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  • General Chemistry  (1,605)
  • PROPULSION SYSTEMS  (295)
  • Cell & Developmental Biology
  • 1970-1974  (1,415)
  • 1910-1914  (703)
  • 1973  (1,415)
  • 1910  (703)
Collection
Keywords
Publisher
Years
  • 1970-1974  (1,415)
  • 1910-1914  (703)
Year
  • 1
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    Investigation of the possibility of use of vibroacoustical signals for purposes of diagnostics in aeronautical engineering (1973)
    In:  CASI
    Details
    Publication Date: 2006-03-28
    Description: The application of vibroacoustic techniques for diagnosing aircraft engine malfunctions is discussed. An experiment was conducted to determine the defects introduced by the nature of change in the amplitude-frequency characteristics of the noises and vibrations of an aircraft jet engine. The manner in which the defects were simulated is explained. The test equipment used during the experiment is identified. The results of the amplitude-frequency characteristics investigation are summarized to show optimum location of the microphone pick-up to record the acoustic data.
    Keywords: PROPULSION SYSTEMS
    Type: Cybernetic Diagnostics of Mech. Systems with Vibro-acoustic Phenomena (NASA-TT-F-14899); p 317-319
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  • 2
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    The problem of carrying out a diagnosis of an internal combustion engine by vibroacoustical parameters (1973)
    In:  CASI
    Details
    Publication Date: 2006-03-28
    Description: The physics of noise formation in an internal combustion engine is discussed. A dependence of the acoustical radiation on the engine operating process, its construction, and operational parameters, as well as on the degree of wear on its parts, has been established. An example of tests conducted on an internal combustion engine is provided. A system for cybernetic diagnostics for internal combustion engines by vibroacoustical parameters is diagrammed.
    Keywords: PROPULSION SYSTEMS
    Type: Cybernetic Diagnostics of Mech. Systems with Vibro-acoustic Phenomena (NASA-TT-F-14899); p 247-249
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  • 3
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    Development of propulsion for high atmospheric pressure or dense environments (1973)
    In:  CASI
    Details
    Publication Date: 2006-02-22
    Description: The development of a propulsion system that employs a detonating propellant is described, and the need for such a system and its use in certain planetary atmospheres are demonstrated. A theoretical formulation of the relevant gas-dynamic processes was developed, and a related series of experimental tests were pursued.
    Keywords: PROPULSION SYSTEMS
    Type: JPL Quart. Tech. Rev., Vol. 3, No. 2 (NASA-CR-133863); p 45-52
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  • 4
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    Miscellaneous losses (1973)
    In:  CASI
    Details
    Publication Date: 2006-01-11
    Description: The boundary-layer losses associated with the flow process in the blade channel are discussed. To determine the overall design point efficiency of a turbine, other losses must also be considered; these include tip clearance loss and disk friction loss. The sum of these losses normally comprises all the losses that are considered in the design of a full admission axial flow turbine. If, however, a partial admission turbine is being considered, there are additional losses that must be included. The partial admission losses usually considered are the pumping loss in the inactive blade channels and the filling-and-emptying loss in the blade passages as they pass through the admission arc. Finally, a loss that occurs at off-design operation of any turbine is the incidence loss.
    Keywords: PROPULSION SYSTEMS
    Type: Turbine Design and Appl., Vol. 2; p 125-148
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  • 5
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    Channel flow analysis (1973)
    In:  CASI
    Details
    Publication Date: 2006-01-11
    Description: The design of a proper blade profile requires calculation of the blade row flow field in order to determine the velocities on the blade surfaces. An analysis theory is presented for several methods used for this calculation and associated computer programs that were developed are discussed.
    Keywords: PROPULSION SYSTEMS
    Type: Turbine Design and Appl., Vol. 2; p 27-56
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  • 6
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    Blade design (1973)
    In:  CASI
    Details
    Publication Date: 2006-01-11
    Description: The design of turbine blading is considered that will produce the flow angles and velocities required by velocity diagrams consistent with the desired efficiency and/or number of stages. The determination of the size, shape, and spacing of the blades is fundamental.
    Keywords: PROPULSION SYSTEMS
    Type: Turbine Design and Appl. Vol. 2; p 1-25
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  • 7
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    Operation of small rocket engines in the JPL high-vacuum molecular space simulator (Molsink) (1973)
    In:  CASI
    Details
    Publication Date: 2006-01-10
    Description: The feasibility of operating small rocket engines in the Molsink facility was demonstrated. A 0.44-N (0.1-lbf) hydrazine engine and a 0.18-N (0.04-lbf) thruster using cold gas from a hydrazine plenum system were operated for both flight duty cycles and off-nominal conditions. The exhaust gases from these thrusters contain NH3, N2, and H2. The chamber was also calibrated for larger bipropellant engines using nitrogen tetroxide/monomethyl hydrazine (NTO/MMH). The exhaust products of these engines contain CO2, CO, H2, H2O, and H2. A mixture of cold gases simulating the engine exhaust was injected through a nozzle under conditions simulating thrust levels up to 26.7-N (6 lbf). Pulsing and continuous operations were investigated. The chamber background pressure traces were compared with the traces obtained for the same thrusters operated with pure nitrogen at approximately equivalent thrust. Satisfactory recuperation times were encountered in all the pulsing modes. Test times greater than 20s were obtained in steady state operation before the vacuum chamber back pressure climbed to prohibitive values.
    Keywords: PROPULSION SYSTEMS
    Type: JPL Quart. Tech. Rev., Vol. 3, No. 1; p 1-13
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  • 8
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    Chemical propulsion - The old and the new challenges (1973)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: The historical background concerning the application of liquid propellant rockets is considered. Progress to date in chemical liquid propellant rocket engines can be summarized as an increase in performance through the use of more energetic propellant combinations and increased combustion pressure. New advances regarding liquid propellant rocket engines are related to the requirement for reusability in connection with the development of the Space Shuttle.
    Keywords: PROPULSION SYSTEMS
    Type: AIAA Student Journal; 11; Dec. 197
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  • 9
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    Technology for low noise aircraft - The NASA Quiet Engines (1973)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: Review of the major goals, procedures, and results of the Quiet Engine Program that was initiated four years ago and is now nearing completion. This program has developed and demonstrated, in full-scale, experimental engine tests, technology advances which, if applied to the design of future aircraft, will help produce equipment with noise levels considerably lower than the older narrow-body aircraft and significantly lower than the new wide-body aircraft flying at present. However, the application of this noise reduction technology will result in increases in aircraft operating costs. Future aircraft noise reduction research should, therefore, consider improvements in the economics associated with noise reduction technology.
    Keywords: PROPULSION SYSTEMS
    Type: Noise Control Engineering; 1; Autumn 1
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  • 10
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    New horizons in chemical propulsion (1973)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: After a review of the work of the late-Fifties on free radicals for propulsion, it is concluded that atomic hydrogen would provide a potentially large increase in specific impulse. Work conducted to find an approach for isolating atomic hydrogen is considered. Other possibilities for obtaining propellants of greatly increased capability might be connected with the technology for the generation of activated states of gases, metallic hydrogen, fuels obtained from other planets, and laser transfer of energy.
    Keywords: PROPULSION SYSTEMS
    Type: Astronautics and Aeronautics; 11; Dec. 197
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  • 11
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    Effect of shear on duct wall impedance. (1973)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: The solution to the equation governing the propagation of sound in a uniform shear layer is expressed in terms of parabolic cylinder functions. This result is used to develop a closed-form solution for acoustic wall impedance which accounts for both the duct liner and the presence of a boundary layer in the duct. The effective wall impedance can then be used as the boundary condition for the much simpler problem of sound propagation in uniform flow.
    Keywords: PROPULSION SYSTEMS
    Type: Journal of Sound and Vibration; 30; Sept. 8
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  • 12
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    Performance of a modified downstream-cathode MPD thruster. (1973)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: Discussion of the improvements incorporated since 1969 in a low-power MPD thruster design which is considered a potential candidate for satellite station keeping and attitude control. The improvements include a new cathode design, and changes in thruster geometry, with xenon used as the propellant. A better thermal design is found to be necessary for further improvement of the thruster.
    Keywords: PROPULSION SYSTEMS
    Type: Journal of Spacecraft and Rockets; 10; Jan. 197
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  • 13
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    A Quiet Engine for stilling complaints. (1973)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: Two experimental Quiet Engines using derated CF6 cores are discussed. One engine has a low-speed fan running at a tip speed of 1160 fps; the other engine has a fan running at the high speed of 1550 fps. The two engines are expected to show the relative advantages of fans operating at low tip speeds with high lift coefficients in comparison with fans operating at high tip speeds with low lift coefficients. Test results obtained with full-scale (6-ft diameter) fans are examined.
