identified. Disposal destinations that met current aerospace nuclear safety criteria were identified and characterized from an operational and energy requirements viewpoint, with delta-V energy requirement as a key parameter. Chemical propulsion methods that met current and near-term application criteria were identified and payload mass and delta-V capabilities were characterized. These capabilities were matched against concept disposal mass and destination delta-V requirements to provide a feasibility of each combination.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 6
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    Alkaline water electrolysis technology for Space Station regenerative fuel cell energy storage (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: The regenerative fuel cell system (RFCS), designed for application to the Space Station energy storage system, is based on state-of-the-art alkaline electrolyte technology and incorporates a dedicated fuel cell system (FCS) and water electrolysis subsystem (WES). In the present study, emphasis is placed on the WES portion of the RFCS. To ensure RFCS availability for the Space Station, the RFCS Space Station Prototype design was undertaken which included a 46-cell 0.93 cu m static feed water electrolysis module and three integrated mechanical components.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 7
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    Evaluation of advanced high rate Li-SOCl2 cells (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: Under NASA sponsorship, JPL is developing advanced, high rate Li-SOCl2 cells for future space missions. As part of this effort, Li-SOCl2 cells of various designs were examined for performance and safety. The cells differed from one another in several aspects, such as: nature of carbon cathode, catalysts, cell configuration, case polarity, and safety devices. Performance evaluation included constant-current discharge over a range of currents and temperatures. Abuse-testing consisted of shortcircuiting, charging, and over-discharge. Energy densities greater than 300 Wh/Kg at the C/2 rate were found for some designs. A cell design featuring a high-surface-area carbon cathode was found to deliver nearly 500 Wh/Kg at moderate discharge rates. Temperature influenced the performance significantly.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 8
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    Parametric and cycle tests of a 40-AH bipolar nickel-hydrogen battery (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: The performance of a 12 V, 40 ampere-hour bipolar battery during various charge current, discharge current, temperature, and pressure operating conditions is investigated. The cell voltages, temperatures, ampere-hours, and watt-hours derived from the charge/discharge cycle tests are studied. Consideration is given to battery voltage and discharge capacity as a function of discharge current, the correlation between energy delivered on a discharge and battery temperature, battery voltage response to pulse discharges, and the voltage-temperature relationship. The data reveal that the bipolar Ni-H battery is applicable to high power systems.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 9
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    Advanced photovoltaic solar array program - A preliminary assessment (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: Two solar array designs developed for the Advanced Photovoltaic Solar Array program are described. The goal of the program is to develop solar arrays with higher mass specific power and power density and good robustness. The specific design requirements are: a beginning-of-life value of 130 W/kg, and end-of-life goals of 105 W/kg and 110 W/sq m. The two array-wing designs consisted of a single blanket. The differences in the blanket material (25 micron-thick Kapton versus 50 micron-thick carbon-loaded Kapton), solar cells (100 micron-thick wrap around versus 50 micron-thick 2 x 4 cm planar contact cells), and performance objectives (proposed industry requirements versus mission objectives) of the two designs are examined.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 10
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    Development of flight ready fifty-micron-thick silicon solar cell array module technology (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: The development of ultrathin silicon solar cell array modules from initial design to flight testing is discussed. Three 80-cell modules were subjected to the thermal soak test, the LEO thermal cycle test, and the solar array flight experiment, and six 48-cell welded modules were evaluated in the geosynchronous orbit thermal cycle test. It is observed that the electrical performance of the modules was not affected by the different environmental conditions. The automatic assembly of the cell modules, in particular the welding and solar cell glassing operation, is described. The specific power capabilities of Kapton, Kapton-Kevlar-Kapton, Kapton-graphite-Kapton, and Kapton-graphite-aluminum honeycomb-graphite solar array designs are assessed.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 11
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    Electrical power system design for the U.S. Space Station (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: The power systems for the Space Station manned core and platforms that have been selected in definition studies are described in this paper. The selected system for the platforms uses silicon arrays and Ni-H2 batteries. The power system for the manned core is a hybrid employing arrays and batteries identical to those on the platform along with solar dynamic modules using either Brayton or organic Rankine engines. The power system requirements, candidate technologies, and configurations that were considered, and the basis for selection, are discussed.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 12
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    Controls, health assessment, and conditional monitoring for large, reusable, liquid rocket engines (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: Past and future progress in the performance of control systems for large, liquid rocket engines typified such as current state-of-the-art, the Shuttle Main Engine (SSME), is discussed. Details of the first decade of efforts, which culminates in the F-1 and J-2 Saturn engines control systems, are traced, noting problem modes and improvements which were implemented to realize the SSME. Future control system designs, to accommodate the requirements of operation of engines for a heavy lift launch vehicle, an orbital transfer vehicle and the aerospace plane, are summarized. Generic design upgrades needed include an expanded range of fault detection, maintenance as-needed instead of as-scheduled, reduced human involvement in engine operations, and increased control of internal engine states. Current NASA technology development programs aimed at meeting the future control system requirements are described.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 13
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    Rocket nozzle lip flow by direct simulation Monte Carlo method (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Spacecraft and Rockets (ISSN 0022-4650); 23; 363-367
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  • 14
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    A 25-LBF gaseous oxygen/gaseous hydrogen thruster for space station application (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: A prototype 25 lb sub f gaseous oxygen/gaseous hydrogen thruster for Space Station propulsion application was designed and fabricated by Rocketdyne and endurance tested at the NASA/Marshall space Flight Center. The thruster incorporates a regeneratively cooled thrust chamber with a nozzle exit area ratio of 30, a 12-element coaxial injector, a spark igniter, and close-coupled propellant valves. Test results indicate that all major technology issues for long-life gaseous oxygen/gaseous hydrogen thrusters for Space Station application have been resolved.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Johns Hopkins Univ., The 1986 JANNAF Propulsion Meeting, Volume 1; p 539-546
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  • 15
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    Uprated OMS engine for upper stage propulsion (1986)
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    Publication Date: 2011-08-19
    Description: The results of a pre-development component demonstration program on the use of a gas generator-driven turbopump that increases the Space Shuttle's Orbital Maneuvering Engine (OME) operating pressure are given. Tests and analysis confirm the the capability of the concept to meet or exceed performance and life requirements. Storable propellant upper stage concepts are also discussed.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Johns Hopkins Univ., The 1986 JANNAF Propulsion Meeting, Volume 1; p 369-378
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  • 16
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    Heat transfer in space power and propulsion systems (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: NASA's planned Space Station has projected power requirements in the 75-300 kW range; attention is presently given to the range of power system configurations thus far proposed. These are a silicon solar cell system incorporating regenerative fuel cell or battery storage, with a 10-year lifetime, a solar-dynamic power system with phase-change or regenerative fuel cell energy storage, and a combination of these two alternatives. A development status evaluation is also given for the propulsion systems that may be used by next-generation boosters. These include such novel airbreathing systems as turboramjets, air liquefaction cycle rockets, airturboramjet/rockets, and supersonic combustion ramjets.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Mechanical Engineering (ISSN 0025-6501); 108; 40-52
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  • 17
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    Space Station 20-kHz power management and distribution system (1986)
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    Publication Date: 2011-08-19
    Description: During the conceptual design phase a 20-kHz power distribution system was selected as the reference for the Space Station. The system is single-phase 400 VRMS, with a sinusoidal wave form. The initial user power level will be 75 kW with growth to 300 kW. The high-frequency system selection was based upon considerations of efficiency, weight, safety, ease of control, interface with computers, and ease of paralleling for growth. Each of these aspects will be discussed as well as the associated trade-offs involved. An advanced development program has been instituted to accelerate the maturation of the high-frequency system. Some technical aspects of the advanced development will be discussed.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 18
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    20 kHz Space Station power system (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: The Space Station represents the next major U.S. commitment in space. The efficient delivery of power to multiple user loads is key to that success. In 1969, NASA Lewis Research Center began a series of studies with component and circuit developments that led to the high frequency bi-directional, four quadrant resonant driven converter. Additional studies and subsequent developments into the early 1980's have shown how the high frequency ac power system could provide overall advantages to many aerospace power systems. Because of its wide versatility, it also has outstanding advantages for the Space Station Program and its wide range of users. High frequency ac power provides higher efficiency, lower cost, and improved safety. The 20 kHz power system has exceptional flexibility, is inherently user friendly, and is compatible with all types of energy sources - photovoltaic, solar dynamic, rotating machines or nuclear Lewis distribution system testbed. The testbed demonstrates flexibility, versatility, and transparency to user technology as well as high efficiency, low mass, and reduced volume.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 19
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    The Space Shuttle, fluid dynamics, and computer graphics (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: Several samples are presented of CFD-generated flow ribbon images of fuel flow in the SSME. The images result from simulations of H2 flow through a design iteration of the main injector, the hot gas manifold and the fuel bowl section. The color-coded ribbons provide three-dimensional perspectives which indicate the direction and orientation of the flow and the velocity magnitude in intervals along the flow path.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: IEEE Computer Graphics and Applications (ISSN 0272-1716); 6; 6
