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  • General Chemistry  (1,705)
  • PROPULSION SYSTEMS  (349)
  • Cell & Developmental Biology
  • 1990-1994
  • 1970-1974  (1,557)
  • 1910-1914  (703)
  • 1972  (1,557)
  • 1910  (703)
Collection
Keywords
Publisher
Years
  • 1990-1994
  • 1970-1974  (1,557)
  • 1910-1914  (703)
Year
  • 1
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    Pulse performance analysis for small hypergolic propellant rocket engines (1972)
    In:  CASI
    Details
    Publication Date: 2005-11-30
    Description: Small rocket engine tests were conducted for the purpose of obtaining pulse performance data to aid in preliminary design and evaluation of attitude control systems. Both monopropellant and hypergolic bipropellant engines of thrust levels from 5 to 445 N (1 to 100 lb) were tested. The performance data for the hypergolic propellant rockets are compared with theoretical performance calculated from idealized chamber filling and evacuation characteristics. Electromechanical delays in valve response and heat transfer characteristics were found to cause substantial deviation between theoretical performance and test performance. The theoretical analysis is modified to obtain a semiempirical model for hypergolic propellant rockets.
    Keywords: PROPULSION SYSTEMS
    Type: Res. Achievements Rev., Vol. 4, No. 6; p 61-74
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  • 2
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    Extendible rocket nozzle development (1972)
    In:  CASI
    Details
    Publication Date: 2005-11-30
    Description: An evaluation is presented of J-2 engine modifications that will simplify operation and improve reliability of the advanced Saturn 1C launch vehicle. Methods of increasing thrust without extensively modifying the S-2 or S-4B stages are also evaluated. A thrust increase was achieved by raising engine combustion through a redesign of the engine thrust chamber and propellant feed system.
    Keywords: PROPULSION SYSTEMS
    Type: Res. Achievements Rev., Vol. 4, No. 6; p 105-115
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  • 3
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    Velocity diagrams (1972)
    In:  CASI
    Details
    Publication Date: 2005-11-30
    Description: The selection and design of velocity diagrams for axial flow turbines are considered. Application is treated in two parts which includes: (1) mean-section diagrams, and (2) radial variation of diagrams. In the first part, the velocity diagrams occurring at the mean section are assumed to represent the average conditions encountered by the turbine. The different types of diagrams, their relation to stage efficiency, and their selection when staging is required are discussed. In the second part, it is shown that in certain cases the mean-section diagrams may or may not represent the average flow conditions for the entire blade span. In the case of relatively low hub- to tip-radius ratios, substantial variations in the velocity diagrams are encountered. The radial variations in flow conditions and their effect on the velocity diagrams are considered.
    Keywords: PROPULSION SYSTEMS
    Type: Turbine Design and Appl., Vol. 1; p 69-99
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  • 4
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    Basic turbine concepts (1972)
    In:  CASI
    Details
    Publication Date: 2005-11-30
    Description: Turbine geometric, flow, energy transfer, efficiency, and performance characteristics are considered by the use of definitions, diagrams, and dimensionless parameters. Emphasis is placed on the determination of the fluid velocity as it passes from one blade row to the next. The general methods for constructing velocity diagrams and relating them to the work and flow capacity of the turbine are discussed.
    Keywords: PROPULSION SYSTEMS
    Type: Turbine Design and Appl., Vol. 1; p 21-67
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  • 5
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    Minimum impulse tests of 0.45-N liquid hydrazine catalytic thrusters (1972)
    In:  CASI
    Details
    Publication Date: 2006-02-22
    Description: An exploratory test series was conducted on three types of 0.45-N (0.1 lbf) liquid hydrazine thrusters to ascertain the minimum impulse bit capability for this class of engine. The test series is described and the results are presented. The testing was performed at 21 and 145 C (70 and 300 F) while maintaining nominal 0.45 N (0.1 lbf) upstream conditions. Valve on-times as low as 0.008 sec were applied. Impulse bits were observed for thruster temperatures of 21 and 145 C (70 and 300 F), respectively.
    Keywords: PROPULSION SYSTEMS
    Type: JPL Quarterly Tech. Rev., Vol. 2, No. 1; p 107-112
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  • 6
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    Mariner 9 propulsion subsystem performance during interplanetary cruise and Mars orbit insertion (1972)
    In:  CASI
    Details
    Publication Date: 2006-02-22
    Description: On 14 November 1971 the Mariner 9 1334-N-(300-lbf)-thrust rocket engine was fired for just over 15 min to place the first man-made satellite into orbit about Mars. Propulsion subsystem data gathered during the 5-month interplanetary cruise and orbit insertion are of significance to future missions of this type. Specific results related to performance predictability, zero g heat transfer, and nitrogen permeation, diffusion, and solubility values are presented.
    Keywords: PROPULSION SYSTEMS
    Type: JPL Quarterly Tech. Rev., Vol. 2, No. 1; p 113-122
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  • 7
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    Solar-electric propulsion breadboard thrust subsystem (1972)
    In:  CASI
    Details
    Publication Date: 2006-02-22
    Description: A solar-electric propulsion breadboard thrust subsystem has been designed, built, and tested. A 1500-h test was performed to demonstrate the functional capabilities of the subsystem. Described are the subsystem functions and testing process. The results show that the ground work has been established for development of an engineering model of the thrust subsystem.
    Keywords: PROPULSION SYSTEMS
    Type: JPL Quart. Tech. Rev., Vol. 2, No. 2; p 100-112
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  • 8
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    STOL propulsion systems (1972)
    In:  CASI
    Details
    Publication Date: 2006-03-27
    Description: The selection and the characteristics of quiet, clean propulsion systems for STOL aircraft are discussed. Engines are evaluated for augmentor wing and externally blown flap STOL aircraft with the engines located both under and over the wings. Some supporting test data are presented. Optimum engines are selected based on achieving the performance, economic, acoustic, and pollution goals presently being considered for future STOL aircraft. The data and results presented were obtained from a number of contracted studies and some supporting NASA inhouse programs, most of which began in early 1972. The contracts include: (1) two aircraft and mission studies, (2) two propulsion system studies, (3) the experimental and analytic work on the augmentor wing, and (4) the experimental programs on Q-Fan. Engines are selected and discussed based on aircraft economics using the direct operating cost as the primary criterion. This cost includes the cost of the crew, fuel, aircraft, and engine maintenance and depreciation.