    Keywords: PROPULSION SYSTEMS
    Type: Astronautics and Aeronautics; 11; Jan. 197
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  • 14
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    Design and evaluation of combustors for reducing aircraft engine pollution (1973)
    In:  CASI
    Details
    Publication Date: 2012-05-22
    Description: Various techniques and test results are briefly described and referenced for detail. The effort arises from the increasing concern for the measurement and control of emissions from gas turbine engines. The greater part of this research is focused on reducing the oxides of nitrogen formed during takeoff and cruise in both advanced CTOL, high pressure ratio engines, and advanced supersonic aircraft engines. The experimental approaches taken to reduce oxides of nitrogen emissions include the use of: multizone combustors incorporating reduced dwell time, fuel-air premixing, air atomization, fuel prevaporization, water injection, and gaseous fuels. In the experiments conducted to date, some of these techniques were more successful than others in reducing oxides of nitrogen emissions. Tests are being conducted on full-annular combustors at pressures up to 6 atmospheres and on combustor segments at pressures up to 30 atmospheres.
    Keywords: PROPULSION SYSTEMS
    Type: AGARD Atmospheric Pollution by Aircraft Engine; 8 p
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  • 15
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    Engine selection for transport and combat aircraft (1973)
    In:  CASI
    Details
    Publication Date: 2012-05-23
    Description: The procedures for selecting engines for transport and combat aircraft during the design process are presented. The types of aircraft considered are: (1) a long haul conventional takeoff and landing transport, (2) a short haul vertical takeoff and landing transport, (3) a long range supersonic transport, and (4) a fighter aircraft. The influence of aircraft noise considerations on engine selection is examined. The aerodynamic characteristics of supercritical wings and their effect on engine selection are reported.
    Keywords: PROPULSION SYSTEMS
    Type: AGARD Aircraft Performance: Prediction Methods and Optimization; 55 p
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  • 16
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    Velocity control propulsion subsystem of the Radio Astronomy Explorer satellite for Goddard Space Flight Center (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The Velocity Control Propulsion Subsystem (VCPS) was designed the propulsion required for trajectory and lunar orbit corrections of the spacecraft. A GFE clamp assembly physically attaches the VCPS to the spacecraft and the unit is ejected after completing the required corrections. The VCPS is physically and functionally separated from the spacecraft except for the electrical and telemetry interfaces. A GFE transtage provides the superstructure on which the VCPS is assembled. The subsystem consists of two 5 foot pound rocket engine assemblies, 4 propellant tanks, 2 latching valves, 2 fill and drain valves, a system filter, pressure transducer, gas and propellant manifolds and electrical heaters and thermostats. The RAE-B VCPS program covered the design, manufacture and qualification of one subsystem. This subsystem was to be manufactured, subjected to qualification tests; and refurbished, if necessary, prior to flight. The VCPS design and test program precluded the need for refurbishing the subsystem and the unit was delivered to GSFC at the conclusion of the program.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-132905 , SVHSER-6226
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  • 17
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    Advanced space engine preliminary design (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A preliminary design was completed for an O2/H2, 89 kN (20,000 lb) thrust staged combustion rocket engine that has a single-bell nozzle with an overall expansion ratio of 400:1. The engine has a best estimate vacuum specific impulse of 4623.8 N-s/kg (471.5 sec) at full thrust and mixture ratio = 6.0. The engine employs gear-driven, low pressure pumps to provide low NPSH capability while individual turbine-driven, high-speed main pumps provide the system pressures required for high-chamber pressure operation. The engine design dry weight for the fixed-nozzle configuration is 206.9 kg (456.3 lb). Engine overall length is 234 cm (92.1 in.). The extendible nozzle version has a stowed length of 141.5 cm (55.7 in.). Critical technology items in the development of the engine were defined. Development program plans and their costs for development, production, operation, and flight support of the ASE were established for minimum cost and minimum time programs.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121237 , PWA-FR-5654
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  • 18
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    Solid motor diagnostic instrumentation (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A review of typical surveillance and monitoring practices followed during the flight phases of representative solid-propellant upper stages and apogee motors was conducted to evaluate the need for improved flight diagnostic instrumentation on future spacecraft. The capabilities of the flight instrumentation package were limited to the detection of whether or not the solid motor was the cause of failure and to the identification of probable primary failure modes. Conceptual designs of self-contained flight instrumentation packages capable of meeting these reqirements were generated and their performance, typical cost, and unit characteristics determined. Comparisons of a continuous real time and a thresholded hybrid design were made on the basis of performance, mass, power, cost, and expected life. The results of this analysis substantiated the feasibility of a self-contained independent flight instrumentation module as well as the existence of performance margins by which to exploit growth option applications.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-136561 , JPL-TM-33-656
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  • 19
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    Single-stage experimental evaluation of tandem-airfoil rotor stator blading for compressors. Part 6: Data and performance for stage D (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An axial flow compressor stage, having single-airfoil blading, was designed for zero rotor prewhirl, constant rotor work across the span, and axial discharge flow. The stage was designed to produce a pressure ratio of 1.265 at a rotor tip velocity of 757 ft/sec. The rotor had an inlet hub/tip ratio of 0.8. The design procedure accounted for the rotor inlet boundary layer and included the effects of axial velocity ratio and secondary flow on blade row performance. The objectives of this experimental program were: (1) to obtain performance with uniform and distorted inlet flow for comparison with the performance of a stage consisting of tandem-airfoil blading designed for the same vector diagrams; and (2) to evaluate the effectiveness of accounting for the inlet boundary layer, axial velocity ratio, and secondary flows in the stage design. With uniform inlet flow, the rotor achieved a maximum adiabatic efficiency of 90.1% at design equivalent rotor speed and a pressure ratio of 1.281. The stage maximum adiabatic efficiency at design equivalent rotor speed with uniform inlet flow was 86.1% at a pressure ratio of 1.266. Hub radial, tip radial, and circumferential distortion of the inlet flow caused reductions in surge pressure ratio of approximately 2, 10 and 5%, respectively, at design rotor speed.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134511 , PWA-FR-5852-PT-6
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  • 20
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    Exploratory development of a glass ceramic automobile thermal reactor (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: This report summarizes the design, fabrication and test results obtained for glass-ceramic (CER-VIT) automotive thermal reactors. Several reactor designs were evaluated using both engine-dynamometer and vehicle road tests. A maximum reactor life of about 330 hours was achieved in engine-dynamometer tests with peak gas temperatures of about 1065 C (1950 F). Reactor failures were mechanically induced. No evidence of chemical degradation was observed. It was concluded that to be useful for longer times, the CER-VIT parts would require a mounting system that was an improvement over those tested in this program. A reactor employing such a system was designed and fabricated.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134513
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  • 21
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    Emission calculations for a scramjet powered hypersonic transport (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Calculations of exhaust emissions from a scramjet powered hypersonic transport burning hydrogen fuel were performed over a range of Mach numbers of 5 to 12 to provide input data for wake mixing calculations and forecasts of future levels of pollutants in the stratosphere. The calculations were performed utilizing a one-dimensional chemical kinetics computer program for the combustor and exhaust nozzle of a fixed geometry dual-mode scramjet engine. Inlet conditions to the combustor and engine size was based on a vehicle of 227,000 kg (500,000 lb) gross take of weight with engines sized for Mach 8 cruise. Nitric oxide emissions were very high for stoichiometric engine operation but for Mach 6 cruise at reduced equivalence ratio are in the range predicted for an advanced supersonic transport. Combustor designs which utilize fuel staging and rapid expansion to minimize residence time at high combustion temperatures were found to be effective in preventing nitric oxide formation from reaching equilibrium concentrations.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-71464 , E-7760
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  • 22
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    Test report for 120-inch-diameter Solid Rocket Booster (SRB) model tests (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The space shuttle solid rocket boosters (SRB's) will be jettisoned to impact in the ocean within a 200-mile radius of the launch site. Tests were conducted at Long Beach, California, using a 12-inch diameter Titan 3C model to simulate the full-scale characteristics of the prototype SRB during retrieval operations. The objectives of the towing tests were to investigate and assess the following: (1) a floating and towing characteristics of the SRB; (2) need for plugging the SRB nozzle prior to tow; (3) attach point locations on the SRB; (4) effects of varying the SRB configuration; (5) towing hardware; and (6) difficulty of attaching a tow line to the SRB in the open sea. The model was towed in various sea states using four different types and varying lengths of tow line at various speeds. Three attach point locations were tested. Test data was recorded on magnetic tape for the tow line loads and for model pitch, roll, and yaw characteristics and was reduced by computer to tabular printouts and X-Y plots. Profile and movie photography provided documentary test data.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-69441 , KSC-TR-1253
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  • 23
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    Radially leaned outlet guide vanes for fan source noise reduction (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Two quiet engine program half scale fans one with a subsonic and the other with a supersonic fan tip speed at takeoff were run with 30 degree leaned and radial outlet guide vanes. Acoustic data at takeoff fan speed on the subsonic tip speed fan showed decreases in 200-foot sideline noise of from 1 to 2 PNdb. The supersonic tip speed fan a takeoff fan speed, however, showed noise increases of up 3 PNdb and a decrease in fan efficiency. At approach fan speed, the subsonic tip speed fan showed a noise decrease of 2.3 PNdb at the 200-foot sideline maximum angle and an increase in efficiency. The supersonic tip speed fan showed noise increase of 3.5 PNdb and no change in efficiency. The decrease in fan efficiency and the nature of the noise increase largely high frequency broadband noise lead to the speculation that an aerodynamic problem occurred.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134486 , R73AEG176
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  • 24
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    Testing and evaluation of the LES-6 pulsed plasma thruster by means of a torsion pendulum system (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Performance characteristics of the LES-6 pulsed plasma thruster over a range of input conditions were investigated by means of a torsion pendulum system. Parameters of particular interest included the impulse bit and time average thrust (and their repeatability), specific impulse, mass ablated per discharge, specific thrust, energy per unit area, efficiency, and variation of performance with ignition command rate. Intermittency of the thruster as affected by input energy and igniter resistance were also investigated. Comparative experimental data correlation with the data presented. The results of these tests indicate that the LES-6 thruster, with some identifiable design improvements, represents an attractive reaction control thruster for attitude contol applications on long-life spacecraft requiring small metered impulse bits for precise pointing control of science instruments.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-135940 , JPL-TM-33-630