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  • 20
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    Power is the keystone (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: An evaluation is made of the various technologies that have been considered for incorporation into the NASA Space Station's solar power system. A major feature of the system is noted to be the use of both 25 kW capacity of photovoltaic power and two 25-kW turbine-driven generators based on the heating of a working fluid by a mirror concentrator dish. Fuel cells will be used to store excess electrical energy, together with nickel-cadmium batteries. The selection of this manned Space Station power system was arrived at through a comparison of six different configurations.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Aerospace America (ISSN 0740-722X); 24; 36-38
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  • 21
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    Test program to provide confidence in liquid oxygen cooling of hydrocarbon fueled rocket thrust chambers (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: In previous tests of liquid oxygen cooling of hydrocarbon fueled rocket engines, small oxygen leaks developed at the throat of the thrust chamber and film cooled the hot gas side of the chamber wall without resulting in catastrophic failure. However, more testing is necessary to demonstrate that a catastropic failure would not occur if cracks developed further upstream between the injector and the throat, where the boundary layer has not been established. Since under normal conditions cracks are expected to form in the throat region of the thrust chamber, cracks must be initiated artificially in order to control their location. Several methods of crack initiation are discussed here.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Johns Hopkins Univ., The 1986 JANNAF Propulsion Meeting, Volume 1; p 361-367
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  • 22
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    Experimental thrust performance of a high area-ratio rocket nozzle (1986)
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    Publication Date: 2011-08-19
    Description: An experimental investigation was conducted to determine the thrust performance attainable from high-area-ratio rocket nozzles. A modified Rao-contoured nozzle with an expansion area of 1030 was test fired with hydrogen-oxygen propellants at altitude conditions. The nozzle was also tested as a truncated nozzle, at an expansion area ratio of 428. Thrust coefficient and thrust coefficient efficiency values are presented for each configuration at various propellant mixture ratios (oxygen/fuel). Several procedural techniques were developed permitting improved measurement of nozzle performance. The more significant of these were correcting the thrust for the aneroid effects, determining the effective chamber pressure, and referencing differential pressure transducers to a vacuum reference tank.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Johns Hopkins Univ., The 23rd JANNAF Combustion Meeting, Volume 1; p 585-599
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  • 23
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    An in-situ measurement of particulates from solid rocket motors fired in space (1986)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Current models exist that predict the damage caused by the impact of aluminum oxide exhaust particles as well as their lifetime in useable space. In these models, two necessary inputs are the size and flux of the particles. An experiment, referred to as the Plume Witness Plate, was designed for the Remote Manipulator System of the space shuttle orbiter to measure in-situ the flux and material effects of a solid rocket motor (SRM) firing in space. Five different types of samples were used to provide a broad range of substances: (1) fused quartz glass (representative of orbiter windows); (2) germanium micrometeroid capture cells; (3) orbiter HRTS tiles from the thermal protection system; (4) Kapton foil; and (5) metallic disks of aluminum, copper, titanium, graphite epoxy, and gold. The analyses of the data show excellent agreement with ground-based SRM firings in terms of particle size distribution and mass distribution. The Particle Impact Damage Integrator computer model used to calculate potential damage of orbiter surfaces by SRM exhaust plumes agrees favorable with the results in terms of particle size and velocity distributions though it may be conservative by as much as 20%.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Lunar and Planetary Inst. Trajectory Determinations and Collection of Micrometeoroids on the Space Station; p 43-44
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  • 24
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    Investigation and modeling of space shuttle main engine shutdown transient chugging (1986)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The space shuttle main engines experience a low frequency pressure pulsation in both the fuel and oxidizer preburners during the shutdown transient. This pressure pulsation, called chugging, has been linked to undesirable bearing loads and possible damage to the spark ignitor supply piping for the fuel preburner. The problem is briefly described and a model is proposed that includes: (1) a transient stirred tank reactor model for the combustion chamber, (2) a resistance capacitance model for the supply piping and (3) purge gas/liquid oxygen interface tracking.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Marshall Space Flight Center Research Reports: 1985 NASA(ASEE Summer Faculty Fellowship Program; 18 p
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  • 25
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    A Monte Carlo study of Weibull reliability analysis for space shuttle main engine components (1986)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The incorporation of a number of additional capabilities into an existing Weibull analysis computer program and the results of Monte Carlo computer simulation study to evaluate the usefulness of the Weibull methods using samples with a very small number of failures and extensive censoring are discussed. Since the censoring mechanism inherent in the Space Shuttle Main Engine (SSME) data is hard to analyze, it was decided to use a random censoring model, generating censoring times from a uniform probability distribution. Some of the statistical techniques and computer programs that are used in the SSME Weibull analysis are described. The methods documented in were supplemented by adding computer calculations of approximate (using iteractive methods) confidence intervals for several parameters of interest. These calculations are based on a likelihood ratio statistic which is asymptotically a chisquared statistic with one degree of freedom. The assumptions built into the computer simulations are described. The simulation program and the techniques used in it are described there also. Simulation results are tabulated for various combinations of Weibull shape parameters and the numbers of failures in the samples.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Marshall Space Flight Center Research Reports: 1985 NASA(ASEE Summer Faculty Fellowship Program; 13 p
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  • 26
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    Rocket nozzle thermal shock tests in an arc heater facility (1986)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: A rocket motor nozzle thermal structural test technique that utilizes arc heated nitrogen to simulate a motor burn was developed. The technique was used to test four heavily instrumented full-scale Star 48 rocket motor 2D carbon/carbon segments at conditions simulating the predicted thermal-structural environment. All four nozzles survived the tests without catastrophic or other structural failures. The test technique demonstrated promise as a low cost, controllable alternative to rocket motor firing. The technique includes the capability of rapid termination in the event of failure, allowing post-test analysis.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA- Goddard Space Flight Center, Greenbelt, Md. Fourteenth Space Simulation Conference: Testing for a Permanent Presence in Space; p 27-41
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  • 27
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    Propulsion issues, options and trades (1986)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: Several different types of propulsion concepts are discussed: pulsed fission; continuous nuclear fission; chemical; and chemical boost with advanced nuclear fission. Some of the key characteristics of each type are provided, and typical concepts of each are shown.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Manned Mars Mission. Working Group Papers, V. 2, Sect. 5, App.; p 815-822]
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  • 28
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    PEGASUS: A multi-megawatt nuclear electric propulsion system (1986)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: A propulsion system (PEGASUS) consisting of an electric thruster driven by a multimegawatt nuclear power system is proposed for a manned Mars mission. Magnetoplasmadynamic and mercury-ion thrusters are considered, based on a mission profile containing a 510-day burn time (for a mission time of approximately 1000 days). Both thrusters are capable of meeting the mission parameters. Electric propulsion systems have significant advantages over chemical systems, because of high specific impulse, lower propellant requirements, and lower system mass. The power for the PEGASUS system is supplied by a boiling liquid-metal fast reactor. The power system consists of the reactor, reactor shielding, power conditioning subsystems, and heat rejection subsystems. It is capable of providing a maximum of 8.5 megawatts of electrical power of which 6 megawatts is needed for the thruster system, leaving 1.5 megawatts available for inflight mission applications.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Marshall Space Flight Center Manned Mars Mission. Working Group Papers, V. 2, Sect. 5, App.; p 769-786
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  • 29
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    Power system technologies for the manned Mars mission (1986)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: The high impulse of electric propulsion makes it an attractive option for manned interplanetary missions such as a manned mission to Mars. This option is, however, dependent on the availability of high energy sources for propulsive power in addition to that required for the manned interplanetary transit vehicle. Two power system technologies are presented: nuclear and solar. The ion thruster technology for the interplanetary transit vehicle is described for a typical mission. The power management and distribution system components required for such a mission must be further developed beyond today's technology status. High voltage-high current technology advancements must be achieved. These advancements are described. In addition, large amounts of waste heat must be rejected to the space environment by the thermal management system. Advanced concepts such as the liquid droplet radiator are discussed as possible candidates for the manned Mars mission. These thermal management technologies have great potential for significant weight reductions over the more conventional systems.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Marshall Space Flight Center Manned Mars Mission. Working Group Papers, V. 2, Sect. 5, App.; p 797-814
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  • 30
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    Management of SSME hardware life utilization (1986)
    In:  CASI
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    Publication Date: 2013-08-31
