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  • Other Sources  (749)
  • SPACECRAFT PROPULSION AND POWER  (749)
  • 1975-1979  (749)
  • 1
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    Nonlinear effects in thermal stress analysis of a solid propellant rocket motor (1976)
    In:  CASI
    Details
    Publication Date: 2006-08-09
    Description: Direct characterization procedures were used to determine the relaxation modulus as a function of time, temperature, and state of strain. Using the quasi-elastic method of linearviscoelasticity, these properties were employed in a finite element computer code to analyze a thick-walled, nonlinear viscoelastic cylinder in the state of plane strain bonded to a thin (but stiff) elastic casing and subjected to slow thermal cooling. The viscoelastic solution is then expressed as a sequence of elastic finite element solutions. The strain-dependent character of the relaxation modulus is included by replacing the single relaxation curve used in the linear viscoelastic theory by a family of relaxation functions obtained at various strain levels. These functions may be regarded as a collection of stress histories or responses to specific loads (in this case, step strains) with which the cooldown solution is made to agree by iterations on the modulus and strain level.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Langley Res. Center Advan. in Eng. Sci., Vol. 1; p 111-135
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  • 2
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    The plume impingement test program at AEDC utilizing the S-2 ullage motors (November 1973), section 1 (1976)
    In:  CASI
    Details
    Publication Date: 2006-04-12
    Description: Proposed experiments for analyzing rocket plumes are reported. Two groups of experiments were studied: (1) those that would help define some of the parameters that characterize the plume and (2) those that would enable evaluation of some of the contamination effects of the plume environment on various items of interest. The items investigated, the purpose of the investigation, are given in tabular form.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: The Effects of Solid Rocket Motor Effluents on Selected Surfaces and Solid Particle Size, Distribution, and Composition for Simulated Shuttle Booster Separation Motors; p 12-95
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  • 3
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    The small motor plume/material impingement test program at MSFC utilizing nine different propellant formulations (April and May 1974), section 2 (1976)
    In:  CASI
    Details
    Publication Date: 2006-04-12
    Description: The relative effects of several candidate SSRM propellant formulations and their plume impingement effects on HRSI and RCC materials were evaluated. Nine solid propellant formulations were tested. The selected propellant matrix allowed an evaluation of propellants with and without metal additives, with and without burning rate catalyst, and low (approximately 1927 C) and high (approximately 2649 C) combustion temperatures. Motors were fired at a simulated SRB staging altitude of 3.96 km (130,000 ft) (nominal). The altitude pressure was predicted to drop approximately 0.6 km (20,000 ft) during a motor firing. All motors were loaded with 1.8 to 2.3 kg (4 to 5 lb) of propellant and burned for approximately 2 s.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: The Effects of Solid Rocket Motor Effluents on Selected Surfaces and Solid Particle Size, Distribution, and Composition for Simulated Shuttle Booster Separation Motors; p 96-151
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  • 4
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    The large motor plume/material impingement test program at AEDC utilizing the four selected propellants from the small motor tests at MSFC (July, August, and September 1974), section 3 (1976)
    In:  CASI
    Details
    Publication Date: 2006-04-12
    Description: Efforts made to determine the vulnerability of Orbiter and ET materials located at various positions within exhaust plumes from test SSRM's using four different propellant formulations are discussed. Data also cover the effect on TPS materials from a single SSRM plume and dual SSRM plumes, and definitions of test SSRM plume environment at material specimen locations.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: The Effects of Solid Rocket Motor Effluents on Selected Surfaces and Solid Particle Size, Distribution, and Composition for Simulated Shuttle Booster Separation Motors; p 152-203
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  • 5
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    Propellant mass usage and flow measurement (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Practices are outlined for the design, installation, checkout, calibration, and operation of a system to be used for measuring propellant flow in a liquid monopropellant rocket engine. Design guidelines rather than detailed specifications are provided for the critical components of each portion of the system. System elemental uncertainties are presented in an appendix.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Handbook of Recommended Practices for the Determination of Liquid Monopropellant Rocket Engine Performance; 32 p
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  • 6
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    Basic stages in the development of the theory of Ramjet Engines (RJE) (1977)
    In:  CASI
    Details
    Publication Date: 2006-01-16
    Description: Basic periods in the history of the development of ramjet engine theory are cited. The periods include the first experimental tests as well as the development of basic ideas and theoretical development of the cosmic ramjet engine.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA, Washington Essays on the History of Rocketry and Astronautics, Vol. 1; p 229-238
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  • 7
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    Pressure measurement (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Practices are outlined for the design, installation, checkout, calibration, and operation of a pressure measuring system to be used during tests of a liquid monopropellant rocket engine. Appendixes include: (1) pressure measurement system elemental uncertainties; (2) short- and long-term pressure measurement system uncertainty; (3) shunt calibration of pressure transducers; (4) special considerations for vacuum measurement; and (5) methods of determining the dynamic characteristics of pressure transducers. Design guidelines are provided for the critical components of each portion of the system to provide a pressure measurement system which meets the performance criteria specified.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Handbook of Recommended Practices for the Determination of Liquid Monopropellant Rocket Engine Performance; 76 p
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  • 8
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    Temperature measurement (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: The design, installation, checkout, calibration, and operation of a temperature measuring system to be used during tests of a liquid monopropellant rocket engine are discussed. Appendixes include: (1) temperature measurement system elemental uncertainties, and (2) tables and equations for use with thermocouples and resistance thermometers. Design guidelines are given for the critical components of each portion of the system to provide an optimum temperature measurement system which meets the performance criteria specified.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Handbook of Recommended Practices for the Determination of Liquid Monopropellant Rocket Engine Performance; 36 p
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  • 9
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    Force and impulse measurement (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Practices are outlined for the design, installation, checkout, calibration, and operation of a thrust measurement system to be used during tests of a liquid monopropellant rocket engine. Appendixes include: (1) thrust measurement system elemental uncertainties; (2) short- and long-term thrust measurement system uncertainty; and (3) shunt calibration of force transducers.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Handbook of Recommended Practices for the Determination of Liquid Monopropellant Rocket Engine Performance; 28 p
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  • 10
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    Data reduction and performance determination (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Definitions, algorithms, and procedures for the reduction of monopropellant thruster measurements to performance parameters are provided. A brief discussion of acquisition and recording systems is also included. Emphasis is placed upon monopropellant hydrazine engines, and some parameters relate specifically to the catalytic decomposition of hydrazine (e.g., percent ammonia dissociation). Performance of other types of monopropellant thrusters may, however, be determined by using procedures similar to those discussed. Two appendixes are included: (1) theoretical performance of monopropellant hydrazine; and (2) calculation of rotational performance.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Handbook of Recommended Practices for the Determination of Liquid Monopropellant Rocket Engine Performance; 45 p
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  • 11
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    Materials and processes for space shuttle's engines (1975)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: It is pointed out that over 50 different alloys are used in construction of the space shuttle main engines (SSME). Primary construction of the SSME is by welding or brazing of wrought and cast components. Welding processes involve both gas tungsten-arc welds and electron-beam welds. Electroforming has been developed as a process to fabricate and bond structural members for the SSME. Important aspects in the selection of materials and processes are related to weight saving considerations and the high-pressure hydrogen environment. Special problems and their solution in the case of various engine components are discussed, giving attention to the oxidizer preburner, the high pressure oxidizer turbopump, and the heat exchanger.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 12
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    Parallel-burn options for dual-fuel single-stage orbital transports (1978)
    In:  Other Sources
    Details
    Publication Date: 2011-08-17
    Description: A parallel-burn version of a single-stage vehicle for transport from the earth to low-earth orbit using two fuels and rocket propulsion is considered. New engine results were incorporated in vehicle performance and design studies. The results indicate that a hydrogen-cooled gas generator cycle engine provides attractive vehicle performance and that there is little incentive for increasing the chamber pressure beyond 27 MPa.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Spacecraft and Rockets; 15; Jan
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  • 13
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    Comment on 'Heat-pipe reactors for space power applications' (1978)
    In:  Other Sources
    Details
    Publication Date: 2011-08-17
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Energy; 2; May-June
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  • 14