    Keywords: PROPULSION SYSTEMS
    Type: STOL Technol.; p 475-509
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  • 9
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    Engine noise technology (1972)
    In:  CASI
    Details
    Publication Date: 2006-03-27
    Description: The characteristics of aircraft engine noise are discussed. Data are provided to show the noise produced by the following aircraft components: (1) fan noise, (2) noise suppressing structures, (3) sonic inlets, (4) jet mixing noise due to nozzle flow, and (5) thrust reversers. Charts are developed to show the sound pressure level and the frequencies for each type of noise source. The use of laminates and composite materials to dissipate acoustic power is examined.
    Keywords: PROPULSION SYSTEMS
    Type: STOL Technol.; p 371-412
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  • 10
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    Quiet STOL propulsion session introductory remarks (1972)
    In:  CASI
    Details
    Publication Date: 2006-03-27
    Description: Research activities, preliminary design activities, and system optimization studies in support of the development of advanced, quiet, STOL propulsion systems are discussed. Noise alleviation by means of controlling the source and by means of acoustical treatment receive considerable emphasis. A STOL airplane designed for a given payload has essentially double the installed thrust of a comparable CTOL airplane. Unless compensated for during the design process, this alone will tend to increase the source noise by 3 db. The propulsive lift introduces flap impingement noise or duct and flap scrubbing noise, noise sources not present in CTOL airplanes to any significant degree. These additional noise sources are illustrated. Depending on the specific configuration, this will tend to increase the noise by several db or more. Although the propulsive lift characteristics of STOL airplanes will tend to increase source noise significantly, the proximity of STOL airfields to populated areas leads to STOL noise objectives considerably lower than those currently applicable to CTOL airplanes.
    Keywords: PROPULSION SYSTEMS
    Type: STOL Technol.; p 367-370
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  • 11
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    Preliminary results of large supersonic burning combustor testing. (1972)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: A hydrogen fueled supersonic-burning combustor 18 in. in diameter, which is equivalent to that of an engine about 6 ft in diameter, was tested as a direct-connected duct at inlet conditions which simulated Mach 8 flight at 115,000 ft alt. A synthetic air consisting of oxygen with 39% nitrogen and 38% water vapor at a total temperature of 4500 deg R and a total pressure of 300 psia was supplied to the combustor inlet by a hydrazine-nitrogen tetroxide hot gas generator which maintained a uniform inlet flow Mach number of 2.8. The large combustor size required a new approach to fuel injector design. Some hydrogen was injected through flush-wall injectors, but most was injected from two rows of swept and tapered struts immersed in the flow stream. Supersonic combustion was obtained at hydrogen equivalence ratios of 0.94 without encountering thermal cho king. Wall static pressures, and the radial distribution of hydrogen, Pitot pressure, and Mach number were determined at the combustor exit.
    Keywords: PROPULSION SYSTEMS
    Type: Journal of Aircraft; 9; Jan. 197
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  • 12
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    Effect on supersonic jet noise of nozzle plenum pressure fluctuations. (1972)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: Demonstration that pressure fluctuations in the plenum chamber to a supersonic nozzle can strongly increase the noise radiated from the jet plume. The correlation shows that jet noise acoustic efficiency increases from 0.3% to 0.8% (or 4 dB) when the chamber roughness intensity increases from essentially no plenum chamber roughness to 2%. A roughness level of 2% has been observed in some turbojet engines. It is concluded that the reduction or elimination of plenum chamber pressure fluctuations may be an important method of reducing the total noise from jet engines.
    Keywords: PROPULSION SYSTEMS
    Type: AIAA Journal; 10; July 197
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  • 13
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    Electrodynamic sailing - Beating into the solar wind. (1972)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: The recent suggestion by Alfven (1972) of a novel means of spacecraft propulsion based upon energy extraction from the electromagnetic field of the solar wind is critically reviewed. In response to this review, the original suggestion is somewhat amplified and clarified by its author.
    Keywords: PROPULSION SYSTEMS
    Type: Science; 178; Dec. 8
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  • 14
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    Swirling base injection for supersonic combustion ramjets. (1972)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: Experimental evaluation of the swirling base injection proposed by Swithenbank and Chigier (1969) for application in supersonic combustion ramjets or scramjets. This concept of accelerated mixing in supersonic streams through swirl was tested, but the results indicate that swirl does not produce any enhancement of mixing.
    Keywords: PROPULSION SYSTEMS
    Type: AIAA Journal; 10; Sept
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  • 15
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    Fusion power for space propulsion. (1972)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: Principles of operation, interplanetary orbit-to-orbit mission capabilities, technical problems, and environmental safeguards are examined for thermonuclear fusion propulsion systems. Two systems examined include (1) a fusion-electric concept in which kinetic energy of charged particles from the plasma is converted into electric power (for accelerating the propellant in an electrostatic thrustor) by the van de Graaf generator principle and (2) the direct fusion rocket in which energetic plasma lost from the reactor has a suitable amount of added propellant to obtain the optimum exhaust velocity. The deuterium-tritium and the deuterium/helium-3 reactions are considered as suitable candidates, and attention is given to problems of cryogenic refrigeration systems, magnet shielding, and high-energy particle extraction and guidance.
    Keywords: PROPULSION SYSTEMS
    Type: New Scientist; 54; Apr. 20
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  • 16
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    Ion-thruster propellant utilization. (1972)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: Maximum propellant utilization in a mercury electron-bombardment thrustor is evaluated. The primary-electron region in the ion chamber of a bombardment thrustor is analyzed at maximum utilization. Both the analysis and experimental data from a range of ion-chamber configurations show a nearly constant loss rate for un-ionized propellant at maximum utilization over a wide range of total propellant flow rate. The discharge loss level of 1000 eV/ion was used to define maximum utilization. The exact level of this definition has no effect on the qualitative results and little effect on the quantitative results. The results obtained are particularly significant whenever efficient throttled operation is required.
    Keywords: PROPULSION SYSTEMS
    Type: Journal of Spacecraft and Rockets; 9; July 197
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  • 17
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    Comparison of advanced propulsion concepts for deep space exploration. (1972)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: Equations and charts are presented which permit rapid estimation of propulsion system performance requirements for some typical deep space missions. The simplicity results from use of gravity-free equations of motion, which are shown to yield good approximations to trip times obtained with solar gravity and planetary motion included. The agreement is satisfactory for missions that do not enter or depart from low orbits about the major planets. A number of advanced propulsion concepts for which performance estimates are available are compared with respect to their capability for fly-by, rendezvous, and round-trip planetary missions. Based on these estimates, the gas-core nuclear fission rocket and the pulsed fusion rocket yield the fastest trip times to the near planets. For round trips to Jupiter and beyond, the controlled fusion rocket shows progressively superior capabilities. Several propulsion concepts based on use of impinging laser beams are found to be noncompetitive with the other advanced concepts for deep space missions.