    Description: Statistical and probabilistic reliability methodologies were developed for the determination of hardware life limits for the Space Shuttle Main Engine (SSME). Both methodologies require that a mathematical reliability model of the engine (system) performance be developed as a function of the reliabilities of the components and parts. The system reliability model should be developed from the Failute Modes and Effects Analysis/Critical Items List. The statistical reliability methodology establishes hardware life limits directly from the failure distributions of the components and parts obtained from statistically-designed testing. The probabilistic reliability methodology establishes hardware life limits from a decision analysis methodology which incorporates the component/part reliabilities obtained from a probabilistic structural analysis, a calibrated maintenance program, inspection techniques, and fabrication procedures. Probilistic structural analysis is recommended as a tool to prioritize upgrading of the components and parts. The Weibull probability distribution is presently being investigated by NASA/MSFC to characterize the failure distribution of the SSME hardware from a limited data base of failures.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Marshall Space Flight Center Research Reports: 1986 NASA(ASEE Summer Faculty Fellowship Program; 30 p
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  • 31
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    Evaluation of moisture barrier coatings on carbon-phenolic SRM nozzle materials (1986)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: The carbon-phenolic composite ablative material used on the Solid Rocket Motor (SRM) nozzle is known to absorb moisture from the atmosphere. This could cause problems such as pocketing during firing. Several moisture barrier coatings were tested on the SRM nozzle material. Data are presented for six of the 12 coatings to be tested. The data were obtained from immersion of coated samples in an environmental chamber at 100 F and 100% relative humidity and by using a modified TGA (thermal gravimetric analysis) technique. The TGA technique involved allowing wet nitrogen (25 C, 80% relative humidity) to flow across a small sample at about 65 cu cm per minute while continually monitoring the weight increase. These preliminary results show Kel-F-800, a material supplied by 3M Corporation to be the better moisture barrier. A second task was to collect data on the relative absorption of water and kerosene into the carbon-phenolic SRM nozzle material. These data indicate that water absorbs into the nozzle material to a much greater extent than kerosene. Thus kerosene is the more likely solvent in which to make specific gravity measurements on the SRM nozzle material.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Marshall Space Flight Center Research Reports: 1986 NASA(ASEE Summer Faculty Fellowship Program; 19 p
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  • 32
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    A study of the possible effect of a thermal post cure on the mechanical properties of the 404 carbon-phenolic ring of the space shuttle solid rocket motor nozzle (1986)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: Erratic pockets of erosion have occurred on the inner perimeter of the 404 rocket nozzle ring during liftoff firing. It is thought that it may be caused by pockets of volatile matter entrapped during manufacture. A thermal post cure was suggested as a possible means of outgassing such pockets, if they in fact do exist. To confirm an outgassing during a post cure and to establish a working upper temperature limit, thermal gravimetric and differential calorimetric analyses were made on a number of samples from two 404 rings supplied by the manufacturer. Continuous weight loss was observed over the temperature range explored (750 F) indicating outgassing, and a strong exothermic reaction occurs beginning about 390 F. Thus, an upper post cure temperature of 350 F is recommended. To determine the possible effect of a post cure on physical properties, the following tests will be made on matched sets of cured and post cured material: x-radiography (internal structure), linear dimensions, weight, porosity, cross ply thermal expansion, drop and double notch shear strengths, and tensional strength in the ply direction.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Marshall Space Flight Center Research Reports: 1986 NASA(ASEE Summer Faculty Fellowship Program; 24 p
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  • 33
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    Investigation of potential driver modules and transmission lines for a high frequency power system on the space station (1986)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: The feasibility of using Series Resonant Inverter as the driver module for high frequency power system on the Space Station was assessed. The performance of the Series Resonant Inverter that was used in the testing of the single-phase, 2.0-kw resonant AC power system breadboard is summarized. The architecture is descirbed and the driver modules of the 5.0 kw AC power system breadboard are analyzed. An investigation of the various types of transmission lines is continued. Measurements of equivalent series resistor and inductor and equivalent parallel capacitors are presented. In particular, a simplified approach is utilized to describe the optimal transmission line.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Marshall Space Flight Center Research Reports: 1986 NASA(ASEE Summer Faculty Fellowship Program; 35 p
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  • 34
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    Vacuum chamber pressure effects on thrust measurements of low Reynolds number nozzles (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: Tests were conducted to investigate the effect of vacuum facility pressure on the performance of small thruster nozzles. Thrust measurements of two converging-diverging nozzles with an area ratio of 140 and an orifice plate flowing unheated nitrogen and hydrogen were taken over a wide range of vacuum facility pressures and nozzle throat Reynolds numbers. In the Reynolds number range of 2200 to 12,000 there was no discernable viscous effect on thrust below an ambient to total pressure ratio of 1000. In nearly all cases, flow separation occurred at a pressure ratio of about 1000. This was the upper limit for obtaining an accurate thrust measurement for a conical nozzle with an area ratio of 140.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Propulsion and Power (ISSN 0748-4658); 2; 385-389
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  • 35
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    Proven, long-life hydrogen/oxygen thrust chambers for space station propulsion (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: A 25 lb sub f hydrogen/oxygen thruster has been developed and proven as a viable candidate to meet the needs of the Space Station Program. Likewise, a 50 lb sub f hydrogen/oxygen thrust chamber has been developed and has demonstrated reliable, long-life expectancy at anticipated Space Station operating conditions. Both these thrust chambers were based on design criteria developed in previous thruster programs. Extensive thermal analysis and models were used to design the thrusters to achieve total impulse goals of 2 million lb sub f sec. Test data from each thruster are compared to the analytical predictions for the performance and heat transfer characteristics. Also, the results of thrust chamber life verification tests are presented.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Johns Hopkins Univ., The 1986 JANNAF Propulsion Meeting, Volume 1; p 547-564
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  • 36
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    Mixing of cryogenic fluids and predicted detonation properties for multiphase liquid oxygen and liquid hydrogen (1986)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Description: The mixing and explosion of LOX and LH sub 2 is a concern for Space Transportation System operations. To understand this problem, cryogenic mixing is experimentally studied by pouring 1,2,2 trichloro 2,1,1 trifluoroethane (Freon 113) into LN sub 2, LN sub 2 into LH sub 2, and LH sub 2 into into LN sub 2 in a 1 m by 15 cm cylindrical glass vessel. Data from these experiments is compared with previous studies and a hypothesis is advanced that LOX/LH sub 2 mixing will result in a complex, heterogenous, multiphase aggregation including LOX, SOX, LH sub 2, and VH sub 2 using the multiphase hypothesis. Visual and x ray observations of the process and mass measurements are reported.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Johns Hopkins Univ., The 1986 JANNAF Propulsion Meeting, volume 1; p 463-475
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  • 37
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    Resistojet- and arcjet-propulsion power processing (1986)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: Power-processing unit (PPU) designs for resistojet and arcjet propulsion systems were developed. Various PPU power converter and power circuit technologies were considered. Arcjet, resistojet and chemical propulsion system performance were compared. Significant propellant mass reductions are enabled with electric propulsion.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Johns Hopkins Univ., The 1986 JANNAF Propulsion Meeting, Volume 1; p 387-397
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  • 38
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    Analytical study of reusable flight-weight cryogenic propellant tank design (1986)
    In:  Other Sources
    Details
    Publication Date: 2011-08-19
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Spacecraft and Rockets (ISSN 0022-4650); 23; 149-157
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  • 39
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    Propulsion Working Group report (1986)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: Existing technology limits and performances, high payoff technologies, propulsion technologies, electric propulsion systems, and advanced bipropellant systems are outlined.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-Lewis Research Center, Spacecraft 2000; p 187-200
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  • 40
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    Solar dynamic power systems for space station (1986)
    In:  CASI
    Details
    Publication Date: 2013-08-31
    Description: The Parabolic Offset Linearly Actuated Reflector (POLAR) solar dynamic module was selected as the baseline design for a solar dynamic power system aboard the space station. The POLAR concept was chosen over other candidate designs after extensive trade studies. The primary advantages of the POLAR concept are the low mass moment of inertia of the module about the transverse boom and the compactness of the stowed module which enables packaging of two complete modules in the Shuttle orbiter payload bay. The fine pointing control system required for the solar dynamic module has been studied and initial results indicate that if disturbances from the station are allowed to back drive the rotary alpha joint, pointing errors caused by transient loads on the space station can be minimized. This would allow pointing controls to operate in bandwidths near system structural frequencies. The incorporation of the fine pointing control system into the solar dynamic module is fairly straightforward for the three strut concentrator support structure. However, results of structural analyses indicate that this three strut support is not optimum. Incorporation of a vernier pointing system into the proposed six strut support structure is being studied.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Langley Research Center NASA(DOD Control)Structures Interaction Technology, 1986; p 149-166
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  • 41
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    The noncavitating performance and life of a small vane-type positive displacement pump in liquid hydrogen (1986)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The low flow rate and high head rise requirements of hydrogen/oxygen auxiliary propulsion systems make the application of centrifugal pumps difficult. Positive displacement pumps are well-suited for these flow conditions, but little is known about their performance and life characteristics in liquid hydrogen. An experimental and analytical investigation was conducted to determine the performance and life characteristics of a vane-type, positive displacement pump. In the experimental part of this effort, mass flow rate and shaft torque were determined as functions of shaft speed and pump pressure rise. Since liquid hydrogen offers little lubrication in a rubbing situation, pump life is an issue. During the life test, the pump was operated intermittently for 10 hr at the steady-state point of 0.074 lbm/sec (0.03 kg/sec) flow rate, 3000 psid (2.07 MPa) pressure rise, and 9000 rpm (938 rad/sec) shaft speed. Pump performance was monitored during the life test series and the results indicated no loss in performance. Material loss from the vanes was recorded and wear of the other components was documented. In the analytical part of this effort, a comprehensive pump performance analysis computer code, developed in-house, was used to predict pump performance. The results of the experimental investigation are presented and compared with the results of the analysis. Results of the life test are also presented.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1438