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    Monte Carlo investigation of thrust imbalance of solid rocket motor pairs (1976)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: The Monte Carlo method of statistical analysis is used to investigate the theoretical thrust imbalance of pairs of solid rocket motors (SRMs) firing in parallel. Sets of the significant variables are selected using a random sampling technique and the imbalance calculated for a large number of motor pairs using a simplified, but comprehensive, model of the internal ballistics. The treatment of burning surface geometry allows for the variations in the ovality and alignment of the motor case and mandrel as well as those arising from differences in the basic size dimensions and propellant properties. The analysis is used to predict the thrust-time characteristics of 130 randomly selected pairs of Titan IIIC SRMs. A statistical comparison of the results with test data for 20 pairs shows the theory underpredicts the standard deviation in maximum thrust imbalance by 20% with variability in burning times matched within 2%. The range in thrust imbalance of Space Shuttle type SRM pairs is also estimated using applicable tolerances and variabilities and a correction factor based on the Titan IIIC analysis.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Spacecraft and Rockets; 13; Apr. 197
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  • 15
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    Quality assurance and control in the production and static tests of the solid rocket boosters for the Space Shuttle (1979)
    In:  Other Sources
    Details
    Publication Date: 2011-08-17
    Description: The paper surveys the various aspects of design and overhaul of the solid rocket boosters. It is noted that quality control is an integral part of the design specifications. Attention is given to the production process which is optimized towards highest quality. Also discussed is the role of the DCA (Defense Contract Administration) in inspecting the products of subcontractors, noting that the USAF performs this role for prime contractors. Fabrication and construction of the booster is detailed with attention given to the lining of the booster cylinder and the mixing of the propellant and the subsequent X-ray inspection.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Astronautik; 16; 2, 19; 1979
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  • 16
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    Current leakage for low altitude satellites - Modeling applications (1979)
    In:  Other Sources
    Details
    Publication Date: 2011-08-17
    Description: To simulate the behavior of a high voltage solar cell array in the ionospheric plasma environment, the large (90 ft x 55 ft diameter) vacuum chamber was used to measure the high-voltage plasma interactions of a 3 ft x 30 ft conductive panel. The chamber was filled with Nitrogen and Argon plasma at electron densities of up to 1,000,000 per cu cm. Measurements of current flow to the plasma were made in three configurations: (a) with one end of the panel grounded, (b) with the whole panel floating while a high bias was applied between the ends of the panel, and (c) with the whole panel at high negative voltage with respect to the chamber walls. The results indicate that a simple model with a constant panel conductivity and plasma resistance can adequately describe the voltage distribution along the panel and the plasma current flow. As expected, when a high potential difference is applied to the panel ends more than 95% of the panel floats negative with respect to the plasma.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 17
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    High temperature reformation of aluminum and chlorine compounds behind the Mach disk of a solid-fuel rocket exhaust (1976)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: Chemical reactions expected to occur among the constituents of solid-fuel rocket engine effluents in the hot region behind a Mach disk are analyzed theoretically. With the use of a rocket plume model that assumes the flow to be separated in the base region, and a chemical reaction scheme that includes evaporation of alumina and the associated reactions of 17 gas species, the reformation of the effluent is calculated. It is shown that AlClO and AlOH are produced in exchange for a corresponding reduction in the amounts of HCl and Al2O3. For the case of the space shuttle booster engines, up to 2% of the original mass of the rocket fuel can possibly be converted to these two new species and deposited in the atmosphere between the altitudes of 10 and 40 km. No adverse effects on the atmospheric environment are anticipated with the addition of these two new species.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Atmospheric Environment; 10; 9, 19; 1976
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  • 18
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    Optimum exhaust velocity for laser-driven rockets (1975)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: The laser-driven rocket in which remotely generated laser power is used to heat propellant belongs basically to the class of specific-impulse limited propulsion systems if difficult missions are considered. It was previously established that trip time reaches a minimum as specific impulse is varied for payload transfers from low earth orbit to synchronous orbit and return via laser-driven rocket propulsion, the computations being based on the perigee-propulsion laser drive described by Minovitch (1972). The present study shows that such minimum occur for all missions and that optimum specific impulse is primarily determined by the mission difficulty. More generally, this optimum specific impulse maximizes payload kinetic energy achievable with a fixed jet power and propulsion time. A formula relating propulsion time parameter to payload ratio is obtained for estimating mission capabilities of laser-driven rockets.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Spacecraft and Rockets; 12; Nov. 197
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  • 19
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    ATS-6 - Cesium bombardment engine north-south stationkeeping experiment (1975)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: Two 0.004 N thrust cesium bombardment ion thrustors have been developed and used for north-south stationkeeping in the geostationary Applications Technology Satellite-6 (ATS-6). The thrustor subsystems are mounted on the north and south faces of the earth viewing module such that 0.0026 N of thrust is applied normal to the orbit plane and 0.0036 N is applied radially upward. The change in the orbit inclination of the satellite is maintained at zero by operating the two thrustors alternately so that their thrust components, normal to the orbital plane, are symmetrically applied about the nodal crossings. Initial operation of the thrustors was successful. There was no interference with the satellite communications systems and the predicted spacecraft operating potential was verified. Subsequent trials failed due to a defect in the operation of the propellant reservoirs in zero g. A feed line valve is under development to correct this difficulty.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: IEEE Transactions on Aerospace and Electronic Systems; AES-11; Nov. 197
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  • 20
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    Annealing kinetics of electron-irradiated GaAs heteroface solar cells in the range 175-200 C (1979)
    In:  Other Sources
    Details
    Publication Date: 2011-08-17
    Description: Annealing of electron-irradiation damage in GaAs solar cells is important for space applications. This paper describes studies conducted to understand this annealing process. GaAs heteroface solar cells were irradiated with 10 to the 15th 1-MeV electrons/sq cm followed by thermal annealing. An activation energy for annealing of 1.25 + or - 0.14 eV and a frequency factor of (3.7 + or - 1.9) x 10 to the 9th per sec were determined. A small component of the irradiation damage which does not anneal measurably at 200 C was observed.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Applied Physics Letters; 35; Sept. 15
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  • 21
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    Electrets used in measuring rocket exhaust effuents from a Space Shuttle model (1979)
    In:  Other Sources
    Details
    Publication Date: 2011-08-17
    Description: The study evaluates the use of electrets as a new contamination-detecting device designed to assess the chemical composition of rocket effluents. Evaluation of electret effectiveness revealed that electrets have multipollutant-measuring capability, simplicity of deployment and rapidity of assessment. Advantages of electrets are small size, light weight and cost-effectiveness. It is shown that electrets compare favorably with other HCl measuring devices. In particular, the summary of the measured data from electrets and HCl detectors is within the limits of computed HCl concentrations from the NASA/MSFC multilayer diffusion model.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Applied Meteorology; 18; Jan. 197
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  • 22
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    Aerospace power systems - A building surge (1979)
    In:  Other Sources
    Details
    Publication Date: 2011-08-17
    Description: Dramatic changes in both aircraft and space power systems are being projected. The primary driver is mission cost, although more stringent performance requirements are also being imposed. The order-of-magnitude reductions anticipated in space power, both in cost and specific mass, are partially related to the advent of the Shuttle and the continuing evolution of the STS over the next decade. Automation, microcircuitry, batteries, and structural materials are common aircraft and space technology needs.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Astronautics and Aeronautics; 17; Feb. 197
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  • 23
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    Battery workshop (1978)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The workshop was attended by representatives from industry and government. The requirements for energy storage and the plans for battery development were reviewed. The workshop followed a debate format, with the objective of recommending improvements to the development plans presented by NASA and the Air Force. The issues addressed were: (1) significant technology deficiencies which can be identified; (2) adequacy of current and proposed programs to resolve the technology deficiencies identified; (3) additional tasks which should be undertaken, including benefits and timing; and (4) lowest priority items in the presently planned program, both in content and in timing.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Res. Center Future Orbital Power Systems Technol. Requirements; p 283-287
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  • 24
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    Solar array workshop (1978)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The solar workshop began with a review of the needs and objectives in this area as presented by the various government representatives during the preceding sessions. The major problem noted with respect to needs was the potentially conflicting requirements of low cost and low weight. Since the importance of weight and cost and relationship between them are strongly mission dependent, the workshop concluded that the requirements of military missions in synchronous orbit could be quite different from the requirements of NASA low-orbit missions and that an assignment of specific technology deficiencies could only be related to specific mission classes.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Res. Center Future Orbital Power Systems Technol. Requirements; p 279-282