    Keywords: PROPULSION SYSTEMS
    Type: Journal of Spacecraft and Rockets; 9; Dec. 197
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  • 18
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    Advanced propulsion - Cleaner and quieter. (1972)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: Studies were conducted to determine the factors which are significant in advancing propulsion technology. The studies surveyed a wide distribution of variables including aircraft configuration, payload, range, and speed. System studies placed major emphasis on reducing noise and exhaust emissions while attaining good economies and performance. An engine for an advanced transport will probably superficially resemble the presently emerging generation of modern high-bypass and high-temperature turbofan engines, but would incorporate the advances in component and system technology identified by the propulsion system studies. These advances could be used to improve aircraft economics significantly with no increase in noise, or to significantly reduce noise and pollution with few or no economic penalties.
    Keywords: PROPULSION SYSTEMS
    Type: Astronautics and Aeronautics; 10; Aug. 197
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  • 19
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    Unmanned planetary spacecraft chemical rocket propulsion. (1972)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: Review of some chemical propulsion technology advances suitable for future unmanned spacecraft applications. Discussed system varieties include liquid space-storable propulsion systems, advanced liquid monopropellant systems, liquid systems for rendezvous and landing applications, and low-thrust high-performance solid-propellant systems, as well as hybrid space-storable systems. To optimize the performance and operational characteristics of an unmanned interplanetary spacecraft for a particular mission, and to achieve high cost effectiveness of the entire system, it is shown to be essential that the type of spacecraft propulsion system to be used matches, as closely as possible the various requirements and constraints. The systems discussed are deemed to be the most promising candidates for some of the anticipated interplanetary missions.
    Keywords: PROPULSION SYSTEMS
    Type: Journal of Spacecraft and Rockets; 9; Oct. 197
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  • 20
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    HEUS-RS applications study, volume 1 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Studies are given for sizing and integrating a high energy upper stage restartable solid motor into a flight stage with various payloads for use with Titan 3 and Thor launch vehicles. Motor and stage configurations are given along with performance evaluation of the HEUS-RS with the space shuttle.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-130274 , D2-116262-1-VOL-1
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  • 21
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    Apollo 15 mission report. Supplement 3: Ascent propulsion system final flight evaluation (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Results from the postflight analysis of the ascent propulsion system (APS) performance during the Apollo 15 mission are presented. The duty cycle for the LM-10 APS consisted of two firings, and ascent stage liftoff from the lunar surface and the terminal phase ignition (TPI) burn. An evaluation was made of APS performance for the first firing and found to be satisfactory. No propulsion data was received from the second APS burn; however, all indications were that the burn was nominal. All performance parameters were well within their LM-10 3-sigma limits. Calculated throat erosion at engine cutoff for the LM-10 APS was approximately 3 percent greater than predicted.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-68932 , TRW-20029-H062-R0-00 , MSC-05161-SUPPL-3
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  • 22
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    Study of solid rocket motor for space shuttle booster, Volume 3: Program acquisition planning (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The program planning acquisition functions for the development of the solid propellant rocket engine for the space shuttle booster is presented. The subjects discussed are: (1) program management, (2) contracts administration, (3) systems engineering, (4) configuration management, and (5) maintenance engineering. The plans for manufacturing, testing, and operations support are included.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-124238 , TWR-5672-VOL-3 , PUBL-0372-36176 , A10738
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  • 23
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    Apollo 14 mission report. Supplement 5: Descent propulsion system final flight evaluation (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The performance of the LM-8 descent propulsion system during the Apollo 14 mission was evaluated and found to be satisfactory. The average engine effective specific impulse was 0.1 second higher than predicted, but well within the predicted l sigma uncertainty. The engine performance corrected to standard inlet conditions for the FTP portion of the burn at 43 seconds after ignition was as follows: thrust, 9802, lbf; specific impulse, 304.1 sec; and propellant mixture ratio, 1603. These values are + or - 0.8, -0.06, and + or - 0.3 percent different respectively, from the values reported from engine acceptance tests and were within specification limits.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-69491 , MSC-04112-SUPPL-5 , TRW-17618-H219-R0-00-SUPPL-5
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  • 24
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    Study of solid rocket motors for a space shuttle booster. Volume 1: Executive summary (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An analysis of the solid propellant rocket engines for use with the space shuttle booster was conducted. A definition of the specific solid propellant rocket engine stage designs, development program requirements, production requirements, launch requirements, and cost data for each program phase were developed.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-124236 , TWR-5672-VOL-1 , PUBL-0372-36174 , A09995
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  • 25
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    Liquid propellant gas generators (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The design of gas generators intended to provide hot gases for turbine drive is discussed. Emphasis is placed on the design and operation of bipropellant gas generators because of their wider use. Problems and limitations involved in turbine operation due to temperature effects are analyzed. Methods of temperature control of gas turbines and combustion products are examined. Drawings of critical sections of gas turbines to show their operation and areas of stress are included.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-SP-8081
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  • 26
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    Rocket nozzle coolant channel thermal analysis program (E25107) (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A complete description of the liquid cooled rocket nozzle analysis program (E25107) is presented, including a users manual, program listing, and a sample problem. The program is recommended for use in designing liquid cooled rocket nozzles. In addition, it is adaptable to any system in which a liquid-cooled tubular structure is used to contain and direct the flow of a hot gas.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-132185 , N8110R:72-036
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  • 27
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    Study of solid rocket motor for space shuttle booster, volume 2, book 2 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A technical analysis of the solid propellant rocket engines for use with the space shuttle is presented. The subjects discussed are: (1) solid rocket motor stage recovery, (2) environmental effects, (3) man rating of the solid propellant rocket engines, (4) system safety analysis, (5) ground support equipment, and (6) transportation, assembly, and checkout.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-123926 , TWR-5672-VOL-2-BK-2
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  • 28
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    Study of solid rocket motor for space shuttle booster, volume 2, book 3, appendix A (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A systems requirements analysis for the solid propellant rocket engine to be used with the space shuttle was conducted. The systems analysis was developed to define the physical and functional requirements for the systems and subsystems. The operations analysis was performed to identify the requirements of the various launch operations, mission operations, ground operations, and logistic and flight support concepts.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-123728 , TWR-5672-VOL-2-BK-3-APP-A
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  • 29