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  • 25
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    Performance of a single-stage transonic compressor with a blade-tip solidity of 1.5 and comparison with 1.3 and 1.7 solidity stages (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The overall and blade-element performance of a transonic compressor stage with a tip solidity of 1.5 is presented over the stable operating range at rotative speeds from 50 to 100 percent of design speed. State peak efficiency of 0.82 was obtained at a weight flow of 29.4 kg.sec (200.4 (kg/sec)/m2 of annulus area) and a pressure ratio of 1.71. Stall margin at design speed was 14 percent. A comparison of three stages in a solidity study showed that the performance of the 1.5 solidity stage and the 1.3 solidity stage were nearly identical but that the performance of the 1.7 solidity stage was significantly lower.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2926 , E-7255
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  • 26
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    DC-9/JT8D refan, Phase 1 (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Analyses and design studies were conducted on the technical and economic feasibility of installing the JT8D-109 refan engine on the DC-9 aircraft. Design criteria included minimum change to the airframe to achieve desired acoustic levels. Several acoustic configurations were studied with two selected for detailed investigations. The minimum selected acoustic treatment configuration results in an estimated aircraft weight increase of 608 kg (1,342 lb) and the maximum selected acoustic treatment configuration results in an estimated aircraft weight increase of 809 kg (1,784 lb). The range loss for the minimum and maximum selected acoustic treatment configurations based on long range cruise at 10 668 m (35,000 ft) altitude with a typical payload of 6 804 kg (15,000 lb) amounts to 54 km (86 n. mi.) respectively. Estimated reduction in EPNL's for minimum selected treatment show 8 EPNdB at approach, 12 EPNdB for takeoff with power cutback, 15 EPNdB for takeoff without power cutback and 12 EPNdB for sideline using FAR Part 36. Little difference was estimated in EPNL between minimum and maximum treatments due to reduced performance of maximum treatment. No major technical problems were encountered in the study. The refan concept for the DC-9 appears technically feasible and economically viable at approximately $1,000,000 per airplane. An additional study of the installation of JT3D-9 refan engine on the DC-8-50/61 and DC-8-62/63 aircraft is included. Three levels of acoustic treatment were suggested for DC-8-50/61 and two levels for DC-8-62/63. Results indicate the DC-8 technically can be retrofitted with refan engines for approximately $2,500,000 per airplane.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121252 , MDC-J5738
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  • 27
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    Study of small turbofan engines applicable to general-aviation aircraft (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The applicability of small turbofan engines to general aviation aircraft is discussed. The engine and engine/airplane performance, weight, size, and cost interrelationships are examined. The effects of specific engine noise constraints are evaluated. The factors inhibiting the use of turbofan engines in general aviation aircraft are identified.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-114630 , AIRESEARCH-73-210148
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  • 28
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    Correlation of ion and beam current densities in Kaufman thrusters. (1973)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: In the absence of direct impingement erosion, electrostatic thruster accelerator grid lifetime is defined by the charge exchange erosion that occurs at peak values of the ion beam current density. In order to maximize the thrust from an engine with a specified grid lifetime, the ion beam current density profile should therefore be as flat as possible. Knauer (1970) has suggested this can be achieved by establishing a radial plasma uniformity within the thruster discharge chamber; his tests with the radial field thruster provide an example of uniform plasma properties within the chamber and a flat ion beam profile occurring together. It is shown that, in particular, the ion density profile within the chamber determines the beam current density profile, and that a uniform ion density profile at the screen grid end of the discharge chamber should lead to a flat beam current density profile.
    Keywords: PROPULSION SYSTEMS
    Type: Journal of Spacecraft and Rockets; 10; Sept
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  • 29
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    Earth storable bimodal engine, phase 1 (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An in-depth study of an Earth Storable Bimodal (ESB) Engine using earth storable propellants N2O/N2H4 and operating in either a monopropellant or bipropellant mode was conducted. Detailed studies were completed for both a hot-gas, regeneratively cooled thrust chamber and a ducted hot-gas, film cooled thrust chamber. Hydrazine decomposition products were used for cooling in either configuration. The various arrangements and configurations of hydrazine reactors, secondary injectors, chambers and gimbal methods were considered. The two basic materials selected for the major components were columbium alloys and L-605. The secondary injector types considered were previously demonstrated by JPL and consisted of a liquid-on-gas triplet, a liquid-on-gas doublet, and a liquid-on-gas coaxial injector. Various design tradeoffs were made with different reactor types located at: the secondary injector station, the thrust chamber throat, and the nozzle/extension interface. Associated thermal, structural, and mass analyses were completed.
    Keywords: PROPULSION SYSTEMS
    Type: BELL-8706-933005 , NASA-CR-133617
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  • 30
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    Preliminary study of turbojets with rotary flow inductors for a low-noise supersonic transport (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: In a simplified airplane-mission study for a Mach 2.61 supersonic transport, dry turbojets with and without real suppressors and dry turbojets with ideal rotary flow inductors were studied for sideline noise levels as low as FAR 36-20. Compressor pressure ratio was varied from 5 to 30 and turbine temperature from 1800 to 3000 F. For no noise constraint and without a suppressor, the best dry turbojet gave a payload of 9.0 percent of gross weight and a sideline noise of 126 effective perceived noise decibels. Payload dropped rapidly for lower noise goals, becoming 6.3 percent of gross weight at FAR 36. At FAR 36, the turbojet with suppressor gave a payload of 8.3 percent and the turbojet with ideal rotary flow inductor, 7.3 percent. Below FAR 36, the ideal inductor was far superior to the real suppressor, giving payloads of 6.6 percent at FAR 36-10 and 5.7 percent at FAR 36-20.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-68233 , E-7450
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  • 31
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    Steady-state and dynamic pressure phenomena in the propulsion system of an F-111A airplane (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Flight tests were conducted with two F-111A airplanes to study the effects of steady-state and dynamic pressure phenomena on the propulsion system. Analysis of over 100 engine compressor stalls revealed that the stalls were caused by high levels of instantaneous distortion. In 73 percent of these stalls, the instantaneous circumferential distortion parameter, k sub theta, exhibited a peak just prior to stall higher than any previous peak. The K sub theta parameter was a better indicator of stall than the distortion factor, k sub d, and the maximum-minus-minimum distortion parameter, d, was poor indicator of stall. Inlet duct resonance occurred in both F-111A airplanes and is believed to have been caused by oscillations of the normal shock wave from an internal to an external position. The inlet performance of the two airplanes was similar in terms of pressure recovery, distortion, and turbulence, and there was good agreement between flight and wind-tunnel data up to a Mach number of approximately 1.8.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-7328 , H-741
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  • 32
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    Performance optimized, small structurally integrated ion thruster system (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A 5-cm structurally integrated ion thruster has been developed for attitude control and stationkeeping of synchronous satellites. As optimized with a conventional ion extraction system, the system demonstrates a thrust T = 0.47 mlb at a beam voltage of 1600 V, total mass efficiency of 76%, and electrical efficiency of 56%. Under the subject contract effort, no significant performance change was noted for operation with two dimensional electrostatic thrust-vectoring grids. Structural integrity with the vectoring grids was demonstrated for shock (+ or - 30 G), sinusoidal (9 G), and random (19.9 G rms) accelerations. System envelope is 31.2 cm long by 13.4 cm flange bolt circle, with a mass of 9.0 Kg, including 6.8 Kg mercury propellant.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121183
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  • 33
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    Flyover and static tests to study flight velocity effects on jet noise of suppressed and unsuppressed plug nozzle configurations (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Two spoke-type suppressor plug nozzles and a basic plug nozzle were tested for noise and thrust performance. The nozzles were mounted on an underwing nacelle on an F-106B aircraft, and tests were made both statically and in flyovers at Mach 0.4 at an altitude of 91 meters (300 ft). The flight and static data were adjusted to common reference conditions so that direct comparisons could be made. The noise characteristics that these nozzles would have on a large multiengine aircraft at a 640-meter (2100-ft) sideline distance are also presented. Flight noise levels for all three nozzles were higher than static at comparable conditions; and a shift in the frequency spectra was seen from static to flight, indicating the presence of a forward velocity effect on the noise characteristics.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2856 , E-7186
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  • 34
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    Investigation of noise suppression by sonic inlets for turbofan engines. Volume 1: Program summary (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Results of a program for sonic inlet technology development are presented. This program includes configuration and mechanical design selection of concepts, aerodynamic design description of the models, and results of test evaluation. Several sonic inlet concepts were tested and compared for aerodynamic and acoustic performance. Results of these comparative evaluations are presented. Near-field measurements were taken inside several of the inlet models. Results of these tests are discussed with respect to the effect of Mach number gradients on noise attenuation and rotor shock wave attenuation, and boundary layer effects on noise propagation. The test facilities and experimental techniques employed are described briefly.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121126 , D6-40855-VOL-1
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  • 35
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    Investigation of noise suppression by sonic inlets for turbofan engines. Volume 3: An experimental investigation of the internal noise field of two axisymmetric sonic inlet models (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: For abstract, see
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121128 , D6-40818-VOL-3
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  • 36
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    Investigation of noise suppression by sonic inlets for turbofan engines. Volume 2: Appendixes (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: For abstract, see
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121127 , D6-40855-1-VOL-2
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  • 37
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    Natural environment design requirements for the Solar Electric Propulsion Stage (SEPS) (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The natural environment design requirements for the solar electric propulsion stage are presented. Environment criteria for the SEP stage will cover earth orbital operations out to geosynchronous altitudes and also interplanetary missions including comet and asteroid missions.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-64761
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  • 38
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    Dynamic response of Mach 2.5 axisymmetric inlet with 40 percent supersonic internal area contraction (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Results of experimental tests conducted on a supersonic, mixed compression, axisymmetric inlet are presented. The inlet is designed for operation at Mach 2.5 with a turbofan engine (TF-30). The inlet was terminated with either a choked-orifice plate or a long pipe with variable area choked exit plug. Frequency responses were obtained for selected static pressures in the diffuser. These pressures were selected as potential control signals for terminal shock control. Frequency responses were obtained for the Mach 2 and 2.5 conditions for different terminations. Responses also were obtained with and without cowl bleed. Internal disturbances were produced by sinusoidally varying the inlet overboard bypass doors at frequencies out to 100 hertz.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2833 , E-7426