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  • 42
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    Performance characteristics of ring-cusp thrusters with xenon propellant (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The performance characteristics and operating envelope of several 30-cm ring-cusp ion thrusters with xenon propellant were investigated. Results indicate a strong performance dependence on the discharge chamber boundary magnetic fields and resultant distribution of electron currents. Significant improvements in discharge performance over J-series divergent-field thrusters were achieved for large throttling ranges, which translate into reduced cathode emission currents and reduced power dissipation which should be of significant benefit for operation at thruster power levels in excess of 10 kW. Mass spectrometer of the ion beam was documented for both the ring-cusp and J-series thrusters with xenon propellant for determination of overall thruster efficiency, and lifetime. Based on the lower centerline values of doubly charged ions in the ion beam and the lower operating discharge voltage, the screen grid erosion rate of the ring-cusp thruster is expected to be lower than the divergent-field J-series thruster by a factor of 2.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1392
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  • 43
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    Space power - Emerging opportunities (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: NASA programs directed towards the development of technologies to meet the cost-effective energy needs of future space missions are described. Consideration is given to the space photovoltaic program, which was developed along two paths: one leading to high-performance ultralight weight solar arrays, the other to high output arrays. The space power materials and energy storage technology are discussed, together with the developmental aspects of an advanced solar dynamic power system and its subsystems. Special attention is given to the Nasa SP-100 Advanced Technology Project and the free-piston Stirling engine technology for nuclear power application.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: IAF PAPER 86-152
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  • 44
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    Electrodynamic tethers for energy conversion (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Conductive tethers have been proposed as a new method for converting orbital mechanical energy into electrical power for use on-board a satellite (generator mode) or conversely (motor mode) as a method of providing electric propulsion using electrical energy from the satellite. The operating characteristics of such systems are functionally dependent on orbit altitude and inclination. Effects of these relationships are examined to determine acceptable regions of application. To identify system design considerations, a specific set of system performance goals and requirements are selected. The case selected is for a 25 kW auxiliary power system for use on Space Station. Appropriate system design considerations are developed, and the resulting system is described.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 45
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    Propulsion recommendations for space station free flying platforms (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Propulsion system candidates have been defined for Space Station free flying platforms for the purpose of comparison and to understand the impact of the various mission requirements on the candidate designs. Recommendations for propulsion for each of the various platforms are given.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Johns Hopkins Univ., The 1986 JANNAF Propulsion Meeting, Volume 1; p 515-523
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  • 46
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    Experimental aerodynamic evaluation of the SSME two-duct, hot gas manifold (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An investigation was performed on an improved Space Shuttle Main Engine (SSME) hot gas manifold (HGM) design using a full scale HGM model in cold-flow, air-blowdown tests. The new HGM design replaces the current three transfer ducts with two enlarged elliptical transfer ducts, faired duct inlets, and enlarged fuel bowl. Extensive experimental results have been obtained to verify previous results showing that the two-duct design improved the HGM internal flow path by eliminating separated flow regions, reducing high local velocities, and providing more uniform flow.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Johns Hopkins Univ., The 1986 JANNAF Propulsion Meeting, Volume 1; p 349-360
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  • 47
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    Concepts for space maintenance of OTV engines (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Concepts for space maintainability of the Orbital Transfer Vehicle (OTV) engines are examined. An engine design is developed which is driven by space maintenance requirements and by a failure modes and effects analysis (FMEA). Modularity within the engine is shown to offer cost benefits and improved space maintenance capabilities. Space-operable disconnects are conceptualized for both engine change-out and for module replacement. A preliminary space maintenance plan is developed around a controls and condition monitoring system using advanced sensors, controls, and condition monitoring concepts.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Johns Hopkins Univ., The 1986 JANNAF Propulsion Meeting, Volume 1; p 99-110
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  • 48
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    Block 2 Solid Rocket Motor (SRM) conceptual design study. Volume 1: Appendices (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The design studies task implements the primary objective of developing a Block II Solid Rocket Motor (SRM) design offering improved flight safety and reliability. The SRM literature was reviewed. The Preliminary Development and Validation Plan is presented.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-179051 , NAS 1.26:179051 , TR-PL-12126-VOL-1
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  • 49
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    Block 2 SRM conceptual design studies. Volume 1, Book 1: Appendix A: CEI specification CPW1-1900, Appendix B: Composite motor case (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The preliminary Contract End Item (CEI) specification presents the performance, design, and verification requirements for Morton Thiokol's Space Shuttle Block 2 Solid Rocket Motor (SRM).
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-179049 , NAS 1.26:179049 , MTI-PUB-87354-VOL-1-BK-1-APP
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  • 50
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    Block 2 SRM conceptual design studies. Volume 1, Book 2: Preliminary development and verification plan (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Activities that will be conducted in support of the development and verification of the Block 2 Solid Rocket Motor (SRM) are described. Development includes design, fabrication, processing, and testing activities in which the results are fed back into the project. Verification includes analytical and test activities which demonstrate SRM component/subassembly/assembly capability to perform its intended function. The management organization responsible for formulating and implementing the verification program is introduced. It also identifies the controls which will monitor and track the verification program. Integral with the design and certification of the SRM are other pieces of equipment used in transportation, handling, and testing which influence the reliability and maintainability of the SRM configuration. The certification of this equipment is also discussed.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-179054 , NAS 1.26:179054 , MTI-PUB-87354-VOL-1-BK-2
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  • 51
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    Block 2 SRM conceptual design studies. Volume 1, Book 1: Conceptual design package (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The conceptual design studies of a Block 2 Solid Rocket Motor (SRM) require the elimination of asbestos-filled insulation and was open to alternate designs, such as case changes, different propellants, modified burn rate - to improve reliability and performance. Limitations were placed on SRM changes such that the outside geometry should not impact the physical interfaces with other Space Shuttle elements and should have minimum changes to the aerodynamic and dynamic characteristics of the Space Shuttle vehicle. Previous Space Shuttle SRM experience was assessed and new design concepts combined to define a valid approach to assured flight success and economic operation of the STS. Trade studies, preliminary designs, analyses, plans, and cost estimates are documented.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-179048 , NAS 1.26:179048 , MTI-PUB-87354-VOL-1-BK-1
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  • 52
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    Space Transportation Booster Engine (STBE) configuration study (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The overall objective of this Space Transportation Booster Engine (STBE) study is to identify candidate engine configurations which enhance vehicle performance and provide operational flexibility at low cost. The specific objectives are as follows: (1) to identify and evaluate candidate LOX/HC engine configurations for the Advanced Space Transportation System for an early 1995 IOC and a late 2000 IOC; (2) to select one optimum engine for each time period; 3) to prepare a conceptual design for each configuration; (4) to develop a technology plan for the 2000 IOC engine; and, (5) to prepare preliminary programmatic planning and analysis for the 1995 IOC engine.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-178941 , NAS 1.26:178941 , RI/RD-BD86-149 , QR-2
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  • 53
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    Advanced propulsion for polar-orbiting and coorbiting free flyers (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Advanced propulsion systems for polar-orbiting and coorbiting free flyers were investigated. Resistojet, arcjet, ion, magnetoplasmadynamic and chemical-bipropellant nitrogen tetroxide/monomethyl hydrazine (NTO/MMH) propulsion systems were compared to the baseline free-flyer hydrazine (N2H4)-propulsion-system performance. Advanced resistojet, arcjet, ion, and NTO/MMH propulsion systems enable significant propellant-mass savings over the baseline N2H4-propulsion system. Using free-flyer mission requirements from the Langley Research Center Mission-Data Base, detailed propulsion requirements for over thirty free-flyer missions were analyzed. The Polar-Platform trip-time constraints may preclude using a low-thrust electric-propulsion system. Electric propulsion will, however, allow a significant coorbiter propellant-mass reduction. Frequent servicing and nodal-regression effects on the coorbiting free-flyer's orbit increase the required mission velocity change and propellant mass. For many coorbiter missions high-specific-impulse resistojet-, arcjet- and ion-propulsion systems allow substantial life-cycle propellant-mass savings.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1564
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  • 54