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    Study of solid rocket motor for space shuttle booster, volume 2, book 5, appendices E thru H (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Preliminary parametric studies were performed to establish size, weight and packaging arrangements for aerodynamic decelerator devices that could be used for recovery of the expended solid propellant rocket motors used in the launch phase of the Space Shuttle System. Computations were made using standard engineering analysis techniques. Terminal stage parachutes were sized to provide equilibrium descent velocities for water entry that are presently thought to be acceptable without developing loads that could exceed the boosters structural integrity. The performance characteristics of the aerodynamic parachute decelerator devices considered are based on analysis and prior test results for similar configurations and are assumed to be maintained at the scale requirements of the present problem.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-123730 , TWR-5672-VOL-2-BK-5-APP-E-H
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  • 30
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    Study of solid rocket motor for space shuttle booster, volume 2, book 1 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The technical requirements for the solid propellant rocket engine to be used with the space shuttle orbiter are presented. The subjects discussed are: (1) propulsion system definition, (2) solid rocket engine stage design, (3) solid rocket engine stage recovery, (4) environmental effects, (5) manrating of the solid rocket engine stage, (6) system safety analysis, and (7) ground support equipment.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-123729 , PUBL-0372-36175 , TWR-5672-VOL-2-BK-1
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  • 31
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    Apollo 15 mission report. Supplement 3: Ascent propulsion system final flight evaluation (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The results are presented of the postflight analysis of the ascent propulsion system (APS) performance during the Apollo 15 Mission. The information presented includes: (1) calculated performance values for the APS lunar liftoff burn; (2) disucssion of analysis techniques, problems and assumptions; (3) comparison of postflight analysis and preflight prediction; (4) reaction control system (RCS) duty cycle included in the APS performance analysis; (5) transient performance analysis; and (6) the APS propellant consumption values.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-69204 , MSC-05161-SUPPL-3
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  • 32
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    Experimental quiet engine program aerodynamic performance of Fan C (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: This report presents the aerodynamic component test results of Fan C, a high-bypass-ratio, low-aerodynamic-loading, 1550 feet per second (472.4 m/sec), single-stage fan, which was designed and tested as part of the NASA Experimental Quiet Engine Program. The fan was designed to deliver a bypass pressure ratio of 1.60 with an adiabatic efficiency of 84.2 percent at a total fan flow of 915 lb/sec (415.0 kg/sec). It was tested with and without inlet distortion. A bypass total-pressure ratio of 1.61 and an adiabatic efficiency of 83.9 percent at a total fan flow of 921 lb/sec (417.8 kg/sec) were actually achieved. An operating margin in excess of 14.6 percent was demonstrated at design speed.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-120981
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    Strengthening of oxidation resistant materials for gas turbine applications (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Compressive surface layers were formed on hot-pressed silicon carbide and nitride. The objective of these treatments was to improve the impact resistance of these materials at 1590 K (2400 F). Quenching was used to form compressive surface layers on silicon carbide. The presence of the compressive stresses was demonstrated by slotted rod tests. Compressive stresses were retained at elevated temperatures. Improvements in impact resistance at 1590 K (2400 F) and flexural strength at room temperature were achieved using cylindrical rods 3.3 mm (0.13 in.) in diameter. Carburizing treatments were used to form the surface layers on silicon nitride. In a few cases using rectangular bars improvements in impact resistance at 1590 K (2400 F) were observed.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121002
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    Development of a multikilowatt ion thruster power processor (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A feasibility study was made of the application of silicon-controlled, rectifier series, resonant inverter, power conditioning technology to electric propulsion power processing operating from a 200 to 400 Vdc solar array bus. A power system block diagram was generated to meet the electrical requirements of a 20 CM hollow cathode, mercury bombardment, ion engine. The SCR series resonant inverter was developed as a primary means of power switching and conversion, and the analog signal-to-discrete-time-interval converter control system was applied to achieve good regulation. A complete breadboard was designed, fabricated, and tested with a resistive load bank, and critical power processor areas relating to efficiency, weight, and part count were identified.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-120928
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    Technology development of a biowaste resistojet, volume 1 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The materials research effort conducted in support of a NASA-sponsored biowaste resistojet development program is summarized. The resistojet concept under development is the concentric tube design wherein the final pass of the gases through the thruster is through the resistance heated center tube. To produce high specific impulses, this center tube must operate at very high temperatures and it is this element that is most critical in the design. Because of the corrosive nature of the biowaste gases at high temperature, and because of the limited data available for many potential materials, the subject materials study was conducted.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-112149
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    Experimental investigation of combustor effects on rocket thrust chamber performance (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A design and experimental program to develop special instrumentation systems, design engine hardware, and conduct tests using LOX/GH2 propellants in which the propellant flow stratification was controlled is described. The mixture ratio was varied from 4.6 to 6 overall. The mixture ratios in the core and outer zone were varied from 3.5 to 6 and 5 to 8, respectively. The range in boundary layer coolant was from 0 to 10 percent of the fuel. The nominal chamber pressure and thrust were 225 psia and 7000 pounds, respectively. Pressure and heat flux profiles as well as gas sampling of the exhaust products were obtained. Specific impulse efficiencies of approximately 94 percent and characteristic velocity efficiencies of approximately 97 percent were obtained during the experiments.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-128318 , R-8903
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    Filament-wound graphite/epoxy rocket motor case (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The fabrication procedures are described for a filament-wound rocket motor case, approximately 56 cm long x 71 cm diameter, utilizing high tensile strength graphite fibers. The process utilized Fiberite Hy-E-1330B prepreg tape which consists of Courtaulds HTS fibers in a temperature-sensitive epoxy matrix. This fabrication effort, with resultant design, material and process recommendations, substantiates the manufacturing feasibility of graphite/epoxy rocket motor cases in the 56 cm x 71 cm size range.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-128417 , BC-8845-FAB
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    A concept for jet noise suppression for an afterburning turbojet engine (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A conceptual design of an afterburner system for turbojet engines which may reduce the jet exhaust noise by approximately 10 decibels is presented in this report. The proposed system consists of an array of swirl-can combustors and jet dividing nozzle tubes. The nozzle tubes translate axially upstream of the swirl cans when not in use. Results of preliminary design calculations and photographs of a kinematic model as applied to a hypothetical turbojet engine are presented.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-68144 , E-7167
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  • 39
