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  • Other Sources  (173)
  • SPACECRAFT PROPULSION AND POWER  (173)
  • Cell & Developmental Biology
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  • 1980-1984
  • 1975-1979  (173)
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  • 1975  (173)
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  • 1980-1984
  • 1975-1979  (173)
  • 1960-1964
  • 1925-1929
  • 1915-1919
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  • 1
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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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  • 2
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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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  • 3
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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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  • 4
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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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  • 5
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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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  • 6
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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
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  • 7
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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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  • 8
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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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  • 9
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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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  • 10
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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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  • 11
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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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  • 12
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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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  • 13
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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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  • 14
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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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  • 15
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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
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  • 16
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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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  • 17
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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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  • 18
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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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  • 19
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    Space shuttle SRM plume expansion sensitivity analysis (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The exhaust plumes of the space shuttle solid rocket motors can have a significant effect on the base pressure and base drag of the shuttle vehicle. A parametric analysis was conducted to assess the sensitivity of the initial plume expansion angle of analytical solid rocket motor flow fields to various analytical input parameters and operating conditions. The results of the analysis are presented and conclusions reached regarding the sensitivity of the initial plume expansion angle to each parameter investigated. Operating conditions parametrically varied were chamber pressure, nozzle inlet angle, nozzle throat radius of curvature ratio and propellant particle loading. Empirical particle parameters investigated were mean size, local drag coefficient and local heat transfer coefficient. Sensitivity of the initial plume expansion angle to gas thermochemistry model and local drag coefficient model assumptions were determined.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-147528 , LMSC-HREC-TM-D496636
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  • 20
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    Design of multi-mission chemical propulsion modules for planetary orbiters. Volume 2: Technical report (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: For abstract, see N76-14196.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-137790 , TRW-26085-6001-TU-00-VOL-2
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  • 21
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    TC-2 post Helios experiment data review (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Data are presented from a systems postflight analysis of the Centaur Launch Vehicle and Helios. Also given is a comparison of data from preflight analyses. Topics examined are: (1) propellant behavior; (2) helium usage; (3) propellant tank pressurization; (4) propellant tank thermodynamics; (5) component heating; thermal control; and thermal protection system; (6) main engine system; (7) H2O2 consumption; (8) boost pump post-meco performance; and (9) an overview of other systems.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-148459
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  • 22
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    Study of cryogenic propellant systems for loading the space shuttle. Part 2: Hydrogen systems (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Computer simulation studies of liquid hydrogen fill and vent systems for the space shuttle are studied. The computer programs calculate maximum and minimum permissible flow rates during cooldown as limited by thermal stress considerations, fill line cooldown time, pressure drop, flow rates, vapor content, vent line pressure drop and vent line discharge temperature. The input data for these programs are selected through graphic displays which schematically depict the part of the system being analyzed. The computed output is also displayed in the form of printed messages and graphs. Digital readouts of graph coordinates may also be obtained. Procedures are given for operation of the graphic display unit and the associated minicomputer and timesharing computer.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-148508 , NBSIR-75-820
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  • 23
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    Solid rocket motor nozzles (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The steps in the nozzle design process are examined. The nozzle designer's role in defining design requirements and constraints is included along with discussions of each of the three basic phases of the nozzle design process itself: (1) aerodynamic design, in which the gas-contacting surfaces are configured to produce the required performance within the envelope limits; (2) thermal design, in which termal liners and thermal insulators are selected and configured to maintain the surfaces as closely as practical against effects of erosion and to limit the structure temperature to acceptable levels; and (3) structural design, in which materials are selected and configured to support the thermal components and to sustain the predicted loads. Analytical techniques that are used to establish thermal and structural design integrity and to predict nozzle performance are discussed along with methods for nozzle quality assurance. Emphasis is placed on nozzle design and materials for modern high-temperature aluminized propellants. Recurring nozzle design problems of graphite cracking and ejection, differential erosion at material interfaces, lack of sufficient proven nondestructive testing (NDT) techniques, the uncertainty of adhesive bonding, and inadequate definition of material properties, particularly at high temperatures are considered.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-SP-8115
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  • 24
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    Turbulence in a gaseous hydrogen-liquid oxygen rocket combustion chamber (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The intensity of turbulence and the Lagrangian correlation coefficient for a LOX-GH2 rocket combustion chamber was determined from experimental measurements of tracer gas diffusion. A combination of Taylor's turbulent diffusion theory and a numerical method for solving the conservation equations of fluid mechanics was used to calculate these quantities. Taylor's theory was extended to consider the inhomogeneity of the turbulence field in the axial direction of the combustion chamber, and an exponential function was used to represent the Lagrangian correlation coefficient. The results indicate that the value of the intensity of turbulence reaches a maximum of 14% at a location about 7" downstream from the injector. The Lagrangian correlation coefficient associated with this value is given by the above exponential expression where alpha = 10,000/sec.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-134909
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  • 25
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    Nuclear energy waste-space transportation and removal (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A method for utilizing the decay heat of actinide wastes to power an electric thrust vehicle is proposed. The vehicle, launched by shuttle to earth orbit and to earth escape by a tug, obtains electrical power from the actinide waste heat by thermionic converters. The heavy gamma ray and neutron shielding which is necessary as a safety feature is removed in orbit and returned to earth for reuse. The problems associated with safety are dealt with in depth. A method for eliminating fission wastes via chemical propulsion is briefly discussed.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-64973
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  • 26
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    Design of multi-mission chemical propulsion modules for planetary orbiters. Volume 3: Appendixes (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: For abstract, see N76-14196.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-137791 , TRW-26085-6001-TU-00-VOL-3
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  • 27
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    Design of multi-mission chemical propulsion modules for planetary orbiters. Volume 1: Summary report (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Results are presented of a conceptual design and feasibility study of chemical propulsion stages that can serve as modular propulsion units, with little or no modification, on a variety of planetary orbit missions, including orbiters of Mercury, Saturn, and Uranus. Planetary spacecraft of existing design or currently under development, viz., spacecraft of the Pioneer and Mariner families, are assumed as payload vehicles. Thus, operating requirements of spin-stabilized and 3-axis stabilized spacecraft have to be met by the respective propulsion module designs. As launch vehicle for these missions the Shuttle orbiter and interplanetary injection stage, or Tug, plus solid-propellant kick motor was assumed. Accommodation constraints and interfaces involving the payloads and the launch vehicle are considered in the propulsion module design. The applicability and performance advantages were evaluated of the space-storable high-energy bipropellants. The incentive for using this advanced propulsion technology on planetary missions is the much greater performance potential when orbit insertion velocities in excess of 4 km/sec are required, as in the Mercury orbiter. Design analyses and performance tradeoffs regarding earth-storable versus space-storable propulsion systems are included. Cost and development schedules of multi-mission versus custom-designed propulsion modules are examined.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-137789 , TRW-26085-6001-TU-00-VOL-1