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  • 25
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    Solar cell workshop (1978)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The workshop addressed three issues in respect to the NASA solar cell technology requirements for future orbital missions. First, technology areas were identified that were considered most significant and the deficiencies and concerns that were had with each area are indicated. Second, the tasks that should be undertaken to reduce the costs and risks of future orbital power systems are recommended. Third, an attempt to identify the lowest priority items in the present program in terms of content and timing are made.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Res. Center Future Orbital Power Systems Technol. Requirements; p 275-277
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  • 26
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    Photovoltaic power systems workshop (1978)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Discussions are presented on apparent deficiencies in NASA planning and technology development relating to a standard power module (25-35 kW) and to future photovoltaic power systems in general. Topics of discussion consider the following: (1) adequate studies on power systems; (2) whether a standard power system module should be developed from a standard spacecraft; (3) identification of proper approaches to cost reduction; (4) energy storage avoidance; (5) attitude control; (6) thermal effects of heat rejection on solar array configuration stability; (7) assembly of large power systems in space; and (8) factoring terrestrial photovoltaic work into space power systems for possible payoff.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Res. Center Future Orbital Power Systems Technol. Requirements; p 271-273
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  • 27
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    An economical approach to space power systems (1978)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Projected energy demand for all NASA, DoD and civil missions for the time span 1981 to 1995 are illustrated. Typical energy cost range from about $300 to $2000 per kW-hr, with an average of about $800 per kW-hr for long-duration missions. At these levels, the cost of the required energy would be several billion dollars per year by about 1985 and might constrain the number and types of NASA programs to be carried out. NASA is extensively pursuing approaches for reducing nonrecurring costs. Two programs are presented for the development of an economical approach to space power systems. They are: (1) Economical Orbital Power (ECOP) with the objective to demonstrate the applicability of a commercial approach to the development of a low cost photovoltaic space power system; and (2) Space Power Experiment (SPEX) which has the objective to demonstrate the application of industrial hardware for space power systems.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Future Orbital Power Systems Technol. Requirements; p 265-270
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  • 28
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    Power modules and projected power systems evolution (1978)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Photovoltaic, solar thermal, and nuclear power systems are considered to supply future earth orbital electrical power requirements. A growth scenario from a 25-kW Power Module in the early Shuttle era to the 5- to 10-GW Satellite Power System in the year 2000 is presented. Photovoltaic systems are presently baselined in this evolution. The Photovoltaic Power System and subsystem growth projections, consistent with this scenario, were developed and are summarized.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Res. Center Future Orbital Power Systems Technol. Requirements; p 235-246
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  • 29
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    JSC space base/power module studies (1978)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Users of the Orbiter/Spacelab combination will require both higher electrical power and longer duration than is available with the current baseline system. Present Orbiter/Spacelab mission capability is primarily constrained by the hydrogen and oxygen available to generate power in the Orbiter fuel cells. It is also necessary to assure that considerable attitude or point flexibility is retained to assure efficient operation of the Orbiter radiator cooling system. Beyond these early limitations, it is foreseen that orbital operations will eventually need even greater quantities of the basic space utilities: electrical power; heat rejection; and attitude control. Such operations, forecasted for the mid to late 1980's, will be best accommodated by a module stored in orbit that can furnish these to a docked Orbiter/Spacelab or other vehicles. The Orbital Service Module concept to provide for these services is presented.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Res. Center Future Orbital Power Systems Technol. Requirements; p 247-264
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  • 30
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    Spacecraft active thermal control technology status (1978)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Four advanced space radiator concepts that were pursued in an integrated effort to develop multi-mission-use and low cost heat rejection systems which can overcome the limitations of current radiator systems are briefly discussed and described. Also, in order to establish a firm background to compare the advanced space radiator concepts, the Orbiter active thermal control system is also briefly described.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Res. Center Future Orbital Power Systems Technol. Requirements; p 213-233
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  • 31
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    Laser power transmission for space power and propulsion (1978)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: An early status report is presented for a review being made of the state of development of major components and subsystems required for ground-to-space, space-to-space, or space-to-ground laser power transmission for electric or thermal power or propulsion. System characteristics are being evaluated from an applications viewpoint, and major problem areas are being identified. The object is to identify a rewarding first application of lasers for space power and propulsion. An evolution of laser power transmission capabilities over the next 20 years is projected. Supporting technology requirements are to be identified, priorities set, and continued developments coordinated with other government agencies.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Res. Center Future Orbital Power Systems Technol. Requirements; p 209-211
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    Technology Status: Fuel Cells and Electrolysis Cells (1978)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The status of the baselined shuttle fuel cell as well as the acid membrane fuel cell and space-oriented water electrolysis technologies are presented. The more recent advances in the alkaline fuel cell technology area are the subject of a companion paper. A preliminary plan for the focusing of these technologies towards regenerative energy storage applications in the multi-hundred kilowatt range is also discussed.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Res. Center Future Orbital Power Systems Technol. Requirements; p 167-194
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    Power management and control for space systems (1978)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Power management and control technology for the large, high-power spacecraft of the 1980's is discussed. Systems weight optimization that indicate a need for higher bus voltages are shown. Environmental interactions that are practical limits for the maximum potential on exposed surfaces are shown. A dual-voltage system is proposed that would provide the weight savings of a high-voltage distribution system and take into account the potential environmental interactions. The technology development of new components and circuits is also discussed.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Future Orbital Power Systems Technol. Requirements; p 195-207
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    Technology status: Batteries and fuel cells (1978)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The current status of research and development programs on batteries and fuel cells and the technology goals being pursued are discussed. Emphasis is placed upon those technologies relevant to earth orbital electric energy storage applications.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Future Orbital Power Systems Technol. Requirements; p 157-166
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    Solar array systems (1978)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The recent past, present state-of-the-art, and future needs in the area of large photovoltaic solar arrays are discussed. In the past most attention was focused upon performance whereas in the future most of the effort should go into cost reduction. Suggestions are made regarding possible approaches to reducing cost such as on-orbit maintenance, extended lifetime, solar concentrators, and high-voltage modular concepts.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Res. Center Future Orbital Power Systems Technol. Requirements; p 147-155
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  • 36
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    A brief survey of the solar cell state-of-the-art (1978)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Modern high performance cells made for space are discussed. The major recent developments that are expected to influence what solar cells will be available in five years or so are described.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Future Orbital Power Systems Technol. Requirements; p 133-146
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  • 37
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    Alternative power-generation systems (1978)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The present state of the art of thermal power systems is surveyed. Because of the great potential variety of thermal power systems, the heat sources, the power conversion systems, and the integration of thermal power systems with missions are treated sequentially.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Future Orbital Power Systems Technol. Requirements; p 113-131
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  • 38