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    Development of an arcjet nuclear electric propulsion system for a 1993 flight demonstration (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The design and performance of an arcjet NEP spacecraft, suitable for use in the Space Nuclear Power System (SNPS) Reference Mission, are outlined. Several arcjet technology levels were considered in this study, and the vehicle design was based on an 30 kW ammonia arcjet system operating at an Isp of 1000 s and an efficiency of 45 percent. The arcjet/gimbal system, PPU and propellant feed-system are described. A 100 kWe SNPS was assumed and the spacecraft mass was baselined at 5500 kg excluding the propellant feed system. A radiation/arcjet efflux diagnostics package was included in the performance analysis. This spacecraft, launched from Kennedy, can perform a 50 degree inclination change and reach a final orbit of GEO with a 180 day trip time providing a six month active load for the SNPS. Advanced ammonia and hydrogen systems were examined for precursor SDI platform applications.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1510
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  • 55
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    Bipropellant rocket exhaust plume analysis on the Galileo spacecraft (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: This paper describes efforts to quantify the contaminant flow field produced by 10 N thrust bipropellant rocket engines used on the Galileo spacecraft. The prediction of the composition of the rocket exhaust by conventional techniques is found to be inadequate to explain experimental observations of contaminant deposition on moderately cold (200 K) surfaces. It is hypothesized that low volatility contaminants are formed by chemical reactions which occur on the surfaces. The flow field calculations performed using the direct simulation Monte Carlo method give the expected result that the use of line-of-sight plume shields may have very little effect on the flux of vapor phase contaminant species to a surface, especially if the plume shields are located so close to the engine that the interaction of the plume with the shield is in the transition flow regime. It is shown that significant variations in the exhaust plume composition caused by nonequilibrium effects in the flow field lead to very low concentrations of species which have high molecular weights in the more rarefied regions of the flow field. Recommendations for the design of spacecraft plume shields and further work are made.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1488
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  • 56
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    Erosive burning research (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A status report is given on the results for the completed tests in a series of motor firings being carried out to measure the effects of the parameters that are considered to most strongly influence the scaling to larger rocket motor sizes of the transition to/or threshold conditions for erosive burning rate augmentation. Propellant burning rates at locations along the axis of the test motors are measured with a newly developed plasma capacitance gauge technique. The measured results are compared with erosive-burning predictions from a supporting ballistics analysis. The completed motor firings have successfully demonstrated response to the designed test variables. The trends with varying propellant burning rate, chamber pressure, and mass flow rate are consistent with existing results, but no pronounced effect of surface roughness has been observed. Rather, the influence of propellant oxidizer particle size on erosive burning is through its effect on the base, no-corssflow burning rate.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1449
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  • 57
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    Effects of geometry and materials on low cycle fatigue life of turbine blades in LOX/hydrogen rocket engines (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: This paper presents the results of an advanced turbine blade test program aimed at improving turbine blade low cycle fatigue (LCF) life. A total of 21 blades were tested in a blade thermal tester. The blades were made of MAR-M-246(Hf)DS and PWA-1480SC in six different geometries. The test results show that the PWA-1480SC material improved life by a factor of 1.7 to 3.0 over the current MAR-M-246(Hf)DS. The geometry changes yielded life improvements as high as 20 times the baseline blade made of PWA-1480SC and 34 times the baseline MAR-M-246DS blade.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1443
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  • 58
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    A review of the multimegawatt MPD thruster and current mission applications (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The magnetoplasmadynamic (MPD) thruster is potentially capable of providing 20-200 N of steady-state thrust, at 1000-20,000 s specific impulse while consuming a variety of propellants and megawatts of DC power. The specific impulse and power capacity put the MPD thruster in a class separate from other electric propulsion engines. Several types of missions are enhanced by the unique capabilities of a nuclear-electric driven MPD thruster and these include earth-orbit raising, cis-lunar transportation, and planetary exploration. This paper describes current MPD thruster technology, and several missions typical of the above cited categories.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1437
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  • 59
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    Numerical analysis of the three-dimensional flow in the main injector assembly of the Space Shuttle Main Engine (1986)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The Main Injector Assembly of the Space Shuttle Main Engine supplies the propellants to the Main Combustion Chamber through a large number of vertical injector elements. The gas flow around these elements (LOX posts) is three-dimensional, turbulent and compressible. This paper presents the results of numerical modeling of the hot and cool sections of the Main Injector Assembly and shows that the shields on the outermost row of injector elements affect the flow and the pressure drop in the hot section significantly. It is suggested that appropriate experiments are undertaken to verify the predictions.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1422
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    A long-life 50 lbf H2/O2 thruster for Space Station auxiliary propulsion (1986)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: In preparation for the development of a manned Space Station, the National Aeronautics and Space Administration (NASA) is conducting a program to develop technology related to on-board Auxiliary Propulsion Systems. To develop the required thruster technology to support the Space Station project, the NASA Lewis Research Center has sponsored a development program based on a unique 'reverse flow' concept where the fuel is injected 'backwards' in the chamber to cool the spherical combustor wall. This combustor was based on previous developments at the 50-lbf, 1000-lbf, and 1500-lbf thrust levels. This paper describes the design and test program carried out to demonstrate a new 50-lbf thruster, the design which was based on this previous technology. Included are the test results for the initial mixture ratio 4 thruster which can operate with uncooled Cres (stainless steel) combustor walls. In addition, the effort to operate a thruster redesigned for operation at a mixture ratio of eight for potential integration with the life support system is described.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1404
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  • 61
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    Thrust imbalance of solid rocket motor pairs on Space Shuttle flights (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: This analysis extends the investigation presented at the 17th Joint Propulsion Conference in 1981 to include fifteen sets of Space Shuttle flight data. The previous report dealt only with static test data and the first flight pair. The objective is to compare the authors' previous theoretical analysis of thrust imbalance with actual Space Shuttle performance. The theoretical prediction method, which involves a Monte Carlo technique, is reviewed briefly as are salient features of the flight instrumentation system and the statistical analysis. A scheme for smoothing flight data is discussed. The effects of changes in design parameters are discussed with special emphasis on the filament wound motor case being developed to replace the steel case. Good agreement between the predictions and the flight data is demonstrated.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1375
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  • 62
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    SSME thrust chamber modeling with Navier Stokes equations (1986)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The capability of predicting two-dimensional, compressible and reacting flow in the combustion chamber and nozzle of the Space Shuttle Main Engine (SSME) is demonstrated. A nonorthogonal body fitted coordinate system has been used to represent the combustor and nozzle geometry. The Navier-Stokes equations are solved for the entire thrust chamber with the k-epsilon turbulence model accounting for compressibility and large pressure gradients effects. Results of the computational test cases reveal all expected features of the transonic nozzle flows including location of sonic line, internal shock and boundary layer build-up. Calculated performance parameters such as thrust, flow rate, and specific impulse are also in reasonble agreement with available data. The results show promising potential of solving full Navier-Stokes equations with heat transfer and two-phase combustion in truly comprehensive modeling of rocket engines.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1517
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  • 63
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    The design and operating characteristics of a 30-kW thermal arcjet engine for space propulsion (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Attention is given to the design features of a radiation-cooled, 30-kW thermal arcjet thruster, whose laboratory tests have yielded specific impulses of up to 935 sec at 36-44 percent thrust efficiency, together with a cumulative lifetime of over 400 hours. All materials used, including seals, can sustain operation at temperatures sufficiently elevated to require the radiation of all waste heat. This electric propulsion system is ideally suited to missions such as the Space-Based Radar. A detailed consideration is conducted for the thruster's seals, which are the most critical element of the design.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1508
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  • 64
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    Diffuser investigation for advanced multi-megawatt MPD engine development (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Steady-state testing of advanced, multimegawatt MPD engines requires vacuum facilities with pumping speeds which are not presently available. Gas dynamic diffusers have been proposed as a possible solution to this problem. An analytical investigation into the feasibility of using a diffuser for pumping the MPD engine exhaust is presented here. This analysis uses a detailed equation of state based on the partition function for the high temperature argon exhaust. On the basis of the electron-ion momentum exchange collision frequency in the exhaust plasma, it was concluded that the diffuser gas dynamics could be modeled to a first approximation with ordinary continuum equations. Calculations of the stagnation pressure in the diffuser, downstream of a strong normal shock, yielded pressures on the order of 10 torr suggesting that the diffuser is feasible for pumping the MPD engine exhaust.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1436
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  • 65
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    Performance and endurance tests of a multipropellant resistojet for space station auxiliary propulsion (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: This paper presents the results of an effort to demonstrate the technology readiness of a long-life multipropellant resistojet for space station auxiliary propulsion. Experiments were performed to evaluate the compatibility of grain-stabilized platinum tubes at temperatures up to 1400 deg C in environments of CO2, CH4, NH3, H2O, and H2. All samples tested showed extrapolated lifetimes in excess of 10,000 hr based on 10 percent mass loss as end-of-life. However, samples tested in an ammonia atmosphere at 1400 deg C showed severe pitting, which raised concerns about the compatibility of grain-stabilized platinum with ammonia-containing atmospheres. Additional tests showed that reducing the metal temperature to about 900 deg C (+ or - 100 deg C) significantly reduced this adverse effect.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1435