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  • 28
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    Space shuttle orbit maneuvering engine reusable thrust chamber program (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Reusable thrust chamber and injector concepts were evaluated for the space shuttle orbit maneuvering engine (OME). Parametric engine calculations were carried out by computer program for N2O4/amine, LOX/amine and LOX/hydrocarbon propellant combinations for engines incorporating regenerative cooled and insulated columbium thrust chambers. The calculation methods are described including the fuel vortex film cooling method of combustion gas temperature control, and performance prediction. A method of acceptance of a regeneratively cooled heat rejection reduction using a silicone oil additive was also demonstrated by heated tube heat transfer testing. Regeneratively cooled thrust chamber operation was also demonstrated where the injector was characterized for the OME application with a channel wall regenerative thrust chamber. Bomb stability testing of the demonstration chambers/injectors demonstrated recovery for the nominal design of acoustic cavities. Cavity geometry changes were also evaluated to assess their damping margin. Performance and combustion stability was demonstrated of the originally developed 10 inch diameter combustion pattern operating in an 8 inch diameter thrust chamber.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-144632 , BAC-8693-950001
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  • 29
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    Research study on multi-KW DC distribution system (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Power distribution system noise and transient stress on switchgear in large space vehicle power systems were investigated in terms of the effect of flight designs of long power distribution cables on load interface EMI requirements. A fifty meter cable pair was simulated to study interactions between the cable, load, and power source terminations. Power system noise characteristics were evaluated based on current spacecraft data, interface hardware filter designs, and power cable parameters. Parametric approaches were defined for evaluating switching transients at various distribution voltage levels. It is concluded that the state-of-the-art semiconductor switches represent a viable approach toward the implementation of power system design with distribution voltages of 120 VDC or less. The interface definition and design for the bus control unit was updated to be consistent with the established requirements.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-144091
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  • 30
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    Analysis of liquid rocket combustion chamber turbulence levels from diffusion data (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: The intensity of turbulence and the Lagrangian correlation coefficient in a liquid-rocket combustion chamber have been analytically determined from an analysis of experimental diffusion data obtained in a small rocket engine which operated at 300-psia chamber pressure and produced approximately 250 pounds thrust. Results of gas-sample measurements obtained by Orsat and gas-chromatograph techniques to determine helium-concentration profiles were analyzed on the basis of Taylor's (1921) turbulent diffusion theory to obtain turbulence flow-field parameters. The results of the analysis indicate that turbulent diffusion in a combustion chamber can be adequately modeled by the one-dimensional Taylor theory, which assumes that the intensity of turbulence is a function only of axial distance in the chamber and that the Lagrangian correlation coefficient is expressed by a power law. The results indicate a higher intensity of turbulence and lower correlation than previously expected.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Combustion and Flame; 25; Oct. 197
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    Application of blast wave theory to explosive propulsion (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: An analysis was carried out by using blast wave theory to delineate the important aspects of detonating explosives in nozzles, such as flow and wave phenomena, characteristic length and time scales, and the parameters on which the specific impulse is dependent. The propulsive system utilizes the momentum of the ambient gas set into motion in the nozzle by the explosion. A somewhat simplified model was considered for the situation where the mass of ambient gas in the nozzle is much greater than the mass of gas produced in the explosion, a condition of interest for dense atmospheres, e.g., near the surface of Venus. Instantaneous detonation and energy release was presumed to occur at the apex of a conical nozzle, and the shock wave generated by the explosion was taken to propagate as a spherical wave, thereby setting the ambient gas in the nozzle into one-dimensional radially outward motion.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Acta Astronautica; 2; May-June
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  • 32
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    Space shuttle orbit maneuvering engine reusable thrust chamber program (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The feasibility of potential reusable thrust chamber concepts is studied. Propellant condidates were examined and analytically combined with potential cooling schemes. A data base of engine data which would assist in a configuration selection was produced. The data base verification was performed by the demonstration of a thrust chamber of a selected coolant scheme design. A full scale insulated columbium thrust chamber was used for propellant coolant configurations. Combustion stability of the injectors and a reduced size thrust chamber were experimentally verified as proof of concept demonstrations of the design and study results.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-147859 , BAC-8693-950002
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  • 33
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    Orbital Maneuvering Engine Feed System Coupled Stability Investigation, Computer User's Manual (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An operating manual for the feed system coupled stability model was given, in partial fulfillment of a program designed to develop, verify, and document a digital computer model that can be used to analyze and predict engine/feed system coupled instabilities in pressure-fed storable propellant propulsion systems over a frequency range of 10 to 1,000 Hz. The first section describes the analytical approach to modelling the feed system hydrodynamics, combustion dynamics, chamber dynamics, and overall engineering model structure, and presents the governing equations in each of the technical areas. This is followed by the program user's guide, which is a complete description of the structure and operation of the computerized model. Last, appendices provide an alphabetized FORTRAN symbol table, detailed program logic diagrams, computer code listings, and sample case input and output data listings.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-150944 , R-9808
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  • 34
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    Solid-propellant rocket motor ballistic performance variation analyses (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Results are presented of research aimed at improving the assessment of off-nominal internal ballistic performance including tailoff and thrust imbalance of two large solid-rocket motors (SRMs) firing in parallel. Previous analyses using the Monte Carlo technique were refined to permit evaluation of the effects of radial and circumferential propellant temperature gradients. Sample evaluations of the effect of the temperature gradients are presented. A separate theoretical investigation of the effect of strain rate on the burning rate of propellant indicates that the thermoelastic coupling may cause substantial variations in burning rate during highly transient operating conditions. The Monte Carlo approach was also modified to permit the effects on performance of variation in the characteristics between lots of propellants and other materials to be evaluated. This permits the variabilities for the total SRM population to be determined. A sample case shows, however, that the effect of these between-lot variations on thrust imbalances within pairs of SRMs is minor in compariosn to the effect of the within-lot variations. The revised Monte Carlo and design analysis computer programs along with instructions including format requirements for preparation of input data and illustrative examples are presented.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-144264
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  • 35
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    Space shuttle orbital maneuvering engine platelet injector program (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A platelet-face injector for the fully reusable orbit maneuvering system OMS on the space shuttle was evaluated as a means of obtaining additional design margin and low cost. Performance, heat transfer, and combustion stability were evaluated over the anticipated range of OMS operating conditions. The effects of acoustic cavity configuration on combustion stability, including cavity depth, open area, inlet contour, and other parameters, were investigated using sea level bomb tests. Prototype injector and chamber behavior was evaluated for a variety of conditions; these tests examined the effects of film cooling, helium saturated propellants, chamber length, inlet conditions, and operating point, on performance, heat transfer and engine transient behavior. Helium bubble ingestion into both propellant circuits was investigated, as was chugging at low pressure operation, and hot and cold engine restart with and without a purge.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-147480 , REPT-13133-F-2
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  • 36
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    Chemically Reacting One-Dimensional Gas-Particle Flows (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The governing equations for the one-dimensional flow of a gas-particle system are discussed. Gas-particle effects are coupled via the system momentum and energy equations with the gas assumed to be chemically frozen or in chemical equilibrium. A computer code for calculating the one-dimensional flow of a gas-particle system is discussed and a user's input guide presented. The computer code provides for the expansion of the gas-particle system from a specified starting velocity and nozzle inlet geometry. Though general in nature, the final output of the code is a startline for initiating the solution of a supersonic gas-particle system in rocket nozzles. The startline includes gasdynamic data defining gaseous startline points from the nozzle centerline to the nozzle wall and particle properties at points along the gaseous startline.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-147388 , LMSC-HREC-TR-D390876
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  • 37
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    Investigation of the Centaur boost pump overspeed condition at main engine shutdown on the Titan Centaur TC-2 flight (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An investigation was conducted to evaluate a potential boost pump overspeed condition which could exist on the Titan/Centaur launch vehicle after main engine shut-off. Preliminary analyses indicated that the acceleration imparted to the unloaded boost pump-turbine assembly, caused by purging residual hydrogen peroxide from the turbine supply lines, could result in a pump-turbine overspeed. Previous test experience indicated that turbine damage occurs at speeds in excess of 75,000 rpm. Detailed theoretical analyses, in conjunction with pump tests, were conducted to establish the maximum pump-turbine speed at main engine shut-off. The analyses predicted a maximum speed of 68,000 rpm. Testing showed the pump-turbine speed to be 66,700 rpm in the overspeed condition. Inasmuch as both the analysis and tests showed the overspeed to be sufficiently less than the speed at which damage could occur, it was concluded that no corrective action would be required for the launch vehicle.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-71822 , E-8521