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    Military needs for orbital power (1978)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Results of the DoD/ERDA (now Department of Energy) Space Power Study completed in October 1977 are presented. The major new thrust of Air Force Advanced Technology Plans center on the development of military solar power systems which will extend capabilities to the 10 - 50 KW sub e power range for new classes of missions while maintaining technology applicability to the 0.5 - 10 KW sub e present mission class. The status of FY78 efforts for Project 682J (Air Force Space Power Advanced Development Program) are reported. Project 682J is divided into the following tasks: (1) high efficiency solar panel; (2) nickel-hydrogen battery; (3) gallium arsenide solar concentrator hardness study; and (4) new-start nuclear dynamic power system applications/integration study.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Res. Center Future Orbital Power Systems Technol. Requirements; p 93-107
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  • 39
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    Historical and projected power requirements (1978)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Policy planning for projected space power requirements is discussed. Topics of discussion cover: (1) historical space power trends (prime power requirements and power system costs); and (2) two approaches to future space power requirements (mission/traffic model approach and advanced system scenario approach). Graphs, tables, and flow charts are presented.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Lewis Res. Center Future Orbital Power Systems Technol. Requirements; p 41-69
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  • 40
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    Exhaust gas composition measurement (1979)
    In:  CASI
    Details
    Publication Date: 2017-10-02
    Description: The design, installation, checkout, and operation of an exhaust gas composition measurement system for collecting and analyzing the exhaust gas from a liquid monopropellant rocket engine are described. Design guidelines are given for the critical components of each portion of the system to provide an exhaust gas composition measurement which meets the performance criteria specified.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Handbook of Recommended Practices for the Determination of Liquid Monopropellant Rocket Engine Performance; 13 p
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  • 41
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    Flow field for an underexpanded, supersonic nozzle exhausting into an expansive launch tube (1976)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Static pressure distributions along the launcher wall and pitot pressure measurements from the annular region between the rocket and the launcher were made as an underexpanded supersonic nozzle exhausted into an expansive launch tube. The flow remained supersonic along the entire length of the launcher for all nozzle locations studied.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA. Langley Res. Center Advan. in Eng. Sci., Vol. 4; p 1665-1671
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  • 42
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    Solar electric propulsion (1975)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The major components of a solar electric propulsion system are discussed and some problems in low thrust mission analysis are detailed. Emphasis is placed on the development of a nominal low thrust trajectory and guidance and navigation aspects.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Flight Mech.(Estimation Theory Symp.; p 73-77
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  • 43
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    Technical status of the Space Shuttle Main Engine (second review) (1979)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-172838 , NAS 1.26:172838 , PB83-181040
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  • 44
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    Concept definition and system analysis study for a solar electric propulsion stage. Volume 1: Executive summary (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-161599 , D180-18553-1
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  • 45
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    Orbit Transfer Vehicle Engine Study. Phase A, extension 1: Study plan update (1979)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The updated study plan for the Space Transportation System orbit transfer vehicle (OTV) engine study is presented. The study program consists of engine system, programmatic, cost, and risk analyses of OTV engine concepts. Detailed task descriptions for the advanced expander cycle engine optimization, alternate low thrust capability, and safety, reliability, and cost comparisons are given.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-161708 , REPT-32999-SP2
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  • 46
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    Orbit Transfer Vehicle Engine Study. Phase A, extension 1: Advanced expander cycle engine optimization (1979)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The performance optimization of expander cycle engines at vacuum thrust levels of 10K, 15K, and 20K lb is discussed. The optimization is conducted for a maximum engine length with an extendible nozzle in the retracted position of 60 inches and an engine mixture ratio of 6.0:1. The thrust chamber geometry and cycle analyses are documented. In addition, the sensitivity of a recommended baseline expander cycle to component performance variations is determined and chilldown/start propellant consumptions are estimated.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-161710 , REPT-3299-E1-T1
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    Plume interference with space shuttle range safety signals (1979)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: The computational procedure for signal propagation in the presence of an exhaust plume is presented. Comparisons with well-known analytic diffraction solutions indicate that accuracy suffers when mesh spacing is inadequate to resolve the first unobstructed Fresnel zone at the plume edge. Revisions to the procedure to improve its accuracy without requiring very large arrays are discussed. Comparisons to field measurements during a shuttle solid rocket motor (SRM) test firing suggest that the plume is sharper edged than one would expect on the basis of time averaged electron density calculations. The effects, both of revisions to the computational procedure and of allowing for a sharper plume edge, are to raise the signal level near tail aspect. The attenuation levels then predicted are still high enough to be of concern near SRM burnout for northerly launches of the space shuttle.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: APL JANNAF 11th Plume Technol. Meeting, Vol. 2; p 1-25
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  • 48
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    A nozzle boundary layer model including the subsonic sublayer usable for determining boundary layer effects on plume flowfields (1979)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: A model for the boundary layer at the exit plane of a rocket nozzle was developed which, unlike most previous models, includes the subsonic sublayer. The equations for the flow near the nozzle exit plane are presented and the method by which the subsonic sublayer transitions to supersonic flow in the plume is described. The resulting model describes the entire boundary layer and can be used to provide a startline for method-of-characteristics calculations of plume flowfields. The model was incorporated into a method of characteristics computer program and comparisons of computed results to experimental data show good agreement. The data used in the comparisons were obtained in tests in which mass fluxes from a 22.2-N (5 lbf) thrust engine were measured at angles off the nozzle centerline of up to 150 deg. Additional comparisons were made with data obtained during tests of a 0.89-N (0.2 lbr) monopropellant thruster and from the OH-64 space shuttle heating tests. The agreement with the data indicates that the model can be used for calculating plume backflow properties.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: APL JANNAF 11th Plume Technol. Meeting, Vol. 2; p 147-165
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    Measurements of reactive gaseous rocket injector admittances (1979)
    In:  Other Sources
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    Publication Date: 2019-06-27
    Description: The paper describes the results of an experimental study of the quantitative determination of the capabilities of the combustion processes associated with coaxial gaseous propellant rocket injectors to drive combustor pressure oscillations. The data, obtained by employing the modified impedance tube technique with compressed air as the oxidizer and acetylene gas as the fuel, describe the frequency dependence of the admittance of the combined injector-combustion process. The measured data are compared with the predictions of the Feiler and Heidmann analytical model utilizing different values for the characteristic combustion time tau sub b. The values of tau sub b which result in a best fit between the measured and predicted data are indicated for different equivalence ratios. It is shown that for the coaxial injector investigated in this study the tau sub b varies between 0.7 and 1.2 msec for equivalence ratios in the range of 0.57 to 1.31. In addition, the experimental data indicate that the tested injector system could drive combustion instabilities over a frequency range that is in qualitative agreement with the predictions of the Feiler and Heidmann model.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Combustion Science and Technology; 20; 5-6,; 1979
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    A computer program for thermal radiation from gaseous rocket exhuast plumes (GASRAD) (1979)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A computer code is presented for predicting incident thermal radiation from defined plume gas properties in either axisymmetric or cylindrical coordinate systems. The radiation model is a statistical band model for exponential line strength distribution with Lorentz/Doppler line shapes for 5 gaseous species (H2O, CO2, CO, HCl and HF) and an appoximate (non-scattering) treatment of carbon particles. The Curtis-Godson approximation is used for inhomogeneous gases, but a subroutine is available for using Young's intuitive derivative method for H2O with Lorentz line shape and exponentially-tailed-inverse line strength distribution. The geometry model provides integration over a hemisphere with up to 6 individually oriented identical axisymmetric plumes, a single 3-D plume, Shading surfaces may be used in any of 7 shapes, and a conical limit may be defined for the plume to set individual line-of-signt limits. Intermediate coordinate systems may specified to simplify input of plumes and shading surfaces.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-161496 , RTR-014-9
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    Field resonance propulsion concept (1979)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A propulsion concept was developed based on a proposed resonance between coherent, pulsed electromagnetic wave forms, and gravitational wave forms (or space-time metrics). Using this concept a spacecraft propulsion system potentially capable of galactic and intergalactic travel without prohibitive travel times was designed. The propulsion system utilizes recent research associated with magnetic field line merging, hydromagnetic wave effects, free-electron lasers, laser generation of megagauss fields, and special structural and containment metals. The research required to determine potential, field resonance characteristics and to evaluate various aspects of the spacecraft propulsion design is described.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-80961 , JSC-16073