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  • 39
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    Experimental quiet engine program. Predicted engine performance (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Three turbofan configurations, each incorporating alternative noise reduction features, were tested under the Quiet Engine Program. Performance data for these engines are shown over a range of flight conditions. The data are presented in tabular form for standard day flight inlet conditions. Procedures for estimating nonstandard day performance are shown. Tabular data and calculation procedures to allow determination of ram recovery, customer bleed, and customer shaft power extraction effects on engine performance can be found in the original Performance Brochure titled, Experimental Quiet Engine Program, Predicted Engine Performance, dated April 8, 1970. Predicted engine noise levels for representative take-off and approach conditions are provided.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121258
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  • 40
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    Development of a biowaste resistojet propulsion system (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The equipment, exclusive of thrustors, required to demonstrate the feasibility of a resistojet propulsion system for space station attitude control application using representative simulated crew biowaste propellants and available resistojet thrustors in the ground simulation tests is discussed. The overall objective of the program was to provide a biowaste resistojet prototype propellant management and control system sufficiently similar to the flight article to permit concept feasibility and system demonstration testing of interface compatibility, operational characteristics, and system flexibility.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-132269 , MDC-G4745
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  • 41
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    Study of unconventional propulsion system concepts for use in a long range transport (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The noise level of an uncoventional propulsion system for the next generation of subsonic, long-range transport aircraft is discussed. The desired noise level may be achieved by: (1) a fixed geometry, high bypass ratio turbofan with a geared two-stage fan and advanced acoustic treatment or (2) a moderate bypass ratio turbofan with a variable pitch two-stage fan, variable primary and duct nozzles, and advanced acoustic treatment. The geared fan system meets the noise goal with minimum economic penalty. Comparison of the noise levels at takeoff and landing in combination with the economic penalties required to achieve the lower noise levels at specific noise measuring stations, indicate that both area reduction and current certification prodedures should be used to ascertain the point of diminishing returns in establishing future noise goals.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121242 , PWA-4692
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  • 42
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    Single grid accelerator for an ion thrustor (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A single grid accelerator system for an ion thrustor is discussed. A layer of dielectric material is interposed between this metal grid and the chamber containing an ionized propellant for protecting the grid against sputtering erosion.
    Keywords: PROPULSION SYSTEMS
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  • 43
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    Monopropellant hydrazine resistoject: Engineering model fabrication and test task (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The monopropellant hydrazine resistojet, termed the electrothermal hydrazine thruster (EHT) by TRW systems, thermally decomposes anhydrous hydrazine propellant to produce a high-temperature, low-molecular-weight gas for expulsion through a propulsive nozzle. The EHT developed for this program required about 3-5 watts of electrical power and produced 0.020 to 0.070 pound of thrust over the inlet pressure range of 100 to 400 psia. The thruster was designed for both pulsed and steady state operation. A summary of the GSFC original requirements and GSFC modified requirements, and the performance of the engineering model EHT is given. The experimental program leading to the engineering model EHT design, modifications necessary to achieve the required thruster life capability, and the results of the life test prgram. Other facets of the program, including analyses, preliminary design, specifications, data correlation, and recommendations for a flight model are discussed.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-132784 , TRW-20266-6024-RO-00
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  • 44
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    Performance of a 1.20 pressure ratio STOL fan stage at three rotor blade setting angles (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A model of a short takeoff and landing (STOL) fan stage was testedin a single-stage compressor research facility. Surveys of the airflow conditions ahead of the rotor, between the rotor and stator, and behind the stator were made over the stable operating range of the stage. At the design speed of 213.3 meters per second and a weight flow of 31.2 kilograms per second, the stage pressure ratio of 1.15 was less than the design value of 1.2. The stage was tested with the rotor blades reset for more flow. Design pressure ratio was achieved and surpassed with the minus 5 deg and minus 7 deg resets, respectively. The stage efficiency was 0.88 for the minus 5 deg reset and 0.85 for the minus 7 deg reset.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2837 , E-7434
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  • 45
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    Monopropellant hydrazine resistojet: Data correlation (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The design, development, and testing of an engineering model nominal 20-millipound thrust monopropellant hydrazine resistojet program is divided into six basic tasks. Included in these tasks are analyses, design, test, and data correlation of the electrothermal hydrazine thruster (EHT). A brief summary is provided of the analyses conducted for the EHT and the design of the engineering model thruster. Some of the results of the engineering model tests are then compared with the analytical performance models generated early in the program.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-130762 , TRW-20266-6025-RO-00
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  • 46
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    Experimental cold-flow evaluation of a ram air cooled plug nozzle concept for afterburning turbojet engines (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A concept for plug nozzles cooled by inlet ram air is presented. Experimental data obtained with a small scale model, 21.59-cm (8.5-in.) diameter, in a static altitude facility demonstrated high thrust performance and excellent pumping characteristics. Tests were made at nozzle pressure ratios simulating supersonic cruise and takeoff conditions. Effect of plug size, outer shroud length, and varying amounts of secondary flow were investigated.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2811 , E-7387
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  • 47
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    Single stage experimental evaluation of tandem-airfoil rotor and stator blading for compressors. Part 4: Data and performance for stage B (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Stage B, composed of tandem-airfoil rotor B and stator B, was tested with uniform inlet flow and with hub radial, tip radial and 90 degree one-per-revolution circumferential distortion of the inlet flow as part of an overall program to evaluate the effectiveness of tandem airfoils for increasing the design point loading capability and stable operating range of rotor and stator blading. The results of this series of tests provide overall performance and blade element data for evaluating: (1) the potential of tandem blading for extending the loading limit and stable operating range of a stage representative of a middle stage of an advanced high pressure compressor, (2) the effect of loading split between the two airfoils in tandem on the performance of tandem blading, and (3) the effects of inlet flow distortion on the stage performance. The rotor had an inlet hub/tip ratio of 0.8 and a design tip velocity of 757 ft/sec. With uniform inlet flow, rotor B achieved a maximum adiabatic efficiency of 88.4% at design equivalent rotor speed and a pressure ratio of 1.31. The stage maximum adiabatic efficiency at design equivalent rotor speed with uniform inlet flow was 82.5% at a pressure ratio of 1.28. Tip radial and circumferential distortion of the inlet flow caused substantial reductions in surge margin.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121145 , FR-5083-PT-4
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  • 48
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    Analytic and experimental performance of two isentropic mixed compression axisymmetric inlets at Mach numbers 0.8 to 2.65 (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A mixed compression axisymmetric inlet model with a capture diameter of 50 cm was tested at Mach numbers ranging from 0.8 to 2.65 at 0 deg angle of attack and a constant total pressure of approximately 1 atm. Analytical methods accounting for the effects of both viscous and inviscid flows and incorporating empirical bleed discharge coefficients were used in the procedure for designing the inlet contours and the bleed system. Experimental results are compared with analytic predictions and are also compared with results from earlier tests of an inlet with the same internal contours but with a bleed system developed by cut and try methods in the wind tunnel. With the bleed configuration predicted by the design procedure, maximum total pressure recovery at the engine face at the design Mach number of 2.65 was 93 percent, with a total pressure distortion less than 10 percent. Corresponding bleed mass flow was approximately 7.5 percent, which was about 1.3 percent less than predicted. At lower supersonic Mach numbers, pressure recovery and bleed were generally lower and distortion generally higher.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-7320 , A-4675
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  • 49
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    A 13000-hour test of a mercury hollow cathode (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A mercury-fed hollow cathode was tested for 12,979 hours in a bell jar at SERT 2 neutralizer operating conditions. The net electron current drawn to a collector was 0.25 ampere at average collector voltages between 21.8 and 36.7 volts. The mercury flow rate was varied from 5.6 to 30.8 equivalent milliamperes to give stable operation at the desired electrode voltages and currents. Variations with time in the neutralizer discharge characteristics were observed and hypothesized to be related to changes in the cathode orifice dimensions and the availability of electron emissive material. A facility failure caused abnormal test conditions for the last 876 hours and led to the cathode heater failure which concluded the test.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2785 , E-7163
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  • 50
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    A simple performance calculation method for LH2/LOX engines with different power cycles (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A simple method for the calculation of the specific impulse of an engine with a gas generator cycle is presented. The solution is obtained by a power balance between turbine and pump. Approximate equations for the performance of the combustion products of LH2/LOX are derived. Performance results are compared with solutions of different engine types.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-64749
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  • 51
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    Parameters controlling nitric oxide emissions from gas turbine combustors (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Nitric oxide forms in the primary zone of gas turbine combustors where the burnt gas composition is close to stoichiometric and gas temperatures are highest. It was found that combustor air inlet conditions, mean primary zone fuel-air ratio, residence time, and the uniformity of the primary zone are the most important variables affecting nitric oxide emissions. Relatively simple models of the flow in a gas turbine combustor, coupled with a rate equation for nitric oxide formation via the Zeldovich mechanism are shown to correlate the variation in measured NOx emissions. Data from a number of different combustor concepts are analyzed and shown to be in reasonable agreement with predictions. The NOx formation model is used to assess the extent to which an advanced combustor concept, the NASA swirl can, has produced a lean well-mixed primary zone generally believed to be the best low NOx emissions burner type.