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    Flight investigation of an air-cooled plug nozzle with afterburning turbojet (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A convectively cooled plug nozzle, using 4 percent of the engine air as the coolant, was tested in 1967 K (3540 R) temperature exhaust gas. No significant differences in cooling characteristics existed between flight and static results. At flight speeds above Mach 1.1, nozzle performance was improved by extending the outer shroud. Increasing engine power improved nozzle efficiency considerably more at Mach 1.2 than at 0.9. The effect of nozzle pressure ratio and secondary weight flow on nozzle performance are also presented.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2607 , E-6676
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  • 40
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    Experimental dynamic response of a two-dimensional, Mach 2.7, mixed compression inlet (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A test program was conducted on a two-dimensional supersonic inlet. Internal disturbances in diffuser exit mass flow were produced by oscillating overboard bypass doors. Open-loop dynamic responses of shock position, throat exit and diffuser exit static pressures are presented. The steady-state and dynamic coupling between ducts were also obtained. The experimental results from the two-dimensional inlet are compared to results from a similar size axisymmetric inlet and also to a transfer function synthesis program.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-6957 , E-7002
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  • 41
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    Cold-air investigation of a turbine for high-temperature-engine application. 4: Two-stage turbine performance (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The efficiency of a two-stage turbine is discussed. The turbine efficiency was 0.932 for equivalent design operating conditions (speed) and specific work, which compares closely to the value of 0.929 that would be estimated using the first-stage efficiency. The mass flow obtained with the two-stage configuration indicated that the mass flow characteristics of the two stages were closely matched at design operating conditions. The stage work split at these conditions was 0.505-0.495, which was close to the design work split of 0.515-0.485.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-6960 , E-6912
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  • 42
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    High temperature compounds for turbine vanes (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Fabrication and microstructure control studies were conducted on SiC, Si3N4, and composites based on these compounds. Charpy mode impact testing to 2400 F established that beta-spodumene, lithium aluminum silicate, coated Si3N4, Si3N4 derived from alpha-Si3N4 powder, and SiC containing 5-25 v/o chopped C fibers had the most promising strengths. Several other composite systems had excellent microstructures and could prove interesting materials in the future. Stress-rupture testing on Si3N4 established that increasing 2000 F - 100 hour strengths were obtained for increasing grain size to at least 5 micrometers, increasing density and possibly increasing phase purity. These parameters became less important at 2400 F where it is thought a grain boundary phase controls strength.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-120966 , AVSD-0336-72-CR
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    Comparison of advanced propulsion concepts for deep space exploration (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Equations and charts are presented that permit rapid estimation of propulsion-system performance requirements for some typical deep-space missions. A number of advanced propulsion concepts for which performance estimates are available are compared with respect to their capability for flyby, rendezvous, and roundtrip planetary missions. Based on these estimates, the gas-core nuclear fission rocket and the pulsed fusion rocket yield the fastest trip times to the near planets. For round trips to Jupiter and beyond, the controlled fusion rocket shows progressively superior capabilities. Several propulsion concepts based on use of impinging laser beams are found to be noncompetitive with the other advanced concepts for deep space missions. Requirements for attainment of interstellar distances within a human lifetime are found to be some orders of magnitude beyond the capabilities of any propulsion concepts for which performance estimates are now possible.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-6968 , E-6798
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    A graphite-lined regeneratively cooled thrust chamber (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Design concepts, based on use of graphite as a thermal barrier for regeneratively cooled FLOX-methane thrust chambers, have been screened and concepts selected for detailed thermodynamic, stress, and fabrication analyses. A single design employing AGCarb-101, a fibrous graphite composite material, for a thermal barrier liner and an electroformed nickel structure with integral coolant passages was selected for fabrication and testing. The fabrication processes and the test results are described and illustrated.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-120853
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    Factors affecting altitude relight performance of a double-annular ram-induction combustor (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A test program was conducted to evaluate the altitude relight capabilities of a short-length, double-annular, ram-induction combustor which was designed for Mach 3 cruise operation. The use of distorted inlet-air flow profiles was tried to evaluate their effect on the relight performance. No significant improvement in altitude relight performance was obtained with this approach. A study was also made to determine the effects of the reference Mach number, the fuel temperature, and the fuel volatility (ASTM-A1 against JP-4) on the altitude relight performance. Decreasing the reference Mach number, increasing the fuel temperature, and using more volatile fuel all decrease the combustor pressure necessary for relight.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2630 , E-6788
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  • 46
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    Sea-level evaluation of digitally implemented turbojet engine control functions (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The standard hydromechanical control system of a turbojet engine was replaced with a digital control system that implemented the same control laws. A detailed discussion of the digital control system in use with the engine is presented. The engine was operated in a sea-level test stand. The effects of control update interval are defined, and a method for extending this interval by using digital compensation is discussed.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-6936 , E-6977
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    Propulsion technology for an advanced subsonic transport (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Engine design studies for future subsonic commercial transport aircraft were conducted in parallel with airframe studies. These studies surveyed a broad distribution of design variables, including aircraft configuration, payload, range, and speed, with particular emphasis on reducing noise and exhaust emissions without severe economic and performance penalties. The results indicated that an engine for an advanced transport would be similar to the currently emerging turbofan engines. Application of current technology in the areas of noise suppression and combustors imposed severe performance and economic penalties.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2625 , E-6913
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    High-loading, 1800 ft/sec tip speed transonic compressor fan stage. 1: Aerodynamic and mechanical design (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A single stage fan with a tip speed of 1800 ft/sec (548.6m/sec) and hub/tip ratio of 0.5 was designed to produce a pressure ratio of 2.285:1 with an adiabatic efficiency of 84.0%. The design flow per inlet annulus area is 38.7 lbm/sq ft-sec (188.9KG/sqm-sec). Rotor blades have modified multiple-circular-arc and precompression airfoil sections. The stator vanes have multiple-circular-arc airfoil sections.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-120907 , PWA-4534
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    Wall temperature distribution calculation for a rocket nozzle contour (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The JANNAF turbulent boundary layer (TBL) computer program, applicable to rocket nozzles, requires a wall temperature distribution among other input parameters to determine boundary layer behavior, heat transfer, and performance degradation. The inclusion of a complete regenerative cooling cycle model with associate geometry, material and fluid property data provides a capability to internally calculate wall temperature profiles on the hot gas and coolant flow-side, as well as the coolant flow bulk temperature variation. Besides the regular heat transfer and performance degradation calculations, the new concept can be used to optimize the cooling cycle, flow requirements, and cooling jacket geometry.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-6825