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    Sealed-cell nickel-cadmium battery applications manual (1979)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The design, procurement, testing, and application of aerospace quality, hermetically sealed nickel-cadmium cells and batteries are presented. Cell technology, cell and battery development, and spacecraft applications are emphasized. Long term performance is discussed in terms of the effect of initial design, process, and application variables. Design guidelines and practices are given.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-RP-1052
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  • 53
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    Inert gas thrusters (1979)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Inert gas thrusters considered for space propulsion systems were investigated. Electron diffusion across a magnetic field was examined utilizing a basic model. The production of doubly charged ions was correlated using only overall performance parameters. The use of this correlation is therefore possible in the design stage of large gas thrusters, where detailed plasma properties are not available. Argon hollow cathode performance was investigated over a range of emission currents, with the positions of the inert, keeper, and anode varied. A general trend observed was that the maximum ratio of emission to flow rate increased at higher propellant flow rates. It was also found that an enclosed keeper enhances maximum cathode emission at high flow rates. The maximum cathode emission at a given flow rate was associated with a noisy high voltage mode. Although this mode has some similarities to the plume mode found at low flows and emissions, it is encountered by being initially in the spot mode and increasing emission. A detailed analysis of large, inert-gas thruster performance was carried out. For maximum thruster efficiency, the optimum beam diameter increases from less than a meter at under 2000 sec specific impulse to several meters at 10,000 sec. The corresponding range in input power ranges from several kilowatts to megawatts.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-159813
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  • 54
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    Solid rocket booster thrust vector control subsystem verification test (V-2) report (1979)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The results of the verification testing sequence V-2 performed on the space shuttle solid rocket booster thrust vector control subsystem are presented. A detailed history of the hot firings plus additional discussion of the auxiliary power unit and the hydraulic component performance is presented. The test objectives, data, and conclusions are included.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-78258
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  • 55
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    Supercharged topping rocket propellant feed system (1979)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A rocket propellant feed system utilizing a bleed turbopump to supercharge a topping turbopump is presented. The bleed turbopump is of a low pressure type to meet the cavitation requirements imposed by the propellant storage tanks. The topping turbopump is of a high pressure type and develops 60 to 70 percent of the pressure rise in the propellant.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 56
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    Performance of a transpiration-regenerative cooled rocket thrust chamber (1979)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The analysis, design, fabrication, and testing of a liquid rocket engine thrust chamber which is gas transpiration cooled in the high heat flux convergent portion of the chamber and water jacket cooled (simulated regenerative) in the barrel and divergent sections of the chamber are described. The engine burns LOX-hydrogen propellants at a chamber pressure of 600 psia. Various transpiration coolant flow rates were tested with resultant local hot gas wall temperatures in the 800 F to 1400 F range. The feasibility of transpiration cooling with hydrogen and helium, and the use of photo-etched copper platelets for heat transfer and coolant metering was successfully demonstrated.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-159742
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  • 57
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    Stability analysis of a liquid fuel annular combustion chamber (1978)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: High frequency combustion instability problems in a liquid fuel annular combustion chamber are examined. A modified Galerkin method was used to produce a set of modal amplitude equations from the general nonlinear partial differential acoustic wave equation in order to analyze the problem of instability. From these modal amplitude equations, the two variable perturbation method was used to develop a set of approximate equations of a given order of magnitude. These equations were modeled to show the effects of velocity sensitive combustion instabilities by evaluating the effects of certain parameters in the given set of equations.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-159734
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  • 58
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    Design, fabrication and delivery of a high pressure LOX-methane injector (1979)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An injector for 3000 psia chamber pressure using liquid oxygen and gaseous methane propellants is presented. The injector is intended to be evaluated during a series of pressure-fed test firings using a water-cooled calorimeter chamber and a milled-slot regenerative chamber. Combustion efficiency, combustion stability, ignition and injector face heat transfer assessments were made for candidate injector body and pattern design approaches. This evaluation resulted in baselining an oxidizer post type manifold with a 60 element platelet coaxial swirler injector pattern. An axial acoustic resonator cavity was created at the injector/chamber interface.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-161343 , REPT-33205F
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  • 59
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    Orbital Transfer Vehicle (OTV) engine study. Phase A: Extension (1979)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An orbital transfer vehicle OTV engine study program was undertaken to provide additional expander and staged combustion cycle data in the design definition of the OTV engine. The proposed program effort optimizes the expander cycle engine concept (consistent with identified OTV engine requirements), investigates the feasibility of kitting the staged combustion cycle engine to provide extended low thrust operation, and conducts in-depth analysis of development risk, crew safety, and reliability for both cycles. Additional tasks to establish the cost of a 10K thrust expander cycle engine and to render support of OTV systems study contractors are reported.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-161338 , ASR79-116 , MPR-2
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  • 60
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    Study of reactor Brayton power systems for nuclear electric spacecraft (1979)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The feasibility of using Brayton power systems for nuclear electric spacecraft was investigated. The primary performance parameters of systems mass and radiator area were determined for systems from 100 to 1000 kW sub e. Mathematical models of all system components were used to determine masses and volumes. Two completely independent systems provide propulsion power so that no single-point failure can jeopardize a mission. The waste heat radiators utilize armored heat pipes to limit meteorite puncture. The armor thickness was statistically determined to achieve the required probability of survival. A 400 kW sub e reference system received primary attention as required by the contract. The components of this system were defined and a conceptual layout was developed with encouraging results. An arrangement with redundant Brayton power systems having a 1500 K (2240 F) turbine inlet temperature was shown to be compatible with the dimensions of the space shuttle orbiter payload bay.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-168942 , NAS 1.26:168942 , AIRESEARCH-31-3321
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  • 61
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    SRB TPS closeout materials characterization (1979)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The recession characteristics of the materials used to closeout the Thermal Protection System (TPS) on the Space Shuttle Solid Rocket Booster (SRB) were assessed. Two candidate closeout materials namely, K5NA and MTA2 is the Marshall Trowelable Ablator. Test panels of each of these two materials were subjected to aerodynamic heat and shear in the NASA Hot Gas Facility. Pretest and Post-test measurements were taken and the rate of recession (R) thus determined was correlated with the heating rate. A least squares curve fit of the data together with another least squares 95% confidence level fit, was performed. The characteristics of the two materials when compared showed that both materials performed equally well at the higher heating rates (10 q 30 Btu/sq ft/sec) but MTA2 was slightly better, not receding as fast, at the lower q values. The comparison of the R versus q curves for K5NA and MTA2 with that of primary TPS P-50 cork on the SRB Aft Skirt was good.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-170885 , NAS 1.26:170885 , LMSC-HREC-TN-D697757
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  • 62
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    Thermal radiation model for solid rocket booster plumes (1977)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: The Monte Carlo method is used to model the thermal radiation field of the plumes for the dual solid rocket boosters astride the Space Shuttle launch configuration. The model accounts for axial and radial variations in radiative properties of the plumes. The plumes are considered to be composed of a dispersion of aluminum oxide (Al2O3) particles immersed in the gaseous products of combustion. The principal emitting gases are taken to be CO, CO2, H2O, and HCl. The thermal model is based on local thermodynamic equilibrium. Scattering of radiant energy by Al2O3 particles may be treated as isotropic or anisotropic. Sample radiant heating rates to the base region of the Space Shuttle are shown. Space Shuttle geometries are simulated as combinations of quadric surfaces.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Spacecraft and Rockets; 14; Nov. 197
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    Solid propellant motor (1978)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A case bonded end burning solid propellant rocket motor is described. A propellant with sufficiently low modulus to avoid chamber buckling on cooling from cure and sufficiently high elongation to sustain the stresses induced without cracking is used. The propellant is zone cured within the motor case at high pressures equal to or approaching the pressure at which the motor will operate during combustion. A solid propellant motor with a burning time long enough that its spacecraft would be limited to a maximum acceleration of less than 1 g is provided by one version of the case bonded end burning solid propellant motor of the invention.