    Keywords: PROPULSION SYSTEMS
    Type: AGARD Atmospheric Pollution by Aircraft Engines; 16 p
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  • 52
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    The starting transient of solid propellant rocket motors with high internal gas velocities (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A comprehensive analytical model which considers time and space development of the flow field in solid propellant rocket motors with high volumetric loading density is described. The gas dynamics in the motor chamber is governed by a set of hyperbolic partial differential equations, that are coupled with the ignition and flame spreading events, and with the axial variation of mass addition. The flame spreading rate is calculated by successive heating-to-ignition along the propellant surface. Experimental diagnostic studies have been performed with a rectangular window motor (50 cm grain length, 5 cm burning perimeter and 1 cm hydraulic port diameter), using a controllable head-end gaseous igniter. Tests were conducted with AP composite propellant at port-to-throat area ratios of 2.0, 1.5, 1.2, and 1.06, and head-end pressures from 35 to 70 atm. Calculated pressure transients and flame spreading rates are in very good agreement with those measured in the experimental system.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-136274 , AMS-1100 , AMS-1100-T
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  • 53
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    Impact resistance of fiber composite blades used in aircraft turbine engines (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Resistance of advanced fiber reinforced epoxy matrix composite materials to ballistic impact was investigated as a function of impacting projectile characteristics, and composite material properties. Ballistic impact damage due to normal impacts, was classified as transverse (stress wave delamination and splitting), penetrative, or structural (gross failure). Steel projectiles were found to be gelatin ice projectiles in causing penetrative damage leading to reduced tensile strength. Gelatin and ice projectiles caused either transverse or structural damage, depending upon projectile mass and velocity. Improved composite transverse tensile strength, use of dispersed ply lay-ups, and inclusion of PRD-49-1 or S-glass fibers correlated with improved resistance of composite materials to transverse damage. In non-normal impacts against simulated blade shapes, the normal velocity component of the impact was used to correlate damage results with normal impact results. Stiffening the leading edge of simulated blade specimens led to reduced ballistic damage, while addition of a metallic leading edge provided nearly complete protection against 0.64 cm diameter steel, and 1.27 cm diameter ice and gelatin projectiles, and partial protection against 2.54 cm diameter projectiles of ice and gelatin.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134502 , PWA-4727
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  • 54
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    Experimental evaluation of a TF30-P-3 turbofan engine in an altitude facility: Effect of steady-state temperature distortion (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The effects of circumferential distortion of the total temperature entering 25, 50, and 75 percent of the inlet circumferential annulus of a turbofan engine were determined. Complete compressor stall resulted from distortions of from 14 to 20 percent of the face averaged temperature. Increasing the temperature level in one sector resulted in that sector moving toward stall by decreasing the equivalent rotor speeds while the pressure ratio remained approximately constant. Stall originated as a rotating zone in the low-pressure compressor which resulted as a terminal stall in the high-pressure compressor. Decreasing the Reynolds number index to 0.25 from 0.5 reduced the required distortion for stall by 50 percent for the conditions investigated.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2921 , E-7499
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  • 55
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    Microfog lubricant application system for advanced turbine engine components, phase 3 (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The wetting characteristics and deposit forming tendencies of a series of lubricants were evaluated using a microfog jet delivery system to wet a flat heated rotating disc. The performances of the nine lubricants are discussed in terms of the various testing parameters which include temperature, disc speed and lubricant gas flow rates. Also discussed are the heat transfer characteristics of two of the lubricants on that same plane disc specimen. The wetting characteristics and heat transfer characteristics of one of the lubricants on a complex disc simulating bearing geometry are also discussed.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121271
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  • 56
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    Processing of thermionic power on an electrically propelled spacecraft (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A study to define the power processing equipment required between a thermionic reactor and an array of mercury-ion thrusters for a nuclear electric propulsion system is reported. Observations and recommendations that resulted from this study were: (1) the preferred thermionic-fuel-element source voltages are 23 V or higher; (2) transistor characteristics exert a strong effect on power processor mass; (3) the power processor mass could be considerably reduced should the magnetic materials that exhibit low losses at high frequencies, that have a high Curie point, and that can operate at 15 to 20 kG become avaliable; (4) electrical component packaging on the radiator could reduce the area that is sensitive to meteoroid penetration, thereby reducing the meteoroid shielding mass requirement; (5) an experimental model of the power processor design should be built and tested to verify the efficiencies, masses, and all the automatic operational aspects of the design.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-135941 , JPL-TM-33-618
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  • 57
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    Propulsion system design for the ATT. (1973)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: The NASA funded Advanced Transport Technology (ATT) systems studies are directed at identifying the optimum propulsion system characteristics required for a low noise, low emissions level engine designed for an advanced commercial transport that employs the supercritical wing technology. This transport could be in service in the late 70s or early 80s and would be designed for transcontinental and international ranges with cruise speeds up to Mach 0.98. This paper reviews the significant results of the propulsion system study, the implications in the propulsion design concept, and the acoustically treated nacelle.
    Keywords: PROPULSION SYSTEMS
    Type: AIAA PAPER 72-760
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  • 58
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    Restart transients of hybrid rocket engines. (1973)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: Experimental investigation of the problems associated with restarting hybrid rocket motors (i.e., motors wherein a liquid or gaseous oxidizer is injected into the port of a solid fuel grain with subsequent mixing and combustion of the oxidizer and fuel) following a brief period of extinguishment. The results include the finding that the ignition delay on restart is decreased because less energy is absorbed by the fuel before the surface reaches the ignition point.
    Keywords: PROPULSION SYSTEMS
    Type: Journal of Spacecraft and Rockets; 10; Mar. 197
    Format: text
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    Ion thruster thermal characteristics and performance. (1973)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: Experimental and analytical results for a typical 20-cm-diam, hollow-cathode ion thrustor are reported. The foundation of the investigation was the application of thermal model correction techniques. Pertinent thermal properties and plasma heating characteristics of the thrustor were determined through correlation and integration of temperature measurement data with a single-state Wiener-Kalman filter. The thrustor self-heating levels on various parts were realistically estimated. Analytically predicted temperatures were forced to agree with the measured values for the purpose of constructing a corrected thermal model, which could then be used to evaluate more realistic thrustor circumstances and environments. The expected accuracy of the resultant analytical network model was demonstrated to be plus or minus 10 K. Thrustor thermal performance data for a typical five-thrustor array are presented as functions of environmental solar intensities. The thermal analyses are also extended to a 30-cm thrustor system.
    Keywords: PROPULSION SYSTEMS
    Type: Journal of Spacecraft and Rockets; 10; Jan. 197
    Format: text
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  • 60
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    Electrostatic Propulsion Beam Divergence Effects on Spacecraft Surfaces. Volume 2, Addendum 1: Ion Time-of-flight Determinations of Doubly to Singly Ionized Mercury Ion Ratios from a Mercury Electron Bombardment Discharge (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The analysis of ion exhaust beam current flow for multiply charged ion species and the application to propellant utilization for the thruster are discussed. The ion engine in use in the experiments is a twenty centimeter diameter electromagnet electron bombardment engine. The experimental technique to determine the multiply charged ion abundance ratios using ion time of flight is described. An analytical treatment of the discharge action in producing various ion species has been carried out.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-131841 , TRW-11985-6002-RU-01-VOL-2-ADD
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    Ion thruster with a combination keeper electrode and electron baffle (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An ion thruster is described in which the cathode front end, surrounded by our insulator, is mounted flush with the front end of the flanged portion of the cathode pole piece. The thruster's baffle positioned in front of the cathode's front end supports the thruster's keeper electrode which is space apart and directed to the cathode's open end. The baffle is at the keeper's electrode potential.