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  • 38
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    Feasibility of an advanced thrust termination assembly for a solid propellant rocket motor (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A total of 68 quench tests were conducted in a vented bomb assembly (VBA). Designed to simulate full-scale motor operating conditions, this laboratory apparatus uses a 2-inch-diameter, end-burning propellant charge and an insulated disc of consolidated hydrated aluminum sulfate along with the explosive charge necessary to disperse the salt and inject it onto the burning surface. The VBA was constructed to permit variation of motor design parameters of interest; i.e., weight of salt per unit burning surface area, weight of explosive per unit weight of salt, distance from salt surface to burning surface, incidence angle of salt injection, chamber pressure, and burn time. Completely satisfactory salt quenching, without re-ignition, occurred in only two VBA tests. These were accomplished with a quench charge ratio (QCR) of 0.023 lb salt per square inch of burning surface at dispersing charge ratios (DCR) of 13 and 28 lb of salt per lb of explosive. Candidate materials for insulating salt charges from the rocket combustion environment were evaluated in firings of 5-inch-diameter, uncured end-burner motors. A pressed, alumina ceramic fiber material was selected for further evaluation and use in the final demonstration motor.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-146312 , E41-75
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  • 39
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    Comparative assessment of out-of-core nuclear thermionic power systems (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The hardware selections available for fabrication of a nuclear electric propulsion stage for planetary exploration were explored. The investigation was centered around a heat-pipe-cooled, fast-spectrum nuclear reactor for an out-of-core power conversion system with sufficient detail for comparison with the in-core system studies completed previously. A survey of competing power conversion systems still indicated that the modular reliability of thermionic converters makes them the desirable choice to provide the 240-kWe end-of-life power for at least 20,000 full power hours. The electrical energy will be used to operate a number of mercury ion bombardment thrusters with a specific impulse in the range of about 4,000-5,000 seconds.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-145972 , JPL-TM-33-749
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  • 40
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    Effect of Nozzle Nonlinearities upon Nonlinear Stability of Liquid Propellant Rocket Motors (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A three dimensional, nonlinear nozzle admittance relation is developed by solving the wave equation describing finite amplitude oscillatory flow inside the subsonic portion of a choked, slowly convergent axisymmetric nozzle. This nonlinear nozzle admittance relation is then used as a boundary condition in the analysis of nonlinear combustion instability in a cylindrical liquid rocket combustor. In both nozzle and chamber analyses solutions are obtained using the Galerkin method with a series expansion consisting of the first tangential, second tangential, and first radial modes. Using Crocco's time lag model to describe the distributed unsteady combustion process, combustion instability calculations are presented for different values of the following parameters: (1) time lag, (2) interaction index, (3) steady-state Mach number at the nozzle entrance, and (4) chamber length-to-diameter ratio. In each case, limit cycle pressure amplitudes and waveforms are shown for both linear and nonlinear nozzle admittance conditions. These results show that when the amplitudes of the second tangential and first radial modes are considerably smaller than the amplitude of the first tangential mode the inclusion of nozzle nonlinearities has no significant effect on the limiting amplitude and pressure waveforms.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-134880
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  • 41
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    Solid rocket technology advancements for space tug and IUS applications (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: In order for the shuttle tug or interim upper stage (IUS) to capture all the missions in the current mission model for the tug and the IUS, an auxiliary or kick stage, using a solid propellant rocket motor, is required. Two solid propellant rocket motor technology concepts are described. One concept, called the 'advanced propulsion module' motor, is an 1800-kg, high-mass-fraction motor, which is single-burn and contains Class 2 propellent. The other concept, called the high energy upper stage restartable solid, is a two-burn (stop-restartable on command) motor which at present contains 1400 kg of Class 7 propellant. The details and status of the motor design and component and motor test results to date are presented, along with the schedule for future work.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-145561 , JPL-TM-33-746
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  • 42
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    Linear rocket engines for advanced space transportation systems (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-07-27
    Description: The linear rocket engine is shown to be a viable candidate propulsion system for post-Space Shuttle single-stage-to-orbit systems. The linear engine system has been developed and fired demonstrating high performance and long life with firing durations exceeding 500 seconds. The application of the split or dual combustor to the linear engine permits the uses of two different propellant combinations in a single engine system. The split combustor possesses the advantages of the two position extendible bell nozzle in a fixed nozzle configuration. Engine power cycles and applications to typical vehicles are discussed.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 75-1251
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  • 43
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    Propulsion technology needs for advanced space transportation systems (1975)
    In:  Other Sources
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    Publication Date: 2019-07-27
    Description: Plans have been formulated for chemical propulsion technology programs to meet the needs of advanced space transportation systems during the two decades from 1980 to the year 2000. The many possible vehicle applications have been reviewed and cataloged to isolate the common threads of primary propulsion technology that will satisfy near term requirements in the first decade and at the same time establish the technology groundwork for various potential far term applications in the second decade. Two thrust classes of primary propulsion engines are apparent: (1) 5,000 to 30,000 pounds thrust for upper stages and space maneuvering; (2) large booster engines of over 250,000 pounds thrust. Six major classes of propulsion systems and the important subdivisions of each class have been identified. The relative importance of each class is discussed in terms of the number of potential applications, the likelihood of that application materializing, and the criticality of the technology needed. Specific technology programs are described and scheduled to fulfill the anticipated primary propulsion technology requirements of the period.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 75-1246
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  • 44
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    Propellant reorientation with off-axis accelerations (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-07-27
    Description: Reorientation of propellant within a tank is important to the design of many acquisition and vent systems. A drop tower test program was conducted to determine the influence of off-axis accelerations on the motion of the liquid. A series of 39 tests was performed with a spherical tank and a cylindrical tank with hemispherical end domes, varying tank orientation, liquid volume and the magnitude of the acceleration. It was shown that reorientation under the effect of a pure axial acceleration, which has been studied in prior test programs, is a special case. The lateral component of the acceleration caused the liquid to reorient along one side of the tank. The motion of the liquid appeared to be independent of the magnitude of the lateral acceleration for the range of values considered.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 75-1195
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  • 45
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    Centaur propellant acquisition system (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-07-27
    Description: The desirability of replacing the hydrogen peroxide settling system of the Centaur D-1S with a capillary acquisition system was evaluated. A comprehensive screening was performed to select the most promising capillary device fluid acquisition, thermal conditioning, and fabrication techniques. Refillable start baskets and bypass feed start tanks were selected for detailed design. Critical analysis areas were settling and refilling, start sequence development with an initially dry boost pump, and cooling the fluid delivered to the boost pump to provide the necessary net positive suction head (NPSH). Design drawings were prepared for start basket and start tank concepts for both the liquid oxygen and liquid hydrogen tanks. System comparisons indicated that the start baskets using wicking flow for thermal conditioning, and thermal subcooling for providing boost pump NPSH, are the most desirable systems for future Centaur acquisition system development.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 75-1194
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  • 46
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    Surface tension propellant acquisition system technology for Space Shuttle reaction control tanks (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-07-27
    Description: A program was conducted to provide the technology base for the SS/RCS flight tankage. Through a combination of analysis, subscale testing and computer predictions, a surface tension acquisition/expulsion system design was developed for the Orbiter RCS application. A full-scale tank system was fabricated and ground verification testing was conducted. Cleaning, inspection, fill and drain, and one-g expulsion performance were demonstrated. Results show that the fine-mesh screen, compartmented tank system provides the performance, flexibility, reusability, and other characteristics required by the pulsing, high flowrate RCS. It provides the required expulsion under widely differing high-g boost abort and reentry vectors oriented 119 deg apart and during on-orbit operation under omnidirectional low-g conditions.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 75-1196
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  • 47
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    Thrust vector control for the Space Shuttle Solid Rocket Motor (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-07-27
    Description: Thrust vector control (TVC) for the Space Shuttle Solid Rocket Motor (SRM) is obtained by omniaxis vectoring of the nozzle. The development and integration of the system are under the cognizance of Marshall Space Flight Center (MSFC). The nozzle and flexible bearing have been designed and will be built by Thiokol Corporation/Wasatch Division. The vector requirements of the system, the impact of multiple reuse on the components, and the unique problems associated with a large flexible bearing are discussed. The design details of each of the major TVC subcomponents are delineated. The subscale bearing development program and the overall development schedule also are presented.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 75-1172