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  • 66
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    A xenon ion propulsion module for enhanced spacecraft capability (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Design, development and preliminary testing of a two-engine xenon ion propulsion module for early flight evaluation is described. Extensive use is made of flight spare propellant system components and also engineering model 30-cm ion engines and assemblies originally developed for the Solar Electric Propulsion System program. Significant design features include a redundant propellant feed system that incorporates a novel gas pulse assembly for rapid and completely reliable engine startup, a central neutralizer subsystem with dual neutralizers for redundancy, and major ion engine performance improvements resulting in a nearly doubling of the 30-cm engine thrust in addition to operating on xenon rather than mercury propellant. At a module input power of 10.0 kw, maximum thrust and specific impulse are projected to be 0.4 N and 3,500 sec. respectively, for a total module efficiency of 67 percent. Total mass of the xenon ion module, including the propellant tank, is only 70.2 kg. The technical approaches taken towards developing and integrating the subsystems comprising this ion propulsion module are presented and discussed in detail.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1393
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  • 67
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    Characterization of real gas properties for space shuttle main engine fuel turbine and performance calculations (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Real thermodynamic and transport properties of hydrogen, steam, the SSME mixture, and air are developed. The SSME mixture properties are needed for the analysis of the space shuttle main engine fuel turbine. The mixture conditions for the gases, except air, are presented graphically over a temperature range from 800 to 1200 K, and a pressure range from 1 to 500 atm. Air properties are given over a temperature range of 320 to 500 K, which are within the bounds of the thermodynamics programs used, in order to provide mixture data which is more easily checked (than H2/H2O). The real gas property variation of the SSME mixture is quantified. Polynomial expressions, needed for future computer analysis, for viscosity, Prandtl number, and thermal conductivity are given for the H2/H2O SSME fuel turbine mixture at a pressure of 305 atm over a range of temperatures from 950 to 1140 K. These conditions are representative of the SSME turbine operation. Performance calculations are presented for the space shuttle main engine (SSME) fuel turbine. The calculations use the air equivalent concept. Progress towards obtaining the capability to evaluate the performance of the SSME fuel turbine, with the H2/H2O mixture, is described.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-175066 , E-2938 , NAS 1.26:175066
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    Orbit Transfer Rocket Engine Technology Program: Advanced engine study, task D.1/D.3 (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Concepts for space maintainability of OTV engines were examined. An engine design was developed which was driven by space maintenance requirements and by a failure mode and effects (FME) analysis. Modularity within the engine was shown to offer cost benefits and improved space maintenance capabilities. Space operable disconnects were conceptualized for both engine change-out and for module replacement. Through FME mitigation the modules were conceptualized to contain the least reliable and most often replaced engine components. A preliminary space maintenance plan was developed around a controls and condition monitoring system using advanced sensors, controls, and condition monitoring concepts. A complete engine layout was prepared satisfying current vehicle requirements and utilizing projected component advanced technologies. A technology plan for developing the required technology was assembled.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-175084 , NAS 1.26:175084 , RI/RD86-116
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  • 69
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    A bibliography of electrothermal thruster technology, 1984 (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Electrothermal propulsion concepts are briefly discussed as an introduction to a bibliography and author index. Nearly 700 citations are given for resistojets, thermal arcjets, pulsed electrothermal thrusters, microwave heated devices, solar thermal thrusters, and laser thermal thrusters.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-86998 , E-2536 , NAS 1.15:86998
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  • 70
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    Technology for Bayton-cycle powerplants using solar and nuclear energy (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Brayton cycle gas turbines have the potential to use either solar heat or nuclear reactors for generating from tens of kilowatts to tens of megawatts of power in space, all this from a single technology for the power generating system. Their development for solar energy dynamic power generation for the space station could be the first step in an evolution of such powerplants for a very wide range of applications. At the low power level of only 10 kWe, a power generating system has already demonstrated overall efficiency of 0.29 and operated 38 000 hr. Tests of improved components show that these components would raise that efficiency to 0.32, a value twice that demonstrated by any alternate concept. Because of this high efficiency, solar Brayton cycle power generators offer the potential to increase power per unit of solar collector area to levels exceeding four times that from photovoltaic powerplants using present technology for silicon solar cells. The technologies for solar mirrors and heat receivers are reviewed and assessed. This Brayton technology for solar powerplants is equally suitable for use with the nuclear reactors. The available long time creep data on the tantalum alloy ASTAR-811C show that such Brayton cycles can evolve to cycle peak temperatures of 1500 K (2240 F). And this same technology can be extended to generate 10 to 100 MW in space by exploiting existing technology for terrestrial gas turbines in the fields of both aircraft propulsion and stationary power generation.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TP-2558 , E-2761 , NAS 1.60:2558
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  • 71
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    Selecting hydrocarbon rocket propulsion technology (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Past studies have shown that the dry weight of future earth-to-orbit vehicles can be reduced by the combined use of hydrogen and hydrocarbon propulsion compared to all-hydrogen propulsion. This paper shows that the use of certain hydrocarbon engines with hydrogen engines produces the lowest vehicle dry mass. These hydrocarbon engines use propane or RP-1 fuel, hydrogen cooling, and hydrogen-rich gas generators. Integration of the hydrogen and hydrocarbon nozzles is also beneficial.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: IAF PAPER 86-167
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  • 72
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    Advances in high chamber pressure propulsion (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: NASA has been involved in the development of improved high thrust booster rocket engines to meet the propulsion requirements of launch vehicles such as the Space Shuttle. Solutions that NASA/Marshall Space Flight Center pursued to accomplish the high performance, long life goals set for SSME are discussed. In addition, currently projected requirements for liquid rocket engines have identified liquid oxygen/hydrocarbon-fueled engines for booster application in the near future. These advanced hydrocarbon-fueled engines will require improvements in performance and life to be suitable for their projected missions. Raising chamber pressure to increase performance and reduce engine envelope are the key objectives in hydrocarbon-fueled engine technology. This paper traces the history of advances in high pressure rocket engine systems and the challenges it presents.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: IAF PAPER 86-166
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  • 73
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    Space station auxiliary thrust chamber technology (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A program to design, fabricate and test a 50 lb sub f (222 N) thruster was undertaken (Contract NAS 3-24656) to demonstrate the applicability of the reverse flow concept as an item of auxiliary propulsion for the space station. The thruster was to operate at a mixture ratio (O/F) of 4, be capable of operating for 2 million lb sub f- seconds (8.896 million N-seconds) impulse with a chamber pressure of 75 psia (52 N/square cm) and a nozzle area ratio of 40. Superimposed was also the objective of operating with a strainless steel spherical combustion chamber, which limited the wall temperature to 1700 F (1200 K), an objective specific impulse of 400 lb sub f sec/lbm (3923 N-seconds/Kg), and a demonstration of 500,000 lb sub f-seconds (2,224,000 N-seconds) of impulse. The demonstration of these objectives required a number of design iterations which eventually culminated in a very successful 1000 second demonstration, almost immediately followed by a changed program objective imposed to redesign and demonstrate at a mixture ratio (O/F) of 8. This change was made and more then 250,000 lb sub f seconds (1,112,000 N-seconds) of impulse was successfully demonstrated at a mixture ratio of 8. This document contains a description of the effort conducted during the program to design and demonstrate the thrusters involved.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-179552 , NAS 1.26:179552 , REPT-8911-950001
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  • 74
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    Space station experiment definition: Advanced power system test bed (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A conceptual design for an advanced photovoltaic power system test bed was provided and the requirements for advanced photovoltaic power system experiments better defined. Results of this study will be used in the design efforts conducted in phase B and phase C/D of the space station program so that the test bed capabilities will be responsive to user needs. Critical PV and energy storage technologies were identified and inputs were received from the idustry (government and commercial, U.S. and international) which identified experimental requirements. These inputs were used to develop a number of different conceptual designs. Pros and cons of each were discussed and a strawman candidate identified. A preliminary evolutionary plan, which included necessary precursor activities, was established and cost estimates presented which would allow for a successful implementation to the space station in the 1994 time frame.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-179502 , NAS 1.26:179502 , WDL-TR10939
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    High voltage characteristics of the electrodynamic tether and the generation of power and propulsion (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The Tethered Satellite System (TSS) will deploy and retrieve a satellite from the Space Shuttle orbiter with a tether of up to 100 km in length attached between the satellite and the orbiter. The characteristics of the TSS which are related to high voltages, electrical currents, energy storage, power, and the generation of plasma waves are described. A number of specific features of the tether system of importance in assessing the operational characteristics of the electrodynamic TSS are identified.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-178949 , NAS 1.26:178949