    Keywords: PROPULSION SYSTEMS
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    Experimental and data analysis techniques for deducing collision-induced forces from photographic histories of engine rotor fragment impact/interaction with a containment ring (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An analysis method termed TEJ-JET is described whereby measured transient elastic and inelastic deformations of an engine-rotor fragment-impacted structural ring are analyzed to deduce the transient external forces experienced by that ring as a result of fragment impact and interaction with the ring. Although the theoretical feasibility of the TEJ-JET concept was established, its practical feasibility when utilizing experimental measurements of limited precision and accuracy remains to be established. The experimental equipment and the techniques (high-speed motion photography) employed to measure the transient deformations of fragment-impacted rings are described. Sources of error and data uncertainties are identified. Techniques employed to reduce data reading uncertainties and to correct the data for optical-distortion effects are discussed. These procedures, including spatial smoothing of the deformed ring shape by Fourier series and timewise smoothing by Gram polynomials, are applied illustratively to recent measurements involving the impact of a single T58 turbine rotor blade against an aluminum containment ring. Plausible predictions of the fragment-ring impact/interaction forces are obtained by one branch of this TEJ-JET method; however, a second branch of this method, which provides an independent estimate of these forces, remains to be evaluated.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134548 , ASRL-TR-154-5
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    Solar electric propulsion/instrument/subsystems interaction study (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The interactive effects between a solar electric propulsion system and an electrically propelled scientific spacecraft were examined. The operation of the ion thrusters may impact upon the acquisition and interpretation of data by the science payload of the spacecraft. The effluents from the operation of the electric propulsion unit may also impact upon the operation of the various subsystems of the vehicle. Specific interactive effects were isolated where meaningful levels of interaction may occur. The level of impact upon elements of the science payload and other affected subsystems is examined, and avenues for the reduction or elimination of impact are defined.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-114732 , TRW-22878-6007-RU-00
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  • 64
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    Hydrogen-methane fuel control systems for turbojet engines (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Design, development, and test of a fuel conditioning and control system utilizing liquid methane (natural gas) and liquid hydrogen fuels for operation of a J85 jet engine were performed. The experimental program evaluated the stability and response of an engine fuel control employing liquid pumping of cryogenic fuels, gasification of the fuels at supercritical pressure, and gaseous metering and control. Acceptably stable and responsive control of the engine was demonstrated throughout the sea level power range for liquid gas fuel and up to 88 percent engine speed using liquid hydrogen fuel.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121247 , R74AEG153 , TM-73-489
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    Analysis of inlet flow distortion and turbulence effects on compressor stability (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The effect of steady state circumferential total pressure distortion on the loss in compressor stall pressure ratio has been established by analytical techniques. Full scale engine and compressor/fan component test data were used to provide direct evaluation of the analysis. Specifically, since a circumferential total pressure distortion in an inlet system will result in unsteady flow in the coordinate system of the rotor blades, analysis of this type distortion must be performed from an unsteady aerodynamic point of view. By application of the fundamental aerothermodynamic laws to the inlet/compressor system, parameters important in the design of such a system for compatible operation have been identified. A time constant, directly related to the compressor rotor chord, was found to be significant, indicating compressor sensitivity to circumferential distortion is directly dependent on the rotor chord.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-114577 , TR-2-57110/3R-3071
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  • 66
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    Solar electric propulsion thrust subsystem development (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The Solar Electric Propulsion System developed under this program was designed to demonstrate all the thrust subsystem functions needed on an unmanned planetary vehicle. The demonstration included operation of the basic elements, power matching input and output voltage regulation, three-axis thrust vector control, subsystem automatic control including failure detection and correction capability (using a PDP-11 computer), operation of critical elements in thermal-vacuum-, zero-gravity-type propellant storage, and data outputs from all subsystem elements. The subsystem elements, functions, unique features, and test setup are described. General features and capabilities of the test-support data system are also presented. The test program culminated in a 1500-h computer-controlled, system-functional demonstration. This included simultaneous operation of two thruster/power conditioner sets. The results of this testing phase satisfied all the program goals.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-131360 , JPL-TR-32-1579
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    An adverse effect of film cooling on the suction surface of a turbine vane (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A J-75 size turbine vane with film cooling holes on the suction surface near the leading edge was tested with and without film cooling flow in a four vane cascade. Results show that the cooling effectiveness on the aft portion of the vane suction surface can decrease with the addition of film cooling near the leading edge. Apparently the film cooling air flow caused a laminar or transitional boundary layer to become a transitional or turbulent boundary layer. The vane was tested at a gas temperature and pressure of 1260 K (1800 F) and 22.7 newtons per square centimeter (33 psia), a coolant temperature of 280 K (50 F), film cooling flow ratios from 0.0 to 0.026, and backside midchord cooling flow ratios of 0.007 and 0.035.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-68210 , E-7392
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    An airline study of advanced technology requirements for advanced high speed commercial transport engines. 2: Engine preliminary design assessment (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The advanced technology requirements for an advanced high speed commercial transport engine are presented. The results of the phase 2 study effort cover the following areas: (1) general review of preliminary engine designs suggested for a future aircraft, (2) presentation of a long range view of airline propulsion system objectives and the research programs in noise, pollution, and design which must be undertaken to achieve the goals presented, (3) review of the impact of propulsion system unreliability and unscheduled maintenance on cost of operation, (4) discussion of the reliability and maintainability requirements and guarantees for future engines.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121133
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  • 69
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    Aerodynamic performance of a core-engine turbine stator vane tested in a two-dimensional cascade of 10 vanes and in a single vane tunnel (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A turbine stator vane was tested in a two-dimensional cascade of 10 vanes and in a single-vane tunnel. The single-vane tunnel was a cold air version of a tunnel which will be used for high temperature heat transfer testing of cooled turbine vanes. The purpose of the investigation was to determine if the flow conditions in the single-vane tunnel were sufficiently similar to those of a 10-vane cascade to permit meaningful heat transfer testing. The vane was tested over a range of ideal exit critical velocity ratios. The principal measurements were vane surface static pressure and cross-channel surveys of exit static pressure, total pressure, and flow angle. A brief description of the test vane and tunnels is included. The results of the exit surveys, the vane surface pressure distributions, and overall performance in terms of flow and loss for the two test configurations are compared.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2766 , E-7293
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    High-Tip-Speed, Low-Loading Transonic Fan Stage. Part 1: Aerodynamic and Mechanical Design (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A high-tip-speed, low-loading transonic fan stage was designed to deliver an overall pressure ratio of 1.5 with an adiabatic efficiency of 86 percent. The design flow per unit annulus area is 42.0 pounds per square foot. The fan features a hub/tip ratio of 0.46, a tip diameter of 28.74 in. and operates at a design tip speed of 1600 fps. For these design conditions, the rotor blade tip region operates with supersonic inlet and supersonic discharge relative velocities. A sophisticated quasi-three-dimensional characteristic section design procedure was used for the all-supersonic sections and the inlet of the midspan transonic sections. For regions where the relative outlet velocities are supersonic, the blade operates with weak oblique shocks only.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121095 , AIRESEARCH-72-8421-PT-1
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    Nuclear electric propulsion mission engineering study. Volume 2: Final report (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Results of a mission engineering analysis of nuclear-thermionic electric propulsion spacecraft for unmanned interplanetary and geocentric missions are summarized. Critical technologies associated with the development of nuclear electric propulsion (NEP) are assessed, along with the impact of its availability on future space programs. Outer planet and comet rendezvous mission analysis, NEP stage design for geocentric and interplanetary missions, NEP system development cost and unit costs, and technology requirements for NEP stage development are studied.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-131600 , DOC-73SD4220-VOL-2
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    Injector design guidelines for gas/liquid propellant systems (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Injector design guidelines are provided for gas/liquid propellant systems. Information was obtained from a 30-month applied research program encompassing an analytical, design, and experimental effort to relate injector design parameters to simultaneous attainment of high performance and component (injector/thrust chamber) compatibility for gas/liquid space storable propellants. The gas/liquid propellant combination studied was FLOX (82.6% F2)/ ambient temperature gaseous methane. Design criteria that provide for simultaneous attainment of high performance and chamber compatibility are presented for both injector types. Parametric data are presented that are applicable for the design of circular coaxial and like-doublet injectors that operate with design parameters similar to those employed. However, caution should be exercised when applying these data to propellant combinations whose elements operate in ranges considerably different from those employed in this study.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-120968 , R-8973-3
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    Nuclear electric propulsion mission engineering study. Volume 1: Executive summary (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Results of a mission engineering analysis of nuclear-thermionic electric propulsion spacecraft for unmanned interplanetary and geocentric missions are summarized. Critical technologies associated with the development of nuclear electric propulsion (NEP) are assessed. Outer planet and comet rendezvous mission analysis, NEP stage design for geocentric and interplanetary missions, NEP system development cost and unit costs, and technology requirements for NEP stage development are studied. The NEP stage design provides both inherent reliability and high payload mass capability. The NEP stage and payload integration was found to be compatible with the space shuttle.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-131599 , DOC-73SD4219-VOL-1
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    Low speed inducers for cryogenic upper stages (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Briefing charts are presented, which were used in an oral presentation of the results and recommendations for the design and analysis of low speed hydrogen and oxygen inducers and their drive systems applicable to the space tug. A discussion of the design of the 15K and RL-10 inducers is included.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-124233 , ASR-73-25