    Keywords: SPACECRAFT PROPULSION AND POWER
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    Qualification test for safety and arming device MSFC PN 10A00465 TMc/S PN 816780 for space shuttle solid rocket booster, volume 2 (1978)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: For abstract, see N78-31159.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-150793 , QTR-4918-245-VOL-2
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  • 65
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    Passive propellant system (1978)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The system utilizes a spherical tank structure A separated into two equal volume compartments by a flat bulkhead B. Each compartment has four similar gallery channel legs located in the principal vehicle axes, ensuring that bulk propellant will contact at least one gallery leg during vehicle maneuvers. The forward compartment gallery channel legs collect propellant and feed it into the aft compartment through communication screens which protrude into the aft compartment. The propellant is then collected by the screened gallery channels in the aft compartment and supplied to the propellant outlet. The invention resides in the independent gallery assembly and screen structure by means of which propellant flow from forward to aft compartments is maintained. Liquid surface tension of the liquid on the screens is used to control liquid flow. The system provides gas-free propellants in low or zero-g environments regardless of axial accelerations and propellant orientation in bulk regions of the vessel.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 66
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    Solid propellant rocket motor internal ballistics performance variation analysis, phase 3 (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Results of research aimed at improving the predictability of off nominal internal ballistics performance of solid propellant rocket motors (SRMs) including thrust imbalance between two SRMs firing in parallel are reported. The potential effects of nozzle throat erosion on internal ballistic performance were studied and a propellant burning rate low postulated. The propellant burning rate model when coupled with the grain deformation model permits an excellent match between theoretical results and test data for the Titan IIIC, TU455.02, and the first Space Shuttle SRM (DM-1). Analysis of star grain deformation using an experimental model and a finite element model shows the star grain deformation effects for the Space Shuttle to be small in comparison to those of the circular perforated grain. An alternative technique was developed for predicting thrust imbalance without recourse to the Monte Carlo computer program. A scaling relationship used to relate theoretical results to test results may be applied to the alternative technique of predicting thrust imbalance or to the Monte Carlo evaluation. Extended investigation into the effect of strain rate on propellant burning rate leads to the conclusion that the thermoelastic effect is generally negligible for both steadily increasing pressure loads and oscillatory loads.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-150577
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  • 67
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    Liquid rocket engine self-cooled combustion chambers (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Self-cooled combustion chambers are chambers in which the chamber wall temperature is controlled by methods other than fluid flow within the chamber wall supplied from an external source. In such chambers, adiabatic wall temperature may be controlled by use of upstream fluid components such as the injector or a film-coolant ring, or by internal flow of self-contained materials; e.g. pyrolysis gas flow in charring ablators, and the flow of infiltrated liquid metals in porous matrices. Five types of self-cooled chambers are considered in this monograph. The name identifying the chamber is indicative of the method (mechanism) by which the chamber is cooled, as follows: ablative; radiation cooled; internally regenerative (Interegen); heat sink; adiabatic wall. Except for the Interegen and heat sink concepts, each chamber type is discussed separately. A separate and final section of the monograph deals with heat transfer to the chamber wall and treats Stanton number evaluation, film cooling, and film-coolant injection techniques, since these subjects are common to all chamber types. Techniques for analysis of gas film cooling and liquid film cooling are presented.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-SP-8124
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  • 68
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    Pretest information for a test to validate plume simulation procedures (FA-17) (1978)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The results of an effort to plan a final verification wind tunnel test to validate the recommended correlation parameters and application techniques were presented. The test planning effort was complete except for test site finalization and the associated coordination. Two suitable test sites were identified. Desired test conditions were shown. Subsequent sections of this report present the selected model and test site, instrumentation of this model, planned test operations, and some concluding remarks.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-150638 , RM-024-1
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  • 69
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    Liquid Rocket Lines, Bellows, Flexible Hoses, and Filters (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Fluid-flow components in a liquid propellant rocket engine and the rocket vehicle which it propels are interconnected by lines, bellows, and flexible hoses. Elements involved in the successful design of these components are identified and current technologies pertaining to these elements are reviewed, assessed, and summarized to provide a technology base for a checklist of rules to be followed by project managers in guiding a design or assessing its adequacy. Recommended procedures for satisfying each of the design criteria are included.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-SP-8123
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  • 70
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    Atmospheric scavenging exhaust (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Solid propellant rocket exhaust was directly utilized to ascertain raindrop scavenging rates for hydrogen chloride. The airborne HCl concentration varied from 0.2 to 10.0 ppm and the raindrop sizes tested included 0.55 mm, 1.1 mm, and 3.0 mm. Two chambers were used to conduct the experiments. A large, rigid walled, spherical chamber stored the exhaust constituents while the smaller chamber housing all the experiments was charged as required with rocket exhaust HCl. Surface uptake experiments demonstrated an HCl concentration dependence for distilled water. Sea water and brackish water HCl uptake was below the detection limit of the chlorine-ion analysis technique employed. Plant life HCl uptake experiments were limited to corn and soybeans. Plant age effectively correlated the HCl uptake data. Metallic corrosion was not significant for single 20 minute exposures to the exhaust HCl under varying relative humidity.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-150491 , IITRI-C6365-17
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  • 71
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    Ion engine auxiliary propulsion applications and integration study (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The benefits derived from application of the 8-cm mercury electron bombardment ion thruster were assessed. Two specific spacecraft missions were studied. A thruster was tested to provide additional needed information on its efflux characteristics and interactive effects. A Users Manual was then prepared describing how to integrate the thruster for auxiliary propulsion on geosynchronous satellites. By incorporating ion engines on an advanced communications mission, the weight available for added payload increases by about 82 kg (181 lb) for a 100 kg (2200 lb) satellite which otherwise uses electrothermal hydrazine. Ion engines can be integrated into a high performance propulsion module that is compatible with the multimission modular spacecraft and can be used for both geosynchronous and low earth orbit applications. The low disturbance torques introduced by the ion engines permit accurate spacecraft pointing with the payload in operation during thrusting periods. The feasibility of using the thruster's neutralizer assembly for neutralization of differentially charged spacecraft surfaces at geosynchronous altitude was demonstrated during the testing program.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-135312 , TRW-29999-6013-RU-00
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  • 72
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    Tripropellant Engine Study (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The feasibility of modifying the space shuttle main engine (SSME) for dual mode operation was investigated. Various high power cycle engine configurations derived from the SSME were configurations that will allow sequential burning of LOX/hydrocarbon and LOX/hydrogen were studied in order to identify concepts that make maximum use of SSME hardware and best satisfy the dual mode booster engine system application. Engine cycles were formulated for LOX/RP-1, LOX/CH4, and LOX/C3H8 propellants. Flow rates and operating cycles were established and the adaptability of the major components of the SSME was evaluated.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-150482 , ASR-77-240 , BMPR-2
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  • 73
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    Hydrogen film cooling of a small hydrogen-oxygen thrust chamber and its effect on erosion rates of various ablative materials (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An experimental investigation was conducted to determine what arrangement of film-coolant-injection orifices should be used to decrease the erosion rates of small, high temperature, high pressure ablative thrust chambers without incurring a large penalty in combustion performance. All of the film cooling was supplied through holes in a ring between the outer row of injector elements and the chamber wall. The best arrangement, which had twice the number of holes as there were outer row injection elements, was also the simplest. The performance penalties, presented as a reduction in characteristic exhaust velocity efficiency, were 0.8 and 2.8 percentage points for the 10 and 20 percent cooling flows, respectively, The best film-coolant injector was then used to obtain erosion rates for 19 ablative materials. The throat erosion rate was reduced by a factor of 2.5 with a 10 percent coolant flow. Only the more expensive silica phenolic materials had low enough erosion rates to be considered for use in the nozzle throat. However, some of the cheaper materials might qualify for use in other areas of small nozzles with large throat diameters where the higher erosion rates are more acceptable.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TP-1098 , E-8909
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  • 74
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    Charge-exchange plasma generated by an ion thruster (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The charge exchange plasma generated by an ion thruster was investigated experimentally using both 5 cm and 15 cm thrusters. Results are shown for wide ranges of radial distance from the thruster and angle from the beam direction. Considerations of test environment, as well as distance from the thruster, indicate that a valid simulation of a thruster on a spacecraft was obtained. A calculation procedure and a sample calculation of charge exchange plasma density and saturation electron current density are included.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-135318