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    Stability evaluation of a rocket engine for gaseous oxygen difluoride (OF2) and gaseous diborane (B2H6) propellants (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Results of an experimental evaluation of the dynamic stability of a candidate combustor for the space storable propellants gaseous OF2/B2H6 show that the combustor is unstable without supplementary damping. A computer analysis indicated that the uninhibited engine could be unstable. The experiments, conducted with O2/C2H4 substitute propellants and with 70-30 FLOX/B2H6 (OF2 simulated with FLOX), show that the uninhibited combustor has a low stability margin to starting transient perturbations, but that is relatively insensitive to bomb disturbances. Damping cavities are shown to provide stability.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-127859 , JPL-TR-32-1561
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    Experimental quiet engine program aerodynamic performance of Fan B (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: This report presents the aerodynamic component test results of Fan B, one of two high-bypass-ratio, 1160 feet per second (353.6 m/sec) single-stage fans, which was designed and tested as part of the NASA Experimental Quiet Engine Program. The fan was designed to deliver a bypass pressure ratio of 1.50 with an adiabatic efficiency of 87.0% at a total fan flow of 950 lb/sec (430.9 kg/sec). It was tested with and without inlet distortion. A bypass total pressure ratio of 1.52 and an adiabatic efficiency of 86.9% at a total fan flow of 966 lb/sec (438.2 kg/sec) were actually achieved. An operating margin of 19.5% was demonstrated at design speed.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-72993
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    Experimental findings from zero-tank net positive suction head operation of the J-2 hydrogen pump (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The results of a series of liquid hydrogen turbopump tests to demonstrate the feasibility of zero-tank net positive suction head are presented. A J-2 engine hydrogen pump and S-IVB stage fuel feed system were used for this investigation. The pump was operated at flows and speeds equivalent to normal J-2 engine operating conditions and at hydrogen bulk temperatures between 39 and 45 R. These tests show zero-tank not positive suction head to be a realistic operating mode that should be considered for future applications.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-6824 , M-368
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    Highly loaded multi-stage fan drive turbine-tandem blade configuration design (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The results of the tandem blade configuration design study are reported. The three stage constant-inside-diameter turbine utilizes tandem blading in the stage two and stage three vanes and in the stage three blades. All other bladerows use plain blades. Blading detailed design is discussed, and design data are summarized. Steady-state stresses and vibratory behavior are discussed, and the results of the mechanical design analysis are presented.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-2097
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    Development and testing of the propulsion subsystem for the Mariner Mars 1971 spacecraft (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The design, testing, fabrication, and problems associated with the development of the Mariner 9 propulsion system are described. Also covered are the design and operation of the associated ground support equipment used to test and service the propulsion system.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-127751 , JPL-TM-33-552
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  • 55
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    Digital-computer normal shock position and restart control of a Mach 2.5 axisymmetric mixed-compression inlet (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Digital computer control of a mixed-compression inlet is discussed. The inlet was terminated with a choked orifice at the compressor face station to dynamically simulate a turbojet engine. Inlet diffuser exit airflow disturbances were used. A digital version of a previously tested analog control system was used for both normal shock and restart control. Digital computer algorithms were derived using z-transform and finite difference methods. Using a sample rate of 1000 samples per second, the digital normal shock and restart controls essentially duplicated the inlet analog computer control results. At a sample rate of 100 samples per second, the control system performed adequately but was less stable.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-6880 , E-6498
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  • 56
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    Single-stage experimental evaluation of tandem-airfoil rotor and stator blading for compressors. Part 3: Data and performance for stage C (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Stage C, comprised of tandem-airfoil rotor C and tandem-airfoil stator B, was designed and tested to establish performance data for comparison with the performance of conventional single-airfoil blading. Velocity diagrams and blade leading and trailing edge metal angles selected for the conventional rotor and stator blading were used in the design of the tandem blading. The rotor had an inlet hub/tip ratio of 0.8 and a design tip velocity of 757 ft/sec. At design equivalent rotor speed, rotor C achieved a maximum adiabatic efficiency of 91.8% at a pressure ratio of 1.31. The stage maximum adiabatic efficiency was 86.5% at a pressure ratio of 1.31.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-120938 , PWA-FR-5028
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  • 57
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    Highly loaded multi-stage fan drive turbine: Leaned stator configuration design (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The results of the high lift blade configuration design study are reported. The three-stage constant-inside-diameter turbine utilizes a ten degree tangentially leaned stator in stage three. All other bladerows use plain blades. Analysis of the leaned stator is discussed, and detailed design data are summarized. Steady-state stresses are discussed, and the results of the mechanical design analysis are presented.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-2096 , R71AEG309
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    A 156-in.-dia SRM mass properties report (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Mass property data for the 156-in.-dia SRM booster with two segments are presented. The SRM baseline booster has a fixed 15 deg canted nozzle and no thrust neutralization system. Summary mass property data for alternative booster configurations are included.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-123727 , REPT-1917-MP1
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    LF467 concept definition (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The LF467 concept is discussed. The LF467 is an advanced turbotip lift fan intended for application with the YJ97-GE-100 turbojet gas generator on a V/STOL transport research aircraft. The program objective was to define a fan that develops a reasonably high (1.30) fan pressure ratio consistent with propulsion requirements of a V/STOL research transport aircraft and that exhibits the ability to achieve a 100 PNdE overall noise objective through the use of modest additional installation treatment. The aerodynamic and mechanical designs of this system and the resulting configuration, weight, and noise predictions are presented.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-120909 , R72AEG207
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  • 60
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    Drive turbine systems for 20-inch turbofan simulators. 1: Duct turbine design (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A study was made to evolve the turbine drive systems for 20-inch turbofan engine simulators. The fan designs used in the simulators included single-stage and two-stage configurations that covered a wide range of rotative speed and power requirement. The objective assumed for the study was to evolve one core turbine design that could drive all of the single-stage fans and, when operated in combination with one duct turbine design, drive all of the two-stage fans. The duct turbine power output is then needed to determine the make-up power required of the core turbine over the range of two-stage fan operating conditions. The duct turbine design analysis is reported and includes the selection of the duct turbine velocity diagram, a description of the blade design, and a determination of its off-design performance. Adjustable stators were found to be quite advantageous to the duct turbine off-design operation. The use of adjustable stators enabled the duct turbine to accommodate fan mass flow at all operating points and caused the duct turbine power output to increase as the total power requirement increased. This in turn resulted in a core turbine make-up power requirement that was not significantly greater than that required for driving the single-stage fans.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-68081 , E-6972