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    Design for pressure regulating components (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The design development for Pressure Regulating Components included a regulator component trade-off study with analog computer performance verification to arrive at a final optimized regulator configuration for the Space Storable Propulsion Module, under development for a Jupiter Orbiter mission. This application requires the pressure regulator to be capable of long-term fluorine exposure. In addition, individual but basically identical (for purposes of commonality) units are required for separate oxidizer and fuel pressurization. The need for dual units requires improvement in the regulation accuracy over present designs. An advanced regulator concept was prepared featuring redundant bellows, all metallic/ceramic construction, friction-free guidance of moving parts, gas damping, and the elimination of coil springs normally used for reference forces. The activities included testing of actual size seat/poppet components to determine actual discharge coefficients and flow forces. The resulting data was inserted into the computer model of the regulator. Computer simulation of the propulsion module performance over two mission profiles indicated satisfactory minimization of propellant residual requirements imposed by regulator performance uncertainties.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-139300 , S-1259
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    Design study of RL10 derivatives. Volume 3, part 2: Operational and flight support plan (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Transportation requirements are considered during the engine design layout reviews and maintenance engineering analyses. Where designs cannot be influenced to avoid transportation problems, the transportation representative is advised of the problems permitting remedies early in the program. The transportation representative will monitor and be involved in the shipment of development engine and GSE hardware between FRDC and vehicle manufacturing plant and thereby will be provided an early evaluation of the transportation plans, methods and procedures to be used in the space tug support program. Unanticipated problems discovered in the shipment of development hardware will be known early enough to permit changes in packaging designs and transportation plans before the start of production hardware and engine shipments. All conventional transport media can be used for the movement of space tug engines. However, truck transport is recommended for ready availability, variety of routes, short transit time, and low cost.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-120148 , FR-6011-VOL-3-PT-2
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  • 77
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    Design study of RL10 derivatives. Volume 2: Engine design characteristics, appendices (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Calculations, curves, and substantiating data which support the engine design characteristics of the RL-10 engines are presented. A description of the RL-10 ignition system is provided. The performance calculations of the RL-10 derivative engines and the performance results obtained are reported. The computer simulations used to establish the control system requirements and to define the engine transient characteristics are included.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-120144 , FR-6011-VOL-2-APP
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    Design study of RL10 derivatives. Volume 3, part 1: Preliminary interface control document (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The Interface Control Document contains engine information necessary for installation of the baseline RL10 Derivative engines in the Space Tug vehicle. The ICD presents a description of the baseline engines and their operating characteristics, mass and load characteristics, and environmental criteria. The document defines the engine/vehicle mechanical, electrical, fluid and pneumatic interface requirements.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-120146 , FR-6011-VOL-3-PT-1
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    Design study of RL10 derivatives. Volume 2: Engine design characteristics (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The design characteristics of the RL-10 rocket engine are discussed. The results from critical elements evaluation, baseline engine design, parametric and special study tasks are presented. Critical element evaluation established the feasibility of various engine features such as tank head idle, pumped idle, autogenous tank pressurization, and two-phase pumping. Three baseline engines, derived from the RL-10 were conceptually designed. Parametric life and performance data were generated. Special studies were conducted to establish the impact on the engine design of environment, safety, interchangeability, and maintenance.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-120145 , FR-6011-VOL-2
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  • 80
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    Space vehicle engine and heat shield environment review. Volume 1: Engineering analysis (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Methods for predicting the base heating characteristics of a multiple rocket engine installation are discussed. The environmental data is applied to the design of adequate protection system for the engine components. The methods for predicting the base region thermal environment are categorized as: (1) scale model testing, (2) extrapolation of previous and related flight test results, and (3) semiempirical analytical techniques.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-120243 , EE-MSFC-1774
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    Development of helicopter engine seals (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An experimental evaluation of main shaft seals for helicopter gas turbine engines was conducted with shaft speeds to 213 m/s(700 ft/sec), air pressures to 148 N/sq cm (215 psia), and air temperatures to 645 K (675 F). Gas leakage test results indicate that conventional seals will not be satisfactory for high-pressure sealing because of excessive leakage. The self-acting face seal, however, had significantly lower leakage and operated with insignificant wear during a 150-hour endurance test at sliding speeds to 145 m/s (475 ft/sec), air pressures to 124 N/sq cm (180 psia), and air temperatures to 408 K (275 F). Wear measurements indicate that noncontact operation was achieved at shaft speeds of 43,000 rpm. Evaluation of the self-acting circumferential seal was inconclusive because of seal dimensional variations.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134647 , LYC-73-48
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    Acoustic testing of a 1.5 pressure ratio low tip speed fan with a serrated rotor (QEP fan B scale model) (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A scale model of the bypass flow region of a 1.5 pressure ratio, single stage, low tip speed fan was tested with a serrated rotor leading edge to determine its effects on noise generation. The serrated rotor was produced by cutting teeth into the leading edge of the nominal rotor blades. The effects of speed and exhaust nozzle area on the scale models noise characteristics were investigated with both the nominal rotor and serrated rotor. Acoustic results indicate the serrations reduced front quadrant PNL's at takeoff power. In particular, the 200 foot (61.0 m) sideline noise was reduced from 3 to 4 PNdb at 40 deg for nominal and large nozzle operation. However, the rear quadrant maximum sideline PNL's were increased 1.5 to 3 PNdb at approach thust and up to 2 PNdb at takeoff thust with these serrated rotor blades. The configuration with the serrated rotor produced the lowest maximum 200 foot (61.0 m) sideline PNL for any given thust when the large nozzle (116% of design area) was employed.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-120846
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    Liquid rocket valve components (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A monograph on valves for use with liquid rocket propellant engines is presented. The configurations of the various types of valves are described and illustrated. Design criteria and recommended practices for the various valves are explained. Tables of data are included to show the chief features of valve components in use on operational vehicles.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-SP-8094
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  • 84
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    Program on ground test of modified quiet, clean, JT3D and JT8D turbofan engines in their respective nacelles (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A program to reduce the community noise levels of commercial jet aircraft is summarized. The program objective is the development of three acoustically treated nacelle configurations for the 707, 727, and 737 series aircraft to provide maximum noise reduction with minimum performance loss, modification requirements, and economic impact. The preliminary design, model testing, data analyses, and economic studies of proposed nacelle configurations are discussed.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134553 , D6-41244
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  • 85
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    Theoretical investigations on plasma processes in the Kaufman thruster (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The lateral neutralization of ion beams is treated by standard mathematical methods for first order, nonlinear partial differential equations. A closed form analytical solution is derived for the transient lateral beam neutralization for electron injection by means of a von Mises transformation. A nonlinear theory of the longitudinal ion beam neutralization is developed using the von Mises transformation. By means of the Lenard-Balescu equation, the intercomponent momentum transfer between stable, collisionless electron and ion components is calculated.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134637
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  • 86
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    Quiet engine program: Turbine noise suppression. Volume 2: Treatment selection, installation, and test results (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: For abstract, see N74-19400.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134586 , R73AEG443-VOL-2
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  • 87
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    Quiet engine program: Turbine noise suppression. -Volume 1: General treatment evaluation and measurement techniques (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Acoustic treatment was developed for jet engine turbine noise suppression. Acoustic impedance and duct transmission loss measurements were made for various suppression systems. An environmental compatibility study on several material types having suppression characteristics is presented. Two sets of engine hardware were designed and are described along with engine test results which include probe, farfield, near field, and acoustic directional array data. Comparisons of the expected and the measured suppression levels are given as well as a discussion of test results and design techniques.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134499 , R73AEG443-VOL-1
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  • 88
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    Space shuttle RCS engine (1973)
    In:  CASI
    Details
    Publication Date: 2019-07-20
    Description: The design of the space shuttle RCS engine has the primary objective of reusability with minimum servicing. Engine S/N FT-2A has successfully completed all ten environmental (salt water spray, sand and dust, vibration and humidity) and hot fire cycles with no change in engine performance (steady state or pulse mode).
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-136021 , REPT-8701-910013
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  • 89
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    The Space Shuttle Main Engine (1973)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: The Space Shuttle Main Engine is the highest performance large rocket engine being developed in the world. This engine will be used to put the Space Shuttle orbiter vehicle into earth orbit and will be the workhorse propulsion system of the 1970s and 1980s. The propulsion requirements demand maximum performance from the liquid oxygen/hydrogen propellants with a minimum hardware weight. The system is being developed at low cost with a required early flight schedule of 1978. The engine system and components are based on the latest technology using advanced materials and high-pressure combustion. Engine weight and performance are discussed. This paper describes engine design, the program and present progress, the technology of advanced components, experimental results, and development areas underway.