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  • 75
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    Monopropellant thruster exhaust plume contamination measurements (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The potential spacecraft contaminants in the exhaust plume of a 0.89N monopropellant hydrazine thruster were measured in an ultrahigh quartz crystal microbalances located at angles of approximately 0 deg, + 15 deg and + or - 30 deg with respect to the nozzle centerline. The crystal temperatures were controlled such that the mass adhering to the crystal surface at temperatures of from 106 K to 256 K could be measured. Thruster duty cycles of 25 ms on/5 seconds off, 100 ms on/10 seconds off, and 200 ms on/20 seconds off were investigated. The change in contaminant production with thruster life was assessed by subjecting the thruster to a 100,000 pulse aging sequence and comparing the before and after contaminant deposition rates. The results of these tests are summarized, conclusions drawn, and recommendations given.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-155270 , JPL-PUB-77-61
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  • 76
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    Extended performance solar electric propulsion thrust system study. Volume 1: Executive summary (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Several thrust system design concepts were evaluated and compared using the specifications of the most advanced 30 cm engineering model thruster as the technology base. The extensions in thruster performance required for the Halley's comet mission were defined and alternative thrust system concepts were designed. Confirmation testing and analysis of thruster and power-processing components were performed, and the feasibility of satisfying extended performance requirements was verified. A baseline design was selected from the alternatives considered, and the design analysis and documentation were refined. A program development plan was formulated that outlines the work structure considered necessary for developing, qualifying, and fabricating the flight hardware for the baseline thrust system within the time frame of a project to rendezvous with Halley's comet. An assessment was made of the costs and risks associated with a baseline thrust system as provided to the mission project under this plan. Critical procurements and interfaces were identified and defined. Results are presented.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-135281-VOL-1
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  • 77
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    Tripropellant engine study (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Work conducted was devoted to three main tasks. Thermochemical equilibrium performance data were assembled to establish the expected performance calculations of the mode 1 engine propellant combinations and thermodynamic and transport data for the products of combustion. Turbine drive gas characteristics were also established. Thrust chamber and nozzle cooling studies were devoted to the evaluation of H2, C3H8, CH4, and RP-1 as coolants in the existing SSME cooling circuit geometry. It was found that all these candidate coolants are feasible without limiting the desired operating conditions with the exception of RP-1, which would limit the maximum P(c) to 2000 psia. RP-1 could be used, however, to cool the nozzle only without imposing the chamber pressure limit. A total of 15 candidate engine system cycles were selected and a preliminary engine system balance was conducted for 12 of these systems to establish component operating flowrates, pressures and temperatures. It was found that the staged combustion cycles employing fuel rich LOX/hydrocarbon turbine drive gases are power limited.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-150444 , ASR-77-213 , BMPR-1
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  • 78
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    Design analogy between optimal time-fuel and rate-ledge relay controllers (1976)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: An empirically obtained reaction control jet relay control law with deadband is used as the basis for determining an equivalent weighted time-fuel optimal switching curve according to a least-squares criterion. The derived transformation from the empirical to the optimal law is found to be reversible and to yield a unique transformed control law. The proposed method provides a basis for determining the behavior of an easily implemented relay control law using well-known optimal control results, as well as determining the equivalent relay law corresponding to an analytically determined optimal control law. A numerical example illustrates the transformation technique and simulation results are presented to compare the two control laws.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Journal of Spacecraft and Rockets; 13; Apr. 197
    Format: text
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  • 79
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    Stability predictions for combustors with acoustic absorbers and continuous combustion distributions (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: The realistic case of a continuous distribution of combustion sources in the axial direction is considered in the investigation. The results obtained are compared with those of an earlier study conducted by Baer et al. (1974) concerning the stability of partially lined combustors with distributed combustion. There is a substantial upward shift of the curves in all cases relative to the curves obtained in the first analysis. The increase in chamber stability indicated is traced to some important damping effects associated with source terms which had been neglected in the previous study.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA Journal; 13; Aug. 197
    Format: text
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  • 80
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    Extended performance solar electric propulsion thrust system study. Volume 5. Capacitor-diode voltage multiplier: Technology evaluation (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A 1-kW capacitor-diode voltage multiplier (CDVM) was designed, fabricated and tested to demonstrate the power of feasibility of high power CDVM's and to verify the analytical techniques that had been used to predict the performance characteristics of a 6-kw CDVM. High efficiency (96.2%), a low ratio of component weight to power (0.55 kg/kW), and low output ripple voltage (less than 1%, peak to peak) were obtained during the operation of a 1-kW CDVM various input line, load current, and load fault conditions.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-135281-VOL-5
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  • 81
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    Tripropellant engine study (1978)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Engine performance data, combustion gas thermodynamic properties, and turbine gas parameters were determined for various high power cycle engine configurations derived from the space shuttle main engine that will allow sequential burning of LOX/hydrocarbon and LOX/hydrogen fuels. Both stage combustion and gas generator pump power cycles were considered. Engine concepts were formulated for LOX/RP-1, LOX/CH4, and LOX/C3H8 propellants. Flowrates and operating conditions were established for this initial set of engine systems, and the adaptability of the major components of shuttle main engine was investigated.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-150562 , ASR-78-17 , BMPR-3
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  • 82
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    Ion Propulsion Module (IPM) technology readiness assessment: 1985 Halley's Comet rendezvous mission (1977)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An assessment of the risk of utilizing ion propulsion to perform a rendezvous mission with Halley's comet in 1985 is presented, and consideration recommendations for reducing identified risks to the lowest possible level at project start in October 1978 were made.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-155308
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  • 83
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    Demonstration of a sterilizable solid rocket motor system (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A solid propellant rocket motor containing 60.9 Kg (134-lb) of propellant was successfully static fired after being subjected to eight heat sterilization cycles (three 54-hour cycles plus five 40-hour cycles) at 125 C (257 F). The test motor, a modified SVM-3 chamber, incorporated a flexible grain retention system of EPR rubber to relieve thermal shrinkage stresses. The propellant used in the motor was ANB-3438, and 84 wt% solids system (18 wt% aluminum) containing 66 wt% stabilized ammonium perchlorate oxidizer and a saturated hydroxylterminated polybutadiene binder. Bonding of the propellant to the EPR insulation (GenGard V-4030) was provided by the use of SD-886, an epoxy urethane restriction.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-2478
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  • 84
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    Results of TC-1 boost pump icing tests in the space power facility (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A series of tests were conducted in the space power facility to investigate the failure of the Centaur oxidizer boost pump during the Titan/Centaur proof flight February 11, 1974. The three basic objectives of the tests were: (1) demonstrate if an evaporative freezing type failure mechanism could have prevented the pump from operating, (2) determine if steam from the exhaust of one of the attitude control engine could have entered a pump seal cavity and caused the failure, and (3) obtain data on the heating effects of the exhaust plume from a hydrogen peroxide attitude control engine.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-71671 , E-8170
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  • 85
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    Phase 2 study of improved materials for use on Scout rocket motor nozzles (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Nozzle material performance data were obtained, and the feasibility was determined of using new materials on the Scout rocket motor nozzles. Stress and heat transfer analyses were conducted to aid in the selection of optimum materials for nozzle tests. A reimpregnated and graphitized throat insert was fabricated along with two nozzles with ablative throats. The dissection and determining of char and erosion of two nozzles fired on X-259 loaded cases are discussed; one of the nozzles used a graphite phenolic ablative throat insert, and the other unit was a standard X-259 nozzle with a reduced area ATJ graphite throat insert.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-132679
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  • 86
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    Heat operated cryogenic electrical generator (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An electrical generator useful for providing electrical power in deep space, is disclosed. The electrical generator utilizes the unusual hydrodynamic property exhibited by liquid helium as it is converted to and from a superfluid state to cause opposite directions of rotary motion for a rotor cell thereof. The physical motion of the rotor cell was employed to move a magnetic field provided by a charged superconductive coil mounted on the exterior of the cell. An electrical conductor was placed in surrounding proximity to the cell to interact with the moving magnetic field provided by the superconductive coil and thereby generate electrical energy. A heat control arrangement was provided for the purpose of causing the liquid helium to be partially converted to and from a superfluid state by being cooled and heated, respectively.