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    Pulse performance analysis for small hypergolic-propellant rocket engines (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Small rocket engine tests were conducted for the purpose of obtaining pulse performance data to aid in preliminary design and evaluation of attitude control systems. Both monopropellant and hypergolic bipropellant engines of thrust levels from 1 to 100 lbs were tested. The performance data for the hypergolic propellant rockets are compared with theoretical performance calculated from idealized chamber filling and evacuation characteristics. Electromechanical delays in valve response and heat transfer characteristics were found to cause substantial deviation between theoretical and test performance. The theoretical analysis is modified to obtain a semi-empirical model for hypergolic propellant rockets which is demonstrated to be reasonably accurate for two different engine configurations over a considerable range of duty cycles.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-64673
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    Space tug propulsion system failure mode, effects and criticality analysis (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: For purposes of the study, the propulsion system was considered as consisting of the following: (1) main engine system, (2) auxiliary propulsion system, (3) pneumatic system, (4) hydrogen feed, fill, drain and vent system, (5) oxygen feed, fill, drain and vent system, and (6) helium reentry purge system. Each component was critically examined to identify possible failure modes and the subsequent effect on mission success. Each space tug mission consists of three phases: launch to separation from shuttle, separation to redocking, and redocking to landing. The analysis considered the results of failure of a component during each phase of the mission. After the failure modes of each component were tabulated, those components whose failure would result in possible or certain loss of mission or inability to return the Tug to ground were identified as critical components and a criticality number determined for each. The criticality number of a component denotes the number of mission failures in one million missions due to the loss of that component. A total of 68 components were identified as critical with criticality numbers ranging from 1 to 2990.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-61388
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    Thrust measurements of a hollow-cathode discharge (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Thrust measurements of a hollow cathode mercury discharge were made with a synthetic mica target on a torsion pendulum. Thrust measurements were made for various target angles, tip temperatures, flow rates, keeper discharge powers, and accelerator electrode voltages. The experimental thrust data are compared with theoretical values for the case where no discharge power was employed.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-6705 , E-6691
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    Preliminary performance of a low loading high tip speed fan stage (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Limited data indicates the feasibility of designing fan stages for operation with weak oblique shocks in the rotor blade tip region. A 1600 ft/sec rotor tip speed fan indicated an overall efficiency of 0.846 at a pressure ratio of 1.51 even though the rotor blades had incurred damage in the leading edge tip region prior to obtaining the design speed performance. After the test it was observed that a number of the rotor blade mid-span dampers had failed, subsequently damaging a number of rotor blades in the leading edge tip region.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-68027 , E-6826
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    Analysis of a supersonic-combustion rocket concept (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A preliminary analysis has been made of a supersonic-combustion rocket engine concept using hydrogen and oxygen propellants. The ejector action of a separate small rocket motor is employed to pump the propellants to high stagnation pressures and supersonic velocities. Therefore complicated heavy turbopumps are eliminated and cooling problems of a sonic throat are reduced. The results of the study show that vacuum specific impulse levels as high as a conventional rocket having the same chamber pressure as the drive motor are possible. The supersonic-combustion rocket would be an attractive alternate for a high-altitude low-thrust conventional rocket operating with a pressure feed propellant system. It would also be a convenient technique for obtaining extremely high thrusts without the need for developing corresponding large turbopumps.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-68020 , E-6825
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    Precomputed state dependent digital control of a nuclear rocket engine (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A control method applicable to multiple-input multiple-output nonlinear time-invariant systems in which desired behavior can be expressed explicitly as a trajectory in system state space is developed. The precomputed state dependent control method is basically a synthesis technique in which a suboptimal control law is developed off-line, prior to system operation. This law is obtained by conducting searches at a finite number of points in state space, in the vicinity of some desired trajectory, to obtain a set of constant control vectors which tend to return the system to the desired trajectory. These vectors are used to evaluate the unknown coefficients in a control law having an assumed hyperellipsoidal form. The resulting coefficients constitute the heart of the controller and are used in the on-line computation of control vectors. Two examples of PSDC are given prior to the more detailed description of the NERVA control system development.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-126649
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  • 67
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    Effects of high combustion chamber pressure on rocket noise environment (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The acoustical environment for a high combustion chamber pressure engine was examined in detail, using both conventional and advanced theoretical analysis. The influence of elevated chamber pressure on the rocket noise environment was established, based on increase in exit velocity and flame temperature, and changes in basic engine dimensions. Compared to large rocket engines, the overall sound power level is found to be 1.5 dB higher, if the thrust is the same. The peak Strouhal number shifted about one octave lower to a value near 0.01. Data on apparent sound source location and directivity patterns are also presented.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-126548 , WR-75-6
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  • 68
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    Upper-stage space shuttle propulsion by means of separate scramjet and rocket engines (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A preliminary mission study of a reusable vehicle from staging to orbit indicates payload advantages for a dual-propulsion system consisting of separate scramjet and rocket engines. In the analysis the scramjet operated continuously and the initiation of rocket operation was varied. For a stage weight of 500,000 lb the payload was 10.4 percent of stage weight or 70 percent greater than that of a comparable all-rocket-powered stage. When compared with a reusable two-state rocket vehicle having 50,000 lb payload, the use of the dual propulsion system for the second stage resulted in significant decreases in lift-off weight and empty weight, indicating possible lower hardware costs.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-6762 , E-6555
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  • 69
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    Preliminary experiments on the noise generated by target-type thrust reverser models (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Experiments are reported on the noise generated by model V-gutter and semicylindrical target-type reversers with circular nozzles. Nozzles were 5.24 and 7.78 cm in diameter. Nozzle pressure ratio ranged from 1.25 to 1.72. The spacing between reversers and nozzle, as well as the reverser orientation, was also varied. More noise was generated with reversers than with the nozzle alone. The measured maximum overall sound pressure level varied with the sixth power of the nozzle exit velocity. Noise levels were more uniform in regard to directivity with reversers than with the nozzle alone. It is concluded that thrust reversers, can be a significant noise problem, especially for STOL aircraft using thrust reversers during approach.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2553 , E-6707
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  • 70