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    PEGASUS - A multi-megawatt nuclear electric propulsion system (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A propulsion system (The PEGASUS Drive) consisting of a magnetoplasmadynamic (MPD) thruster driven by a multimegawatt nuclear power system is proposed as the propulsion system for a manned Mars mission. The propulsion system described is based on a mission profile containing a 510-day burn time (for a mission time of approximately 1000 days). Electric propulsion systems have significant advantages over chemical systems, because of high specific impulse, lower propellant requirements, and lower system mass. The thermal power for the PEGASUS Drive is supplied by a boiling liquid-metal fast reactor. The system consists of the reactor, reactor shielding, power conditioning, heat rejection, and MPD thruster subsystems. It is capable of providing a maximum of 8.5 megawatts of electrical power of which 6 megawatts is needed for the thruster system, 1.5 megawatts is available for spacecraft system operations and inflight mission applications, leaving the balance for power system operation.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1583
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    Space Shuttle Main Engine fuel preburner augmented spark igniter shutdown detonations (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Detonations were experienced in the Space Shuttle Main Engine fuel preburner (FPB) augmented spark igniter (ASI) during engine cutoff. Several of these resulted in over pressures sufficient to damage the FPB ASI oxidizer system. The detonations initiated in the FPB ASI oxidizer line when residual oxidizer (oxygen) in the line mixed with backflowing fuel (hydrogen) and detonated. This paper reviews the damage history to the FPB ASI oxidizer system, an engineering assessment of the problem cause, a verification of the mechanisms, the hazards associated with the detonations, and the solution implemented.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1445
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  • 78
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    A 10,000 hour life multipropellant engine for Space Station applications (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A review of the design background and operating objectives of a multipropellant resistojet is presented. An engine has been designed to operate with carbon dioxide, methane, water, hydrazine decomposition products and hydrogen. Design performance has been constrained to ensure a 10,000-hour life. The engine, constructed primarily of grain stabilized platinum, is to operate at temperatures up to 1400 C. General performance guidelines, design and fabrication methods are reported.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1403
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  • 79
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    Thermofluid analysis of the SSME preburner using a gas-gas diffusion model for oxygen and hydrogen combustion at supercritical pressures (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The paper discusses the thermofluid analysis of the Space Shuttle Main Engine (SSME) fuelside preburner. The governing equations have been solved numerically to predict flow, heat transfer, mixing, and combustion. A two-fluid approach is adopted in which oxygen is regarded as one fluid and hydrogen is regarded as the other fluid. The chemical kinetics is assumed to be very fast so that combustion is primarily controlled by the rate of mixing between oxygen and hydrogen. The preburner pressure is much greater than the critical pressures of oxygen and hydrogen; hence, a gas-gas diffusion model (rather than an evaporation model) has been developed to compute the rate of interphase mixing. Empirical correlations have been incorporated to account for the effect of slip on the interphase exchange. A sensitivity study has been performed with various model parameters. It is observed that the model can predict possibility of incomplete combustion and local regions of high temperatures under steady operating conditions. Some of these anomalies have been observed in actual tests, and the numerical model is useful for understanding possible causes and remedies. At least some measurements are needed for quantitative verification of the model.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1425
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    Computational simulation of liquid rocket injector anomalies (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A computer model has been developed to analyze the three-dimensional two-phase reactive flows in liquid fueled rocket combustors. The model is designed to study the influence of liquid propellant injection nonuniformities on the flow pattern, combustion and heat transfer within the combustor. The Eulerian-Lagrangian approach for simulating polidisperse spray flow, evaporation and combustion has been used. Full coupling between the phases is accounted for. A nonorthogonal, body fitted coordinate system along with a conservative control volume formulation is employed. The physical models built into the model include a kappa-epsilon turbulence model, a two-step chemical reaction, and the six-flux radiation model. Semiempirical models are used to describe all interphase coupling terms as well as chemical reaction rates. The purpose of this study was to demonstrate an analytical capability to predict the effects of reactant injection nonuniformities (injection anomalies) on combustion and heat transfer within the rocket combustion chamber. The results show promising application of the model to comprehensive modeling of liquid propellant rocket engines.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1424
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    Solar dynamic power for the Space Station (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: This paper describes a computer code which provides a significant advance in the systems analysis capabilities of solar dynamic power modules. While the code can be used to advantage in the preliminary analysis of terrestrial solar dynamic modules its real value lies in the adaptions which make it particularly useful for the conceptualization of optimized power modules for space applications. In particular, as illustrated in the paper, the code can be used to establish optimum values of concentrator diameter, concentrator surface roughness, concentrator rim angle and receiver aperture corresponding to the main heat cycle options - Organic Rankine and Brayton - and for certain receiver design options. The code can also be used to establish system sizing margins to account for the loss of reflectivity in orbit or the seasonal variation of insolation. By the simulation of the interactions among the major components of a solar dynamic module and through simplified formulations of the major thermal-optic-thermodynamic interactions the code adds a powerful, efficient and economic analytical tool to the repertory of techniques available for the design of advanced space power systems.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1299
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    Effect of subcooling on the on-orbit pressurization rate of cryogenic propellant tankage (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The SOLA-ECLIPSE code is being developed to enable prediction of the behavior of cryogenic propellants in spacecraft tankage. A brief description of the formulations used for modeling heat transfer and for determining thermodynamic state is presented. Code performance is verified through comparison to experimental data for the self-pressurization of scale model liquid hydrogen tanks. SOLA-ECLIPSE is used to examine the effect of initial subcooling of the liquid phase on the self-pressurization rate of an on-orbit full scale liquid hydrogen tank typical for a chemical propulsion Orbital Transfer Vehicle. The computational predictions show that even small amounts of subcooling will significantly decrease the self-pressurization rate. Further, if the cooling is provided by a Thermodynamic Vent System, it is concluded that small levels of subcooling will maximize propellant conservation.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-1253
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  • 83
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    Numerical analysis of SSME preburner injector atomization and combustion processes (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The coaxial spray injection and combustion flowfields of a Space Shuttle Main Engine preburner injector element have been analyzed using a three-phase numerical code. The processes of atomization, evaporation, secondary droplet breakup, and multispecies chemistry, as well as turbulent diffusion, are included. The model produced realistic pictures of the complex internal flowfield, including liquid jet length, spray shape, flame-zone size and characteristics, and predicted temperatures that seem to be in agreement with test data envelopes. It predicted an external group combustion type of flame. Salient combustion and mixing features are discussed and sources of uncertainty are pointed out for future studies.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-0454
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  • 84
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    Solar dynamic systems for spacecraft power applications (1986)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Results are presented of a parametric study of the potential for using solar dynamic (SD) power supply systems on deep space probes. The SD systems would consist of a parabolic concentrator to focus solar energy on a thermal receiver for conversion by Brayton, organic Rankine or Stirling engines. The net thermal power and efficiencies available from each of the types of conversion devices were analyzed for a power requirement of 0.5 kWe. Examinations were also carried out of the optical, thermodynamic, materials and size limitations of the devices. The subsystem drivers were found to be the quality of concentrator reflectance and the system temperature level. Lower temperature systems are preferred for farther distances from the sun, mainly due to the required concentrator area. The SD system could be used out to 6 A.U. in optimal conditions. It is concluded that Brayon and Stirling engines have the best chances for further development, and that Rankine systems have already been optimized. Further evaluations are dependent on the definition of specific mission requirements.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 86-0382
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  • 85
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    Propulsion apparatus and method using boil-off gas from a cryogenic liquid (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A propulsion system and method are disclosed for controlling the attitude and drag of a space vehicle. A helium dewar contains liquid helium which cools an experiment package. The helium is heated or vented to keep the temperature between 1.5 and 1.7 degrees K to maintain adequate helium boil-off gas as a propellant without adversely affecting the experiment package which is contained in the helium dewar for protection from solar heating. The apparatus includes auxiliary heater and temperature sensor for controlling the temperature of the helium. The boil-off gas propellant is delivered to thruster modules to control vehicle attutude and compensate for drag.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 86
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    A data base and analysis program for shuttle main engine dynamic pressure measurements (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A dynamic pressure data base management system is described for measurements obtained from space shuttle main engine (SSME) hot firing tests. The data were provided in terms of engine power level and rms pressure time histories, and power spectra of the dynamic pressure measurements at selected times during each test. Test measurements and engine locations are defined along with a discussion of data acquisition and reduction procedures. A description of the data base management analysis system is provided and subroutines developed for obtaining selected measurement means, variances, ranges and other statistics of interest are discussed. A summary of pressure spectra obtained at SSME rated power level is provided for reference. Application of the singular value decomposition technique to spectrum interpolation is discussed and isoplots of interpolated spectra are presented to indicate measurement trends with engine power level. Program listings of the data base management and spectrum interpolation software are given. Appendices are included to document all data base measurements.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-178808 , NAS 1.26:178808 , REPT-66338-01
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  • 87