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  • 48
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    Flexible joints for thrust vector control (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-07-27
    Description: Flexible joints have been used to achieve thrust vector control over a wide range of sizes of nozzles and have been demonstrated successfully in bench tests and static firings, and are operational on two motors. From these many joints the problems of flexible joints have been defined as establishment of the movable nozzle envelope, definition of the actuation power requirements, definition of the mechanical properties of joint materials, adhesive bonding, test methods, and quality control. These data and problem solutions are contained in a large number of reports. Data relating to joint configuration, design requirements, materials selection, joint design, structural analysis, manufacture, and testing are summarized.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 75-1221
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  • 49
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    Thrust vector control by liquid injection for solid propellant rockets (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-07-27
    Description: In liquid injection thrust vector control, a rocket jet is deflected for steering purposed by injecting a liquid into the nozzle exit cone. The liquid is preferably both dense and reactive so that it adds mass and energy and generates shocks in the supersonic exhaust. This behavior increases thrust in the affected part of the jet producing not only a side force for steering but an addition to axial thrust. This paper presents a summary of current liquid injection thrust vector control technology, including procedures for design, development, analysis, testing and evaluation, together with supporting data and references.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 75-1225
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  • 50
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    Development and parametric evaluation of the prototype 2 and 3 flash evaporators (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Development of the Prototype 2 and 3 flash evaporator heat sinks which vaporize an expendable fluid to cool a heat transport fluid loop is reported. The units utilize Freon 21 as the heat transport fluid and water as the expendable fluid to meet the projected performance requirements of the space shuttle for both on-orbit and ascent/reentry operations. The evaporant is pulse-sprayed by on-off control onto heat transfer surfaces containing the transport fluid and exhausted to the vacuum environment through fixed area exhaust ducts.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-144446 , T157-76
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  • 51
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    Accelerated life test of sputtering and anode deposit spalling in a small mercury ion thruster (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Tantalum and molybdenum sputtered from discharge chamber components during operation of a 5 centimeter diameter mercury ion thruster adhered much more strongly to coarsely grit blasted anode surfaces than to standard surfaces. Spalling of the sputtered coating did occur from a coarse screen anode surface but only in flakes less than a mesh unit long. The results were obtained in a 200 hour accelerated life test conducted at an elevated discharge potential of 64.6 volts. The test approximately reproduced the major sputter erosion and deposition effects that occur under normal operation but at approximately 75 times the normal rate. No discharge chamber component suffered sufficient erosion in the test to threaten its structural integrity or further serviceability. The test indicated that the use of tantalum-surfaced discharge chamber components in conjunction with a fine wire screen anode surface should cure the problems of sputter erosion and sputtered deposits spalling in long term operation of small mercury ion thrusters.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-3269 , E-8152
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  • 52
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    Monopropellant engine investigation for space shuttle reaction control system. Volume 3: Improvement of metal foam for catalyst retention (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The retention of granular catalyst in a metal foam matrix was demonstrated to greatly increase the life capability of hydrazine monopropellant reactors. Since nickel foam used in previous tests was found to become degraded after long-term exposure the cause of degradation was examined and metal foams of improved durability were developed. The most durable foam developed was a rhodium-coated nickel foam. An all-platinum foam was found to be incompatible in a hot ammonia (hydrazine) environment. It is recommended to scale up the manufacturing process for the improved foam to produce samples sufficiently large for space shuttle APU gas generator testing.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-144417 , REPT-75-R-460-VOL-3
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  • 53
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    Monopropellant engine investigation for space shuttle reaction control. Volume 2: Design, fabrication, and demonstration test of a catalytic gas generator for the space shuttle APU (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The capability of a catalytic gas generator to meet the requirement specified for the space shuttle APU is established. A full-scale gas generator, designed to operate at a chamber pressure of 750 psia and a flow rate of 0.36 lbm/sec, was fabricated and subjected to three separate life test series. The nickel foam metal used for catalyst retention was investigated. Inspection of the foam metal following the first life test revealed significant degradation. Consequently an investigation was conducted to determine the mechanism of degradation and to provide an improved foam metal.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-144416 , REPT-75-R-460-VOL-2
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  • 54
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    Monopropellant engine investigation for space shuttle reaction control system, volume 1 (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The results are presented of an investigation to determine the capability of a monopropellant hydrazine thruster to meet the requirements specified for the space shuttle reaction control system (RCS). Of those requirements, the major concern was whether the 100,000 seconds life could be achieved at thrust levels within the specified range. Although burn times in excess of 200,000 seconds have been demonstrated at low thrust levels, the corresponding total impulse values have been substantially lower than that required for the space shuttle RCS. Two other areas of concern, involving the catalyst, were: (1) the effects of the relatively high vehicle vibration levels on catalyst attrition and (2) the effect of exposure of the catalyst to air during atmospheric reentry of the vehicle.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-144415 , REPT-75-R-460-VOL-1
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  • 55
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    Orbital maneuvering engine feed system coupled stability investigation (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A digital computer model used to analyze and predict engine feed system coupled instabilities over a frequency range of 10 to 1000 Hz was developed and verified. The analytical approach to modeling the feed system hydrodynamics, combustion dynamics, chamber dynamics, and overall engineering model structure is described and the governing equations in each of the technical areas are presented. This is followed by a description of the generalized computer model, including formulation of the discrete subprograms and their integration into an overall engineering model structure. The operation and capabilities of the engineering model were verified by comparing the model's theoretical predictions with experimental data from an OMS-type engine with a known feed system/engine chugging history.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-144409 , R-9807 , MA-129
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  • 56
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    H2O2 space shuttle APU (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A cryogenic H2-O2 auxiliary power unit (APU) was developed and successfully demonstrated. It has potential application as a minimum weight alternate to the space shuttle baseline APU because of its (1) low specific propellant consumption and (2) heat sink capabilities that reduce the amount of expendable evaporants. A reference system was designed with the necessary heat exchangers, combustor, turbine-gearbox, valves, and electronic controls to provide 400 shp to two aircraft hydraulic pumps. Development testing was carried out first on the combustor and control valves. This was followed by development of the control subsystem including the controller, the hydrogen and oxygen control valves, the combustor, and a turbine simulator. The complete APU system was hot tested for 10 hr with ambient and cryogenic propellants. Demonstrated at 95 percent of design power was 2.25 lb/hp-hr. At 10 percent design power, specific propellant consumption was 4 lb/hp-hr with space simulated exhaust and 5.2 lb/hp-hr with ambient exhaust. A 10 percent specific propellant consumption improvement is possible with some seal modifications. It was demonstrated that APU power levels could be changed by several hundred horsepower in less than 100 msec without exceeding allowable turbine inlet temperatures or turbine speed.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-134784 , AIRESEARCH-75-11290
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    Low cost computer subsystem for the Solar Electric Propulsion Stage (SEPS) (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The Solar Electric Propulsion Stage (SEPS) subsystem which consists of the computer, custom input/output (I/O) unit, and tape recorder for mass storage of telemetry data was studied. Computer software and interface requirements were developed along with computer and I/O unit design parameters. Redundancy implementation was emphasized. Reliability analysis was performed for the complete command computer sybsystem. A SEPS fault tolerant memory breadboard was constructed and its operation demonstrated.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-120671 , IBM-75W-00041
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  • 58
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    Multi-kilowatt modularized spacecraft power processing system development (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A review of existing information pertaining to spacecraft power processing systems and equipment was accomplished with a view towards applicability to the modularization of multi-kilowatt power processors. Power requirements for future spacecraft were determined from the NASA mission model-shuttle systems payload data study which provided the limits for modular power equipment capabilities. Three power processing systems were compared to evaluation criteria to select the system best suited for modularity. The shunt regulated direct energy transfer system was selected by this analysis for a conceptual design effort which produced equipment specifications, schematics, envelope drawings, and power module configurations.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-134814 , GE-SD-75SDS4242
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  • 59
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    Space shuttle orbit maneuvering engine (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Data on the performance, stability, and thermal characteristics of an OME operating with an alternate injector configuration and with alternate propellants was obtained. The design, manufacturing, and operating characteristics of an electroformed, regeneratively cooled thrust chamber were also derived. Subscale and full scale tests provide data relating to off-design and transient operation.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-144440 , R-9686