    Keywords: SPACECRAFT PROPULSION AND POWER
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  • 87
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    Space LOX vent system (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: This is the final report summarizing the work completed under contract NAS8-26972. Concept selection, design, fabricating and testing of a prototype compact heat exchanger thermodynamic vent system are discussed. The system is designed to operate in a 2.7m (9 foot) spherical liquid oxygen tank with a heating rate of 32.2 - 35.2 watts (110-120 Btu/hr) and to control pressure to 310 + or - 13.8 kN/sq m (45 + or - 2.0 psia.) the design mission is of 2,590 ks (30 days) duration on board a space shuttle orbiter.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-120769 , CASD-NAS-75-021
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  • 88
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    Acquisition system environmental effects study (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The effects of vibration, warm gas exposure, and feed system startup/shutdown fluid dynamics on capillary-screen propellant retention capabilities are quantified. The existing technology is extended to the point where quantitative conlusions in terms of design criteria may be drawn.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-120768 , MCR-75-21
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  • 89
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    Hydrazine Gas Generator Program (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The design and fabrication of a flight gas generator for the space shuttle were investigated. Critical performance parameters and stability criteria were evaluated as well as a scaling laws that could be applied in designing the flight gas generator. A test program to provide the necessary design information was included. A structural design, including thermal and stress analysis, and two gas generators were fabricated based on the results. Conclusions are presented.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-141795 , R-9690
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  • 90
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    Natural environment design criteria for the Solar Electric Propulsion Stage (SEPS) (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The natural environment design criteria are given for six different solar electric propulsion stage missions. These environment data include the neutral atmosphere; ionosphere, trapped radiation; free-space radiation environment; and meteoroid, asteroid, and comet environments. The electromagnetic radiation environment (direct, reflected, or scattered) at the planets and interplanetary regions is also included.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-64929
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  • 91
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    Liquid rocket engine injectors (1976)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The injector in a liquid rocket engine atomizes and mixes the fuel with the oxidizer to produce efficient and stable combustion that will provide the required thrust without endangering hardware durability. Injectors usually take the form of a perforated disk at the head of the rocket engine combustion chamber, and have varied from a few inches to more than a yard in diameter. This monograph treats specifically bipropellant injectors, emphasis being placed on the liquid/liquid and liquid/gas injectors that have been developed for and used in flight-proven engines. The information provided has limited application to monopropellant injectors and gas/gas propellant systems. Critical problems that may arise during injector development and the approaches that lead to successful design are discussed.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-SP-8089
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  • 92
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    Cusped magnetic field mercury ion thruster (1976)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The importance of a uniform current density profile in the exhaust beam of an electrostatic ion thruster is discussed in terms of thrust level and accelerator system lifetime. A residence time approach is used to explain the nonuniform beam current density profile of the divergent magnetic field thruster. Mathematical expressions are derived which relate the thruster discharge power loss, propellant utilization, and double to single ion density ratio to the geometry and plasma properties of the discharge chamber. These relationships are applied to a cylindrical discharge chamber model of the thruster. Experimental results are presented for a wide range of the discharge chamber length. The thruster designed for this investigation was operated with a cusped magnetic field as well as a divergent field geometry, and the cusped field geometry is shown to be superior from the standpoint of beam profile uniformity, performance, and double ion population.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-135047
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    Space shuttle engineering and operations support. Orbiter to spacelab electrical power interface. Avionics system engineering (1976)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The results are presented of an investigation of the factors which affect the determination of Spacelab (S/L) minimum interface main dc voltage and available power from the orbiter. The dedicated fuel cell mode of powering the S/L is examined along with the minimum S/L interface voltage and available power using the predicted fuel cell power plant performance curves. The values obtained are slightly lower than current estimates and represent a more marginal operating condition than previously estimated.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-147834 , REPT-1.3-DN-C0504-036
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    NASA TECHNICAL REPORTS
  • 94
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    Mariner Jupiter/Saturn LCSSE thruster/valve assembly and injection propulsion unit rocket engine assemblies: 0.2-lbf T/VA development and margin limit test report (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Thruster valve assemblies (T/VA's) were subjected to the development test program for the combined JPL Low-Cost Standardized Spacecraft Equipment (LCSSE) and Mariner Jupiter/Saturn '77 spacecraft (MJS) programs. The development test program was designed to achieve the following program goals: (1) demonstrate T/VA design compliance with JPL Specifications, (2) to conduct a complete performance Cf map of the T/VA over the full operating range of environment, (3) demonstrate T/VA life capability and characteristics of life margin for steady-state limit cycle and momentum wheel desaturation duty cycles, (4) verification of structural design capability, and (5) generate a computerized performance model capable of predicting T/VA operation over pressures ranging from 420 to 70 psia, propellant temperatures ranging from 140 F to 40 F, pulse widths of 0.008 to steady-state operation with unlimited duty cycle capability, and finally predict the transient performance associated with reactor heatup during any given duty cycle, start temperature, feed pressure, and propellant temperature conditions.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-148542 , RRC-76-R-499
    Format: application/pdf
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  • 95
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    Entry reaction control system propellant analysis. Flight control integration (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The effects of winds, sideslip angle feedback, and the data reference (air or inertial) on the Reaction Control System (RCS) propellant requirements during entry were investigated. It was determined that in the presence of a 3 sigma crosswind an addition 188 pounds of RCS propellant was required for entry control which is within the present 200 pound allotment for winds. The absence of air data information does result in slightly higher RCS propellant demands.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-147812
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  • 96
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    Commonality Study of the Pressure-Volume-Temperature Based Propellant Gaging Software Modules for the Auxiliary Power Unit, Reaction Control System, and Orbital Maneuvering System (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Computer storage requirements can be reduced if areas of commonality exist in two or more programs placed in the same computer and identical code can be used by more than one program. The pressure-volume-temperature (P-V-T) relationship for the propellant tank pressurant agent is utilized as the basis for either a primary of a backup propellant gaging program for the auxiliary power unit (APU), the reaction control system (RCS), and the orbital maneuvering system (OMS). These three propellant gaging programs were investigated. It was revealed that a very limited degree of software commonality exits among them. An examination of this common software indicated that only the computation of the helium compressibility factor in an external function subprogram accessible to both the RCS and OMS propellant gaging programs appears to offer a savings in computer storage requirements.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-147808
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  • 97
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    Rocket engine thrust chamber wall temperature distribution calculation and analysis (1976)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An analytical computational concept is presented which predicts the temperature profiles along a regeneratively cooled thrust chamber wall on the hot gas side and on the coolant side, and also the coolant bulk temperature profile. The computational model is based upon a coupling of the boundary layer heat transfer process with the heat transfer process through the chamber wall and the coolant flow heat absorption. The calculation is started with approximate temperature distributions for the hot gas side wall and the coolant flow. The iteration process of the computer program is terminated when the total heat transfer rates from the hot gas boundary layer to the wall and from the wall to the coolant are equal. The computer program for the integration of regenerative cooling process to a thrust chamber is kept general such that this program can be used with any boundary layer analysis computer program for temperature profile and heat transfer studies. A sample application of this concept is shown by using a boundary layer analysis program for the RL10 rocket engine thrust chamber.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-148288
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  • 98
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    Hydrazine engine plume contamination mapping (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Instrumentation for the measurement of plume exhaust specie deposition rates were developed and demonstrated. The instruments, two sets of quartz crystal microbalances, were designed for low temperature operation in the back flow and variable temperature operation in the core flow regions of an exhaust plume. These quartz crystal microbalances performed nominally, and measurements of exhaust specie deposition rates for 8400 number of pulses for a 0.1-lb monopropellant thruster are reported.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-147936 , AD-A018729 , AFRPL-TR-75-16
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  • 99
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    Unified computer codes: Properties data for low cost nozzle materials (1976)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The development of the analytic capability to predict the thermal ablation response of promising low cost materials for rocket nozzles is presented.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-144315 , TR-76-9
    Format: application/pdf
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  • 100
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    Beam efflux measurements (1976)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Measurements have been made of the high energy thrust ions, (Group I), high angle/high energy ions (Group II), and high angle/low energy ions (Group IV) of a mercury electron bombardment thruster in the angular divergence range from 0 deg to greater than 90 deg. The measurements have been made as a function of thrust ion current, propellant utilization efficiency, bombardment discharge voltage, screen and accelerator grid potential (accel-decel ratio) and neutralizer keeper potential. The shape of the Group IV (charge exchange) ion plume has remained essentially fixed within the range of variation of the engine operation parameters. The magnitude of the charge exchange ion flux scales with thrust ion current, for good propellant utilization conditions. For fixed thrust ion current, charge exchange ion flux increases for diminishing propellant utilization efficiency. Facility effects influence experimental accuracies within the range of propellant utilization efficiency used in the experiments. The flux of high angle/high energy Group II ions is significantly diminished by the use of minimum decel voltages on the accelerator grid. A computer model of charge exchange ion production and motion has been developed. The program allows computation of charge exchange ion volume production rate, total production rate, and charge exchange ion trajectories for "genuine" and "facilities effects" particles. In the computed flux deposition patterns, the Group I and Group IV ion plumes exhibit a counter motion.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-135038
    Format: application/pdf
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