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    Evaluation of a double Gimbal IPACS design (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The suitability of various integrated power/attitude control systems (IPACS) rotor materials was analyzed. Three materials were investigated: (1) 6A1-4V-Titanium (the current IPACS rotor material); (2) B120 VCA Titanium; and (3) Custom 455 stainless steel. The preliminary linear vibration analysis was updated to include the weights and stiffnesses of the gimbals design. A belleville washer spring preload mechanism was designed to replace the existing helical spring and interference fit preload mechanism.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-178040 , NAS 1.26:178040
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  • 88
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    Engineering evaluation of SSME dynamic data from engine tests and SSV flights (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An engineering evaluation of dynamic data from SSME hot firing tests and SSV flights is summarized. The basic objective of the study is to provide analyses of vibration, strain and dynamic pressure measurements in support of MSFC performance and reliability improvement programs. A brief description of the SSME test program is given and a typical test evaluation cycle reviewed. Data banks generated to characterize SSME component dynamic characteristics are described and statistical analyses performed on these data base measurements are discussed. Analytical models applied to define the dynamic behavior of SSME components (such as turbopump bearing elements and the flight accelerometer safety cut-off system) are also summarized. Appendices are included to illustrate some typical tasks performed under this study.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-178807 , NAS 1.26:178807 , TR-64058-03
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  • 89
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    Computational analysis of the SSME fuel preburner flow (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A computational fluid dynamics model which simulates the steady state operation of the SSME fuel preburner is developed. Specifically, the model will be used to quantify the flow factors which cause local hot spots in the fuel preburner in order to recommend experiments whereby the control of undesirable flow features can be demonstrated. The results of a two year effort to model the preburner are presented. In this effort, investigating the fuel preburner flowfield, the appropriate transport equations were numerically solved for both an axisymmetric and a three-dimensional configuration. Continuum's VAST (Variational Solution of the Transport equations) code, in conjunction with the CM-1000 Engineering Analysis Workstation and the NASA/Ames CYBER 205, was used to perform the required calculations. It is concluded that the preburner operational anomalies are not due to steady state phenomena and must, therefore, be related to transient operational procedures.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-178803 , NAS 1.26:178803 , CI-FR-0084
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  • 90
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    A data base and analysis program for shuttle main engine dynamic pressure measurements. Appendix F: Data base plots for SSME tests 750-120 through 750-200 (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A dynamic pressure data base and data base management system developed to characterize the Space Shuttle Main Engine (SSME) dynamic pressure environment is presented. The data base represents dynamic pressure measurements obtained during single engine hot firing tests of the SSME. Software is provided to permit statistical evaluation of selected measurements under specified operating conditions. An interpolation scheme is also included to estimate spectral trends with SSME power level.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-178813 , NAS 1.26:178813 , TR-66338-01-APP-F
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  • 91
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    A data base and analysis program for shuttle main engine dynamic pressure measurements. Appendix E: Data base plots for SSME tests 750-112 through 750-118 (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Dynamic pressure measurements and power spectra obtained from hot test firing of the Space Shuttle Main Engine are reported for test stand A3.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-178812 , NAS 1.26:178812 , TR-66338-01-APP-E
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  • 92
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    A data base and analysis program for shuttle main engine dynamic pressure measurements. Appendix D: Data base plots for SSME tests 750-047 through 750-111 (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Space Shuttle Main Engine power spectra and dynamic pressure measurements from hot test firing are reported.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-178811 , NAS 1.26:178811 , TR-66338-01-APP-D
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  • 93
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    A data base and analysis program for shuttle main engine dynamic pressure measurements. Appendix C: Data base plots for SSME tests 902-214 through 902-314 (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A dynamic pressure data base and data base management system developed to characterize the Space Shuttle Main Engine (SSME) dynamic pressure environment is reported. The data base represents dynamic pressure measurements obtained during single engine hot firing tests of the SSME. Software is provided to permit statistical evaluation of selected measurements under specified operating conditions. An interpolation scheme is included to estimate spectral trends with SSME power level. Flow Dynamic Environments in High Performance Rocket Engines are described.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-178810 , NAS 1.26:178810 , TR-66338-01-APP-C
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  • 94
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    A data base and analysis program for shuttle main engine dynamic pressure measurements. Appendix B: Data base plots for SSME tests 901-290 through 901-414 (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A dynamic pressure data base and data base management system developed to characterize the Space Shuttle Main Engine (SSME) dynamic pressure environment is described. The data base represents dynamic pressure measurements obtained during single engine hot firing tesets of the SSME. Software is provided to permit statistical evaluation of selected measurements under specified operating conditions. An interpolation scheme is also included to estimate spectral trends with SSME power level. Flow dynamic environments in high performance rocket engines are discussed.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-178809 , NAS 1.26:178809 , TR-66338-01-APP-B
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  • 95
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    Space Shuttle Main Engine (SSME) LOX turbopump pump-end bearing analysis (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A simulation of the shaft/bearing system of the Space Shuttle Main Engine Liquid Oxygen turbopump was developed. The simulation model allows the thermal and mechanical characteristics to interact as a realistic simulation of the bearing operating characteristics. The model accounts for single and two phase coolant conditions, and includes the heat generation from bearing friction and fluid stirring. Using the simulation model, parametric analyses were performed on the 45 mm pump-end bearings to investigate the sensitivity of bearing characteristics to contact friction, axial preload, coolant flow rate, coolant inlet temperature and quality, heat transfer coefficients, outer race clearance and misalignment, and the effects of thermally isolating the outer race from the isolator.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-178746 , NAS 1.26:178746 , SRS/STD-TR86-007-535
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  • 96
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    Cycle life of nickel-hydrogen cells. II - Accelerated cycle life test (1986)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A cycle life test of nickel-hydrogen (Ni/H2) cells containing electrolytes of various KOH concentrations and a sintered-type nickel electrode were carried out at 23 C using a 45-min accelerated low earth orbit (LEO) cycle regime at 80 percent depth of discharge. Ten cells containing 21 to 36 percent KOH were tested. Since this accelerated test regime accelerated the cycle life roughly twice as fast as a typical LEO regime, the present results indicate that the cells with 26 percent KOH may last over 5 years in an 80 percent depth-of-discharge cycling in an LEO regime. Cells with lower KOH concentrations (21 to 23.5 percent) also showed longer cycle life than those with KOH concentrations of 31 percent or higher, although the life was shorter than those with 26 percent KOH.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 97
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    Galileo probe battery program - An historical overview (1986)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The Galileo Probe mission to Jupiter required the selection of a battery module design that exceeded the known state of the art in 1977. The choice of the lithium-sulfur dioxide system, a technology then under development for nonaerospace applications necessitated an extensive cell and module development program that ultimately resulted in a space-qualified product that satisfies the severe constraints and requirements of the Galileo mission. The development program drew on the data base and experience then available from other government-sponsored lithium-sulfur dioxide programs and is an example of multiapplication synergism that can be derived from industry and government cooperation.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 98
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    Galileo probe battery systems design (1986)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: NASA's Galileo mission to Jupiter will consist of a Jovian orbiter and an atmospheric entry probe. The power for the probe will be derived from two primary power sources. The main source is composed of three Li-SO2 battery modules containing 13 D-size cell strings per module. These are required to retain capacity for 7.5 years, support a 150 day clock, and a 7 hour mission sequence of increasing loads from 0.15 to 9.5 amperes for the last 30 minutes. This main power source is supplemented by two thermal batteries (CaCrO4-Ca) for use in firing the pyrotechnic initiators during the atmospheric staging events. This paper describes design development and testing of these batteries at the system level.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 99
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    Configuration evaluation and criteria plan. Volume 1: System trades study and design methodology plan (preliminary). Space Transportation Main Engine (STME) configuration study (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The System Trades Study and Design Methodology Plan is used to conduct trade studies to define the combination of Space Shuttle Main Engine features that will optimize candidate engine configurations. This is accomplished by using vehicle sensitivities and engine parametric data to establish engine chamber pressure and area ratio design points for candidate engine configurations. Engineering analyses are to be conducted to refine and optimize the candidate configurations at their design points. The optimized engine data and characteristics are then evaluated and compared against other candidates being considered. The Evaluation Criteria Plan is then used to compare and rank the optimized engine configurations on the basis of cost.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-178939 , NAS 1.26:178939 , RPT/BB0324-VOL-1 , DR-9-VOL-1
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  • 100
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    Space shuttle main engine vibration data base (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The Space Shuttle Main Engine Vibration Data Base is described. Included is a detailed description of the data base components, the data acquisition process, the more sophisticated software routines, and the future data acquisition methods. Several figures and plots are provided to illustrate the various output formats accessible to the user. The numerous vibration data recall and analysis capabilities available through automated data base techniques are revealed.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-86576 , NAS 1.15:86576
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