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  • 60
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    Space shuttle orbit maneuvering engine reusable thrust chamber program (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Analyses and preliminary designs of candidate OME propellant combinations and corresponding engine designs were conducted and evaluated in terms of performance, operating limits, program cost, risk, inherent life and maintainability. For the Rocketdyne recommended and NASA approved propellant combination and cooling concept (NTO/MMH regeneratively cooled engine), a demonstration thrust chamber was designed, fabricated, and experimentally evaluated to define operating characteristics and limits. Alternate fuel (50-50) operating characteristics were also investigated with the demonstration chamber. Adverse operating effects on regenerative cooled operation were evaluated using subscale electrically heated tubes and channels. An investigation of like doublet element characteristics using subscale tests was performed. Full scale 8- and 10-inch diameter like-doublet injectors for the OME were designed, fabricated, and tested. Injector stability was evaluated analytically and experimentally.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-144441 , R-9686-1
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  • 61
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    System reliability analysis through corona testing (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A corona vacuum test facility for nondestructive testing of power system components was built in the Reliability and Quality Engineering Test Laboratories at the NASA Lewis Research Center. The facility was developed to simulate operating temperature and vacuum while monitoring corona discharges with residual gases. The facility is being used to test various high-voltage power system components.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-3287 , E-8246
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  • 62
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    Actuation and system design and evaluation OMS engine shutoff valve, Volume 2 (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: For abstract, see N75-27062.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-141875 , EER-5739000-VOL-2
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  • 63
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    Actuation and system design and evaluation OMS engine shutoff valve, Volume 1 (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A technology program was conducted to identify and verify the optimum valve and actuation system concept for the Space Shuttle Orbit Maneuvering System engine. Of major importance to the valve and actuation system selection was the ten-year, 100-mission, 10,000-cycle life requirement, while maintaining high reliability, low leakage, and low weight. Valve and actuation system concepts were comparatively evaluated against past valve failure reports and potential failure modes due to the shuttle mission profile to aid in the selection of the most optimum concept for design, manufacture and verification testing. Two valve concepts were considered during the preliminary design stage; i.e., the moving seat and lifting ball. Two actuation systems were manufactured and tested. Test results demonstrate the viability of a lifting ball concept as well as the applicability of an ac motor actuation system to best meet the requirements of the shuttle mission.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-141874 , EER-5739000-VOL-1
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  • 64
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    Low-speed inducers for cryogenic upper-stage engines (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Two-phase, low-speed hydrogen and oxygen inducers driven by electric motors and applicable to the tug engine were designed and constructed. The oxygen inducer was tested in liquid and two-phase oxygen. Its head and flow performance were approximately as designed, and it was able to accelerate to full speed in 3 seconds and produce its design flow and head. The analysis of the two-phase data indicated that the inducer was able to pump with vapor volume fractions in excess of 60 percent. The pump met all of its requirements (duration of runs and number of starts) to demonstrate its mechanical integrity.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-143872 , R-9655
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  • 65
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    Nozzle exit exhaust products from space shuttle boost vehicle (November 1973 design) (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Principal exhaust species emitted at various altitudes for two trajectories of the space shuttle vehicle are presented. The exhaust composition is given for the nozzle exit plane on the basis of equilibrium chemistry. Afterburning of excess H, H2, and CO in the plume is accounted for. Species considered include HCl and Al2O3, which have been recognized as environmentally significant, as well as others such as H2O (produced by both the solid rocket motor and the orbiter main engine) which, although innocuous, may participate in subsequent chemical reactions in the atmosphere.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-136747 , JPL-TM-33-712
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  • 66
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    Frontiers in propulsion research: Laser, matter-antimatter, excited helium, energy exchange thermonuclear fusion (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Concepts are described that presently appear to have the potential for propulsion applications in the post-1990 era of space technology. The studies are still in progress, and only the current status of investigation is presented. The topics for possible propulsion application are lasers, nuclear fusion, matter-antimatter annihilation, electronically excited helium, energy exchange through the interaction of various fields, laser propagation, and thermonuclear fusion technology.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-142707 , JPL-TM-33-722
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  • 67
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    Internal voltage control of hydrogen-oxygen fuel cells: Feasibility study (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An experimental study was conducted to assess the feasibility of internal voltage regulation of fuel cell systems. Two methods were tested. In one, reactant partial pressure was used as the voltage control parameter and in the other reactant total pressure was used for control. Both techniques were breadboarded and tested on a single alkaline-electrolyte fuel cell. Both methods were found to be possible forms of regulation, however, of the two the total pressure technique would be more efficient, simpler to apply and would provide better transient characteristics.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TN-D-7956 , E-8155
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  • 68
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    Characteristics of response factors of coaxial gaseous rocket injectors (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The results of an experimental investigation undertaken to determine the frequency dependence of the response factors of various gaseous propellant rocket injectors subject to axial instabilities are presented. The injector response factors were determined, using the modified impedance-tube technique, under cold-flow conditions simulating those observed in unstable rocket motors. The tested injectors included a gaseous-fuel injector element, a gaseous-oxidizer injector element and a coaxial injector with both fuel and oxidizer elements. Emphasis was given to the determination of the dependence of the injector response factor upon the open-area ratio of the injector, the length of the injector orifice, and the pressure drop across the injector orifices. The measured data are shown to be in reasonable agreement with the corresponding injector response factor data predicted by the Feiler and Heidmann model.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-134788
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  • 69
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    Evaluation of refractory-metal-clad uranium nitride and uranium dioxide fuel pins after irradiation for times up to 10 450 hours at 990 C (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The effects of some materials variables on the irradiation performance of fuel pins for a lithium-cooled space power reactor design concept were examined. The variables studied were UN fuel density, fuel composition, and cladding alloy. All pins were irradiated at about 990 C in a thermal neutron environment to the design fuel burnup. An 85-percent dense UN fuel gave the best overall results in meeting the operational goals. The T-111 cladding on all specimens was embrittled, possibly by hydrogen in the case of the UN fuel and by uranium and oxygen in the case of the UO2 fuel. Tests with Cb-1Zr cladding indicate potential use of this cladding material. The UO2 fueled specimens met the operational goals of less than 1 percent cladding strain, but other factors make UO2 less attractive than low-density UN for the contemplated space power reactor use.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TN-D-7891 , E-7988
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  • 70
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    Study of liquid oxygen/liquid hydrogen auxiliary propulsion systems for the space tug (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Design concepts are considered that permit use of a liquid-liquid (as opposed to gas-gas) oxygen/hydrogen thrust chamber for attitude control and auxiliary propulsion thrusters on the space tug. The best of the auxiliary propulsion system concepts are defined and their principal characteristics, including cost as well as operational capabilities, are established. Design requirements for each of the major components of the systems, including thrusters, are developed at the conceptual level. The competitive concepts considered use both dedicated (separate tanks) and integrated (propellant from main propulsion tanks) propellant supply. The integrated concept is selected as best for the space tug after comparative evaluation against both cryogenic and storable propellant dedicated systems. A preliminary design of the selected system is established and recommendations for supporting research and technology to further the concept are presented.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-134790 , SD75-SA-0043
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  • 71
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    Power processing for electric propulsion (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: The potential of achieving up to 30 per cent more spacecraft payload or 50 per cent more useful operating life by the use of electric propulsion in place of conventional cold gas or hydrazine systems in science, communications, and earth applications spacecraft is a compelling reason to consider the inclusion of electric thruster systems in new spacecraft design. The propulsion requirements of such spacecraft dictate a wide range of thruster power levels and operational lifetimes, which must be matched by lightweight, efficient, and reliable thruster power processing systems. This paper will present electron bombardment ion thruster requirements; review the performance characteristics of present power processing systems; discuss design philosophies and alternatives in areas such as inverter type, arc protection, and control methods; and project future performance potentials for meeting goals in the areas of power processor weight (10 kg/kW), efficiency (approaching 92 per cent), reliability (0.96 for 15,000 hr), and thermal control capability (0.3 to 5 AU).