    Keywords: PROPULSION SYSTEMS
    Type: International Symposium on Space Technology and Science; Sep 03, 1973 - Sep 08, 1973; Tokyo; Japan
    Format: text
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  • 90
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    Extension of a simplified computer program for analysis of solid-propellant rocket motors (1973)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A research project to develop a computer program for the preliminary design and performance analysis of solid propellant rocket engines is discussed. The following capabilities are included as computer program options: (1) treatment of wagon wheel cross sectional propellant configurations alone or in combination with circular perforated grains, (2) calculation of ignition transients with the igniter treated as a small rocket engine, (3) representation of spherical circular perforated grain ends as an alternative to the conical end surface approximation used in the original program, and (4) graphical presentation of program results using a digital plotter.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-129024
    Format: application/pdf
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  • 91
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    Electrostatic propulsion beam divergence effects on spacecraft surfaces, volume 3 (1973)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Tests were conducted to determine the effects of electrostatic propulsion beam divergence effects on spacecraft surfaces. The subjects discussed are: (1) sensitive surfaces on the ATS 6 spacecraft, (2) the cesium ion source and testing facility, (3) cesium ion effects on thermophysical properties, and (4) simulated charge-exchange ion exposure. The compatibility of the ATS 6 ion engine experiment with the engineering subsystems and other experiments aboard the ATS 6 spacecraft was analyzed.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-136933 , TRW-11985-6003-RU-00
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  • 92
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    Power processor for a 20CM ion thruster (1973)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A power processor breadboard for the JPL 20CM Ion Engine was designed, fabricated, and tested to determine compliance with the electrical specification. The power processor breadboard used the silicon-controlled rectifier (SCR) series resonant inverter as the basic power stage to process all the power to the ion engine. The breadboard power processor was integrated with the JPL 20CM ion engine and complete testing was performed. The integration tests were performed without any silicon-controlled rectifier failure. This demonstrated the ruggedness of the series resonant inverter in protecting the switching elements during arcing in the ion engine. A method of fault clearing the ion engine and returning back to normal operation without elaborate sequencing and timing control logic was evolved. In this method, the main vaporizer was turned off and the discharge current limit was reduced when an overload existed on the screen/accelerator supply. After the high voltage returned to normal, both the main vaporizer and the discharge were returned to normal.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121160 , TRW-20384-6002-RU-00
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  • 93
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    Hollow cathode restartable 15 cm diameter ion thruster (1973)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The effects of substituting high perveance dished grids for low perveance flat ones on performance variables and plasma properties within a 15 cm modified SERT II thruster are discussed. Results suggest good performance may be achieved as an ion thruster is throttled if the screen grid transparency is decreased with propellant flow rate. Thruster startup tests, which employ a pulsed high voltage tickler electrode between the keeper and the cathode to initiate the discharge, are described. High startup reliability at cathode tip temperatures of about 500 C without excessive component wear over 2000 startup cycles is demonstrated. Testing of a single cusp magnetic field concept of discharge plasma containment is discussed. A theory which explains the observed behavior of the device is presented and proposed thruster modifications and future testing plans are discussed.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134532
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  • 94
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    Design considerations for the airframe-integrated scramjet (1973)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Research programs at the NASA Langley Research Center on the development of airframe-integrated scramjet concepts (supersonic combustion ramjet) are reviewed briefly. The design and performance of a specific scramjet configuration are examined analytically by use of recently developed and substantiated techniques on boundary-layer development, heat transfer, fuel-air mixing, heat-release rates, and engine-cycle analysis. These studies indicate that the fixed-geometry scramjet module will provide practical levels of thrust performance with low cooling requirements. Areas which need particular emphasis in further development work are the combustor design for low speeds and the integrated nozzle design.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2895 , L-8152 , Intern. Symp. on Air Breathing Engines; Jun 01, 1972; Marseille; France
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  • 95
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    Advanced hydrogen/oxygen thrust chamber design analysis (1973)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The results are reported of the advanced hydrogen/oxygen thrust chamber design analysis program. The primary objectives of this program were to: (1) provide an in-depth analytical investigation to develop thrust chamber cooling and fatigue life limitations of an advanced, high pressure, high performance H2/O2 engine design of 20,000-pounds (88960.0 N) thrust; and (2) integrate the existing heat transfer analysis, thermal fatigue and stress aspects for advanced chambers into a comprehensive computer program. Thrust chamber designs and analyses were performed to evaluate various combustor materials, coolant passage configurations (tubes and channels), and cooling circuits to define the nominal 1900 psia (1.31 x 10 to the 7th power N/sq m) chamber pressure, 300-cycle life thrust chamber. The cycle life capability of the selected configuration was then determined for three duty cycles. Also the influence of cycle life and chamber pressure on thrust chamber design was investigated by varying in cycle life requirements at the nominal chamber pressure and by varying the chamber pressure at the nominal cycle life requirement.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121213 , R-9258
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  • 96
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    Plume mass flow and optical damage distributions for an MMH/N2O4 RCS thruster (1973)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The data obtained from two recent experiments conducted in a continuing series of experiments at the Lewis Research Center into the contamination characteristics of a 5-pound thrust MMH/N2O4 engine are presented. The primary objectives of these experiments were to establish the angular distribution of condensible exhaust products within the plume and the corresponding optical damage angular distribution of transmitting optical elements attributable to this contaminant. The plume mass flow distribution was measured by five quartz crystal microbalances (QCM's) located at the engine axis evaluation. The fifth QCM was located above the engine and 15 deg behind the nozzle exit plane. The optical damage was determined by ex-situ transmittance measurements for the wavelength range from 0.2 to 0.6 microns on 2.54 cm diameter fused silica discs also located at engine centerline elevation. Both the mass deposition and optical damage angular distributions followed the expected trend of decreasing deposition and damage as the angle between sensor or sample and the nozzle axis increased. A simple plume gas flow equation predicted the deposition distribution reasonably well for angles of up to 55 degrees. The optical damage measurements also indicated significant effects at large angles.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-71465 , Space Simulation Conf.; Nov 12, 1973 - Nov 14, 1973; Los Angeles
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  • 97
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    Space shuttle main engine controller assembly, phase C-D (1973)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: System design and system analysis and simulation are slightly behind schedule, while design verification testing has improved. Input/output circuit design has improved, but digital computer unit (DCU) and mechanical design continue to lag. Part procurement was impacted by delays in printed circuit board, assembly drawing releases. These are the result of problems in generating suitable printed circuit artwork for the very complex and high density multilayer boards.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-135947 , HONEYWELL-W2101-QPR-3-73 , QPR-6
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  • 98
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    Acoustic investigation of the engine-over-the-wing concept using a D-shaped nozzle. (1973)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: Small-model experiments were conducted of the engine-over-the-wing concept using a D-shaped nozzle in order to determine the static-lift and acoustic characteristics at two wing-flap positions. Configurations were tested with the flow attached and unattached to the upper surface of the flaps. Attachment was obtained with a nozzle flow deflector. In both cases, high frequency noise shielding by the wing was obtained. Configurations using the D-shaped nozzle are compared with corresponding ones using a circular nozzle. With flow attached to the flaps, the static lift and acoustic results are almost the same for both nozzles. Without the nozzle flow deflector (unattached flap flow), the D-nozzle is considerably noisier than a circular nozzle in the low and middle frequencies.
    Keywords: PROPULSION SYSTEMS
    Type: AIAA PAPER 73-1030 , Aero-Acoustics Conference; Oct 15, 1973 - Oct 17, 1973; Seattle, WA
    Format: text
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  • 99
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    Multiple pure tone noise generation and control. (1973)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: The generation of multiple pure tones in supersonic fans is discussed. The theoretical results of Kurasaka are reviewed and compared with experimental data obtained on a 36-in. diameter, 1550 ft/sec, 1.6 pressure ratio fan. Detailed measurements on bow shock locations taken with pressure transducers indicate that blade to blade discrepancies are the source of MPT generation. The paper presents some experimental results on an attempt to reduce the shock strength, and subsequently the MPT's, through blade modifications. Other attempts at reducing the MPT's through wall treatment, high inlet flow Mach number, acoustically treated splitters - are discussed. Experimental data is presented on the validity of these noise reduction methods.
    Keywords: PROPULSION SYSTEMS
    Type: AIAA PAPER 73-1021 , Aero-Acoustics Conference; Oct 15, 1973 - Oct 17, 1973; Seattle, WA
    Format: text
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  • 100
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    Noise comparisons from full-scale fan tests at NASA Lewis Research Center. (1973)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: The overall aero and acoustic design features of eight 6-foot-diameter, single-stage fans tested in an outdoor acoustic facility are described. A correlation of the acoustic results for subsonic tip-speed fans showed the total sound power to be proportional to the mechanical power imparted to the fan and the specific work performed on the air to within plus or minus 2 dB. The correlation was relatively insensitive to fan design variables over a broad range of operating conditions. Maximum perceived noise levels were generally proportional to the sound power levels with both noise levels exhibiting a relatively unique increase with fan pressure ratio when normalized by the delivered thrust. The spectra of broadband noise attributed to the fan exhibited a bimodal characteristic for most of the fans. A predominant mode centered near the blade-passage tone and another at 8 to 16 times the tone frequency.
    Keywords: PROPULSION SYSTEMS
    Type: AIAA PAPER 73-1017 , Aero-Acoustics Conference; Oct 15, 1973 - Oct 17, 1973; Seattle, WA
    Format: text
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