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: EASCON ''75; Electronics and Aerospace Systems Convention; Sep 29, 1975 - Oct 01, 1975; Washington, DC
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  • 72
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    NASA thermionic converter research and technology program (1975)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: This paper describes the NASA/ERDA research and technology program that was initiated in mid-FY 1974 with the objective of doubling the efficiency of thermionic power conversion with decreased emitter temperature. Also discussed are the potential uses of thermionic power conversion systems. Emphasis in this paper is placed on potential space applications, especially nuclear-electric propulsion (NEP). Possible development schedules are shown that would allow NEP systems to be ready for use in the 1990 time period for missions to the outer planets.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Energy 10; Annual Intersociety Energy Conversion and Engineering Conference; Aug 18, 1975 - Aug 22, 1975; Newark, DE
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  • 73
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    Charge-exchange plasma generated by an ion thruster (1975)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The use of high voltage solar arrays greatly reduces or eliminates power processing requirements in space electric propulsion systems. This use also requires substantial areas of solar array to be at high positive potential relative to space and most of the spacecraft. The charge exchange plasma conducts electrons from the ion beam to such positive surfaces, and thereby electrically load the high voltage solar array. To evaluate this problem, the charge-exchange plasma generated by an ion beam was investigated experimentally. Based upon the experimental data, a simple model was derived for the charge-exchange plasma. This model is conservative in the sense that both the electron/ion density and the electron current density should be equal to, or less than, the preducted value for all directions in the hemisphere upstream of the ion beam direction. Increasing the distance between a positive potential surface (such as a high voltage solar array) and the thruster is the simplest way to control interactions. Both densities and currents vary as the inverse square of this distance.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-134844
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  • 74
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    Determination of the extent of ion thruster efflux (1975)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: In the studies of proposed electric propulsion missions one of the areas of concern is the possible contamination of spacecraft instruments and thermal control surfaces by exhaust particles from an ion thruster. Vacuum tank tests were conducted in ground facilities to determine the extent of this deposition by thruster exhaust particles, but the application of these results to long term space missions is questionable. The flight thermal data from the SERT II satellite, the only electric propulsion mission with an extensive thruster operational history, was reviewed specifically to see if there is any evidence of contamination that could be attributed to the 5860 hours of mercury bombardment ion thruster operation. This evaluation of the flight data shows that the only evidence of deposition occurred on the contamination experiment solar cells which are located at the edge of the thruster exhaust beam. There is no evidence of any deposition of ion thruster efflux on any other surface of the satellite.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-71642 , E-8202 , 11th Elec. Propulsion Conf; Mar 19, 1975 - Mar 21, 1975; New Orleans, LA; United States
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  • 75
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    Measurement of sputtered efflux from 5-, 8-, and 30-cm diameter mercury ion thrusters (1975)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A study was undertaken to investigate the sputtered efflux from 5-, 8-, and 30-cm diameter mercury ion thrusters. Quartz crystal microbalances and fused silica samples were used to analyze the sputtered flux. Spectral transmittance measurements and spectrographic analysis of the samples were made after they were exposed to different thruster effluence by operating the thrusters at various conditions and durations of time. These measurements were used to locate the source of the efflux and determine its accumulated effect at various locations near the thruster. Comparisons of in situ and ex situ transmittance measurements of samples exposed to thruster efflux are also presented.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-71665 , E-8250 , 11th Elec. Propulsion Conf; Mar 19, 1975 - Mar 21, 1975; New Orleans, LA; United States
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  • 76
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    A 20000-hour endurance test of a structurally and thermally integrated 5-cm diameter ion thruster main cathode (1975)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A 5-cm diameter mercury ion thruster main cathode has completed over 20,000 hours of operation in an ongoing lifetime endurance test. The cathode operating parameters remained at acceptable performance levels throughout the test, the first 9175 hours of which were part of a thruster endurance test. After 20,000 hours, the cathode discharge was easily restarted, the tip orifice indicated negligible erosion and the tip heater showed no degradation. The cathode-isolator- vaporizer assembly, a major thruster subsystem, has thus successfully demonstrated an operational lifetime capability of 20,000 hours, which is the lifetime goal of the 8-cm diameter auxiliary propulsion ion thruster.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-71660 , 11th AIAA Elec. Propulsion Conf; Mar 19, 1975 - Mar 21, 1975; New Orleans, LA; United States
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  • 77
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    Studies of internal sputtering in a 30-cm ion thruster (1975)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Initial studies have been made of the sputtering and deposition phenomena in a 30-cm thruster. Sputtering rates, of the cathode baffle, one of the main sources of sputtered material in a thruster, have been measured by weight loss as a function of several thruster parameters. Sputtering rates were found to increase with both cathode flow rate and beam current when constant discharge voltage of 37 volts and power loses of 185 ev/ion were maintained. Sputtering rates were reduced 24% as discharge voltage was decreased from 37 to 33 volts while keeping discharge power constant. Qualitative agreement was found between sputtering rates obtained by the weight loss and those implied by spectroscopically observed line intensities of the excited iron sputtered atoms. After the completion of the sputtering tests, deposition and sputtering sites inside the thruster were identified.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-71654 , E-8232 , Elec. Propulsion Conf.,; Mar 19, 1975 - Mar 21, 1975; New Orleans, LA; United States
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  • 78
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    Control logic for a 30 cm diameter ion thruster (1975)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The various missions to be performed by the 30-cm diameter mercury bombardment thruster engine are discussed. The operating constraints imposed by the thermal environment, allowable time to reach steady state operation, and the number of start-ups required are examined. The variety of requirements is further analyzed for the impact on the basic control logic for the engine. The control logic is divided into the start-up, run, and shutdown modes of operation. The start-up mode is reported.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-71647 , E-8219 , Elec. Propulsion Conf.,; Mar 19, 1975 - Mar 21, 1975; New Orleans, LA; United States
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  • 79
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    The effects of exposure to LN2 temperatures and 2.5 suns solar radiation on 30-cm ion thruster performance (1975)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: An experimental test program was developed to demonstrate all 30 cm Hg-ion bombardment thruster functions over the thermal environment of several proposed missions. A 30 cm thruster with grids dished 1.25 cm and instrumented with 31 thermocouples, was placed in a vacuum tank equipped with minus 196 C walls. Cold storage of a thruster was simulated and temperatures as low as minus 100 C were attained on the thruster. The thruster started successfully from these cold conditions. The thruster operating at both half and full beam power was exposed to 2.5 suns on axis solar simulation. Various thruster thermal configurations, used to simulate multiple thruster operation, were tested at the above conditions. The results of these tests are reported herein.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-71652 , E-8227 , Elec. Propulsion Conf.,; Mar 19, 1975 - Mar 21, 1975; New Orleans, LA; United States
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  • 80
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    Direct thrust measurement of a 30-cm ion thruster (1975)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A direct thrust measurement of a 30-cm diameter ion thruster was accomplished by means of a laser interferometer thrust stand. The thruster was supported in a pendulum manner by three 3.65-m long wires. Electrical power was provided by means of 18 mercury filled pots. A movable 23-button planar probe rake was used to determine thrust loss due to ion beam divergence. Values of thrust, thrust loss due to ion beam divergence, and thrust loss due to multiple ionization were measured for ion beam currents ranging from 0.5 A to 2.5 A. Measured thrust values indicate an accuracy of approximately 1% and are in good agreement with thrust values calculated by indirect measurements.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-71646 , E-8215 , Elec. Propulsion Conf.,; Mar 19, 1975 - Mar 21, 1975; New Orleans, LA; United States
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  • 81
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    Auxiliary propulsion thruster performance with ion machined accelerator grids (1975)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A substantial improvement in the performance of an 8-centimeter-diameter auxiliary propulsion thruster was achieved by reducing the diameter of the accelerator grid apertures. The accelerator grid hole geometry was defined by ion machining accelerator grids on an 8-centimeter thruster at thrust levels of 2.2, 4.4, and 6.7 millinewtons was (mN). A thruster with an ion machined accelerator grid was operated at a thrust of 4.4 mN for 1000 hours. The discharge propellant utilization was 92% at an eV/ion of 338. Thruster performance and accelerator grid hole geometry was documented as a function of thrust level. It was also determined that the small hole accelerator grid has a very low backstreaming voltage limit. In fact the thruster can be operated with the acclerator grid held at neutralizer tip potential.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-71653 , Elec. Propulsion Conf.,; Mar 19, 1975 - Mar 21, 1975; New Orleans, LA; United States
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  • 82
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    SERT 2 thruster space restart, 1974 (1975)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The results of testing the flight thrusters on the SERT spacecraft during the 1974 test period are presented. The most notable result was the clearing of the high voltage short from thruster 2 and the successful stable operation of its ion beam. Test periods were limited to 70 minutes or less by earth eclipse of the spacecraft solar array and by ground station coverage limitations. Thruster 2 was restarted 26 times with an ion beam produced 21 times. The high voltage short remains in thruster 1, but the cathodes were restarted 12 times to demonstrate continued restart capability. The propellant feed systems, power processors, and spacecraft ancillary equipment were demonstrated to be functional after 4 1/2 years in space. In addition to the thruster tests, a neutralizer cathode was operated separately to demonstrate that the potential level of a spacecraft could be controlled by the neutralizer alone.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-71651 , E-8229 , 11th Elec. Propulsion Conf.; Mar 19, 1975 - Mar 21, 1975; New Orleans, LA; United States
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  • 83
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    Dynamic characteristic of a 30-cm mercury ion thruster (1975)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Measurements of the fluctuations of the discharge and beam plasmas of a 30 centimeter ion thruster were performed using 60 Hertz laboratory type power supplies. The time-varying properties of the discharge voltage and current, the ion beam current, and neutralizer keeper current were measured. The intensities of the fluctuations were found to depend on the beam and magnetic baffle currents. The shape of the frequency spectra of the discharge plasma fluctuations was found to be related to the beam and magnetic baffle currents. The measurements indicated that the discharge current fluctuations directly contribute to the beam current fluctuations and that the power supply characteristics modify these fluctuations.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-71655 , E-8234 , Elec. Propulsion Conf.; Mar 19, 1975 - Mar 21, 1975; New Orleans, LA; United States
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  • 84
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    A multiple thruster array for 30-cm thrusters (1975)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A 3.0 m diameter chamber of the 7.6 m diameter by 21.4 m long vacuum tank was modified to permit testing of an array of up to six 30-cm thrusters with a variety of laboratory and thermal vacuum breadboard power systems. A primary objective of the Multiple Thruster Array (MTA) program is to assess the impact of multiple thruster operation on individual thruster and power processor requirements. The areas of thruster startup, steady-state operation, throttling, high voltage recycle, thrust vectoring, and shutdown are of special concern. The results of initial tests are reported.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-71657 , E-8240 , Elec. Propulsion Conf.; Mar 19, 1975 - Mar 21, 1975; New Orleans; United States
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  • 85
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    Preliminary measurements of plasma fluctuations in an 8-cm mercury ion thruster (1975)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The rms magnitude, spectra, and cross correlations for the fluctuations in the beam current, the neutralizer keeper current, and the discharge current and voltage were measured for an 8-cm diameter, dished grid ion thruster for a beam current of 72 milliamps. The ratio of the rms magnitude of the fluctuations to the time-mean neutralizer keeper current was found to depend significantly on the neutralizer time-mean keeper current, the flow rate, and keeper hold diameter. The maxima of the spectra of the beam current fluctuations did not depend on the discharge fluctuations. It was found that: (1) the discharge current fluctuations do not directly contribute to the beam current fluctuations; and (2) the neutralizer contributions to the beam fluctuations are small (for good neutralizer-to-beam coupling) but not negligible and appear mostly in the higher frequency range measured.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-71650 , E-8226 , Elec. Propulsion Conf.,; Mar 19, 1975 - Mar 21, 1975; New Orleans, LA; United States
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  • 86
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    NiCd battery reliability for shuttle serviced spacecraft (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: Data from life-cycle testing of 22-cell NiCd batteries designed for shuttle-service spacecrafts are analyzed, first by averaging battery cycle lives as a function of temperature and depth of discharge, and then by calculating single cell failure and survival probabilities as a function of temperature, depth of discharge, and number of operating cycles. The single cell data are used to calculate battery survival probabilities. A mathematical failure model which is a function of the quantity of charge through the battery and the temperature is characterized, suggesting a predictive capability for accurate estimation of energy storage subsystem life and performance of accelerated tests. The failure mechanism described is apparently independent of most charge-discharge parameters and common to most cell sizes.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Energy 10; Annual Intersociety Energy Conversion and Engineering Conference; Aug 18, 1975 - Aug 22, 1975; Newark, DE
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  • 87
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    Design and test of a flywheel energy storage unit for spacecraft application (1975)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: This paper summarizes the design and test of a development flywheel energy storage device intended for spacecraft application. The flywheel unit is the prototype for the rotating assembly portion of an Integrated Power and Attitude Control System (IPACS). The paper includes a general description of the flywheel unit; specific design characteristics for the rotor and bearings, motor-generators, and electronics; an efficiency analysis; and test results for a research unit.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Energy 10; Annual Intersociety Energy Conversion and Engineering Conference; Aug 18, 1975 - Aug 22, 1975; Newark, DE
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  • 88
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    Coaxial injector response to axial instability under simulated cold-flow (1975)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: The frequency dependence of the admittances and response factors of various gaseous rocket injector configurations subject to axial instabilities under cold-flow conditions, have been measured using the modified impedance-tube technique. The tested configurations simulate the flows in a gaseous-fuel injector, gaseous-oxidizer injector and a coaxial injector with both fuel and oxidizer elements. Comparison of the measured response data with corresponding data predicted by the Feiler and Heidmann model indicates good agreement between the two sets of data.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 75-230 , American Institute of Aeronautics and Astronautics, Aerospace Sciences Meeting; Jan 20, 1975 - Jan 22, 1975; Pasadena, CA
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  • 89
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    One-millipound mercury ion thruster (1975)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: A mercury ion thruster has been developed for efficient operation at the nominal 1-mlb thrust level with a specific impulse of about 3,000 sec and a total power consumption of about 120 W. At a beam voltage of 1,200 V and beam current of 72 mA, the discharge chamber operates with a propellant efficiency of 93.8% at an ion-generation energy of 276 eV/ion. The 8-cm diameter thruster advances proven component technology to assure the capability for thruster operation over an accumulated beam-on time in excess of 20,000 hours with a capability for 10,000 on-off duty cycles. Discharge chamber optimization has combined stable current-voltage characteristics with high performance efficiency by careful placement of the discharge cathode near the location of a magnetic-field zero just upstream of the thruster endplate.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 75-386 , American Institute of Aeronautics and Astronautics, Electric Propulsion Conference; Mar 19, 1975 - Mar 21, 1975; New Orleans, LA
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  • 90
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    The cesium bombardment engine north-south stationkeeping experiment on ATS-6 (1975)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: Two one millipound-thrust cesium bombardment ion thrusters have been developed and integrated on the ATS-F spacecraft for the purpose of demonstrating compatible north-south stationkeeping of a synchronous communication satellite. Preliminary operation of the two thrusters on ATS-6 was completely successful on the first run of each. In addition to verifying operation, the principal accomplishments were the demonstration of a total absence of interference with the communications systems, verification of the predicted spacecraft operating potential, demonstration of compatibility with the star tracker, and demonstration of spacecraft attitude by thrust vectoring. Subsequent attempts to operate the thrusters have not been successful. Analysis indicates that the problem is associated with operation of the propellant reservoirs in zero-g.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 75-363 , American Institute of Aeronautics and Astronautics, Electric Propulsion Conference; Mar 19, 1975 - Mar 21, 1975; New Orleans, LA
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  • 91
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    Thermal analytical model of a 30 cm mercury ion thruster (1975)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: A lumped parameter thermal nodal network has been developed for a 30 cm engineering model mercury ion thruster. The network consists of approximately 100 nodes coded in SINDA format for use on the Univac 1106/1108 computer. This model takes into account internal dissipation, radiation, and conduction as well as environmental heating. A series of tests were performed at NASA-Lewis Research Center to simulate a wide range of thermal environments on an operating 30 cm thruster, instrumented to measure the temperature distribution within the thruster. The results of these tests were used to calibrate the analytical model. Presented in this paper are a description of the analytical model along with comparisons between analytical and experimental results for the various operating conditions.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 75-344 , American Institute of Aeronautics and Astronautics, Electric Propulsion Conference; Mar 19, 1975 - Mar 21, 1975; New Orleans, LA
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  • 92
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    Solar electric propulsion system thermal analysis (1975)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Thermal control elements applicable to the solar electric propulsion stage are discussed along with thermal control concepts. Boundary conditions are defined, and a thermal analysis was conducted with special emphasis on the power processor and equipment compartment thermal control system. Conclusions and recommendations are included.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-CR-120770 , SD-75-SA-0012
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  • 93
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    Summary of high efficiency silicon solar cell meeting held at NASA-Lewis (1975)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Attempts made to raise the efficiency of solar cells for space use are reported. The Helios, violet, and non-reflective cells were studied and it was concluded that the maximum practical efficiency of silicon solar cells is between 17 and 20%.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-71729 , E-8353 , d at the 11th Photovoltaic Specialists Conf.; May 06, 1975 - May 08, 1975; Phoenix, AZ; United States
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  • 94
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    Thermal analytic model of 30 cm engineering model mercury ion thruster (1975)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A lumped parameter thermal nodal network was developed for a 30 cm Engineering Model Mercury Ion Thruster. The network consists of approximately 100 nodes coded in SINDA format for use on the Univac 1106/1108 computer. This model takes into account internal dissipation, radiation, and conduction as well as environmental heating. A series of tests were performed to simulate a wide range of thermal environments on an operating 30 cm thruster, instrumented to measure the temperature distribution within the thruster. The results of these tests were used to calibrate the analytical model. The analytical model along with comparisons between analytical and experimental results for the various operating conditions are presented.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-71680 , Mar 19, 1975 - Mar 21, 1975; New Orleans, LA; United States
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  • 95
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    Solutions for discharge chamber sputtering and anode deposit spalling in small mercury ion thrusters (1975)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Proposed solutions to the problems of sputter erosion and sputtered material spalling in the discharge chamber of small mercury ion thrusters are presented. The accelerated life test evaluated three such proposed solutions: (1) the use of tantalum as a single low sputter yield material for the exposed surfaces of the discharge chamber components subject to sputtering, (2) the use of a severely roughened anode surface to improve the adhesion of the sputter-deposited coating, and (3) the use of a wire cloth anode surface in order to limit the size of any coating flakes which might spall from it. Because of the promising results obtained in the accelerated life test with anode surfaces roughened by grit-blasting, experiments were carried out to optimize the grit-blasting procedure. The experimental results and an optimal grit-blasting procedure are presented.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-71675 , AIAA PAPER 75-399 , 11th Elec. Propulsion Conf; Mar 19, 1975 - Mar 21, 1975; New Orleans, LA; United States
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  • 96
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    A structural and thermal packaging approach for power processing units for 30-cm ion thrusters (1975)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Solar Electric Propulsion (SEP) is currently being studied for possible use in a number of near earth and planetary missions. The thruster subsystem for these missions would consist of 30 centimeter ion thrusters with Power Processor Units (PPU) clustered in assemblies of from two to ten units. A preliminary design study of the electronic packaging of the PPU has been completed at Lewis Research Center of NASA. This study evaluates designs meeting the competing requirements of low system weight and overall mission flexibility. These requirements are evaluated regarding structural and thermal design, electrical efficiency, and integration of the electrical circuits into a functional PPU layout.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-71686 , E-8282 , Elect Porpulsion Conf; Mar 19, 1975 - Mar 21, 1975; New Orleans, LA; United States
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  • 97
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    A Thruster Sub-System Module (TSSM) for solar electric propulsion (1975)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Solar Electric Propulsion (SEP) is currently being studied for possible use in a number of near-earth and planetary missions. Thruster systems for these missions could be integrated directly into a spacecraft or modularized into a Thruster Sub-System Module (TSSM). A TSSM for electric propulsion missions would consist of a 30-cm ion thruster, thruster gimbal system, propellant storage and feed system, associated Power Processing Unit (PPU), thermal control system and complete supporting structure. The TSSM would be wholly self-contained and be essentially a plug-in or strap-on electric stage with simple mechanical, thermal, electrical and propellant interfaces. The TSSM described in this report is designed for a broad range of missions requiring from two to ten TSSM's mounted in a 2 by x configuration. The thermal control system is designed to accommodate waste heat from the power processor based on realistic efficiencies when the TSSM is operating from 0.7 to 3.5 AU's. The modules are 0.61 M (2 ft) wide by 2.29 M (7.5 ft) long and have a dry weight including propellant tank of 54.4 kg (120 lb). The propellant tank will hold 145.1 kg (320 lb) of mercury.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: NASA-TM-X-71683 , E-8276 , 11 Elec. Propulsion Conf; Mar 19, 1975 - Mar 21, 1975; New Orleans, LA; United States
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  • 98
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    Performance testing of thermoelectric generators at JPL (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: Several thermoelectric generators, ranging in output power from 170 watts to microwatts, are undergoing testing at JPL. They represent a wide range of technologies using advanced PbTe, SiGe and cascaded PbTe and BiTe thermoelectric materials. Several of these generators are of an advanced concept while others are representative of the Nimbus, Transit, Viking and the multi-hundred-watt (MHW) technology. Of interest is the behavior of generators which have been tested for times in excess of 60,000 hours.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: Energy 10; Annual Intersociety Energy Conversion and Engineering Conference; Aug 18, 1975 - Aug 22, 1975; Newark, DE
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  • 99
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    Thyristor power processor for the 30 cm mercury electric propulsion engine (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 75-433 , American Institute of Aeronautics and Astronautics, Electric Propulsion Conference; Mar 19, 1975 - Mar 21, 1975; New Orleans, LA
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  • 100
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    A structural and thermal packaging approach for power processing units for 30-cm ion thrusters (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: Solar Electric Propulsion (SEP) is currently being studied for possible use in a number of near earth and planetary missions. The thruster subsystem for these missions would consist of 30 centimeter ion thrusters with Power Processor Units (PPU) clustered in assemblies of from two to ten units. A preliminary design study of the electronic packaging of the PPU has been completed at Lewis Research Center of NASA. This study evaluates designs meeting the competing requirements of low system weight and overall mission flexibility. These requirements are evaluated regarding structural and thermal design, electrical efficiency, and integration of the electrical circuits into a functional PPU layout.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Type: AIAA PAPER 75-403 , American Institute of Aeronautics and Astronautics, Electric Propulsion Conference; Mar 19, 1975 - Mar 21, 1975; New Orleans, LA
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