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    Static performance of a 13.97 cm (5.5 inch) diameter model VTOL lift fan (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A tip-turbine-driven fan of the type currently being used in wind tunnel tests of VTOL lift fan models was tested. Values of thrust, weight flow, exit total and static pressure, exit swirl angle, and turbine temperature drop were measured as a function of fan speed for several inlet and exit configurations. A standard fan performance map was also obtained.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-2051 , HST-TR-329-0-1
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  • 71
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    Dynamics of a supersonic inlet-engine combination subjected to disturbances in fuel flow and inlet overboard bypass airflow (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An axisymmetric mixed-compression supersonic inlet and a single-spool turbojet engine were dynamically tested at Mach 2.5. The propulsion system was subjected to sweep-frequency sinusoidal disturbances of either inlet overboard bypass airflow. The disturbances were at a logarithmic sweep rate of 1 decade per minute. Dynamic responses were taken of signals throughout the propulsion system. Selected signals were reduced relative to the prime propulsion system parameters. The experimental data are presented in Bode plots. Most of the plots are for a frequency range of 1.0 to 50 hertz.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2558 , E-6115
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  • 72
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    Comparison of the effect of two damper sizes on the performance of a low-solidity axial-flow transonic compressor rotor (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The experimental performance of a 20-inch-diameter axial-flow transonic compressor rotor with small dampers is presented. The compressor rotor was tested earlier with large dampers which were twice in size, and comparisons of overall performance and radial distributions of selected flow and performance parameters are made. The rotor with small dampers experienced lower losses in the damper region which resulted in locally higher values of temperature rise efficiency and total pressure ratio. However, there was no appreciable effect on overall efficiency and pressure ratio. A greater stall margin was measured for the rotor with small dampers at design speed, but at 70 and 90 percent of design speed the rotor with large dampers had somewhat greater flow range.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2536 , E-6536
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    Design and test of porous-tungsten mercury vaporizers (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: New design requirements for porous plug-type vaporizers used with Kaufman thrusters and thruster arrays are discussed. The results of testing samples of porous tungsten for mercury flow rate, liquid intrusion pressure level, and mechanical strength are presented. Nitrogen gas was used instead of mercury vapor for approximate calibration. Liquid intrusion pressure levels will require that flight thruster systems with long feedlines have restrictions in the dynamic line during launch.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-6782 , E-6792
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  • 74
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    Single-stage evaluation of highly-loaded high-Mach-number compressor stages 5. Data and performance of baseline, corner-blow wall suction and combined corner blow wall suction stator (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A single-stage compressor with a rotor tip speed of 1600 ft/sec and a 0.5 hub tip ratio was used to investigate the effects of several stator endwall treatment methods on stage range and performance. These endwall treatment methods consisted of stator corner-blow, annular wall suction upstream of stator leading edge, and combined corner-blow and annular wall suction. The overall stage performance with corner blow was essentially the same as the baseline performance. The performance for the annular wall suction and the combined corner-blow and wall suction showed a reduction in peak efficiency of 2.5 percentage points compared to the baseline data.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-120887 , PWA-4312
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    Study of solid rocket motors for a space shuttle booster. Volume 2: Technical report (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: For abstract, see N72-22784.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-123628 , UTC-4205-72-7-VOL-2
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  • 76
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    Mass properties report on 156-inch diameter SRM (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Mass properties data for 156-in.-dia SRM booster with two segments are presented. The SRM baseline booster has a fixed 15 degrees canted nozzle and no thrust neutralization system. Summary mass properties data for alternative booster configurations are included.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-123613 , REPT-1917-MP1
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  • 77
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    Study of solid rocket motors for a space shuttle booster. Volume 1: Executive summary (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The design, development, production, and launch support analysis for determining the solid propellant rocket engine to be used with the space shuttle are discussed. Specific program objectives considered were: (1) definition of engine designs to satisfy the performance and configuration requirements of the various vehicle/booster concepts, (2) definition of requirements to produce booster stages at rates of 60, 40, 20, and 10 launches per year in a man-rated system, and (3) estimation of costs for the defined SRM booster stages.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-123623 , UTC-4205-72-7-VOL-1
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  • 78
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    Feasibility study of a pressure-fed engine for a water recoverable space shuttle booster (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: For abstract, see N72-22781.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-123605 , SE-019-013-2H , DRD-SE-02
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  • 79
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    Feasibility Study of a Pressure-fed Engine for a Water Recoverable Space Shuttle Booster (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Detailed mass properties are presented for a gimbaled, fixed thrust, regeneratively cooled engine having a coaxial pintle injector. The baseline design parameters for this engine are tabulated. Mass properties are also summarized for several other engine configurations i.e., a hinge nozzle using a Techroll seal, a gimbaled duct cooled engine and a regeneratively cooled engine using liquid injection thrust vector control (LITVC). Detailed engine analysis and design trade studies leading to the selection of a regeneratively cooled gimbaled engine and pertaining to the selection of the baseline design configuration are also given.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-123604 , SE-019-015-2H , DRD-SE-03
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  • 80
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    Monopropellant hydrazine resistojet model specification (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The specifications for the electrothermal hydrazine thruster model are presented including performance, design and qualification requirements, and product configuration and acceptance tests. The contamination control procedures, acceptance test plan, and engineering drawings are included.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-122377 , TRW-20266-6010-R0-00
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  • 81
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    Performance and control study of a low-pressure-ratio turbojet engine for a drone aircraft (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The results of analog and digital computer studies of a low-pressure-ratio turbojet engine system for use in a drone vehicle are presented. The turbojet engine consists of a four-stage axial compressor, single-stage turbine, and a fixed area exhaust nozzle. Three simplified fuel schedules and a generalized parameter fuel control for the engine system are presented and evaluated. The evaluation is based on the performance of each schedule or control during engine acceleration from a windmill start at Mach 0.8 and 6100 meters to 100 percent corrected speed. It was found that, because of the higher acceleration margin permitted by the control, the generalized parameter control exhibited the best dynamic performance.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2537 , E-6688
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  • 82
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    Two-Stage Fan I: Aerodynamic and Mechanical Design (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A two-stage, highly-loaded fan was designed to deliver an overall pressure ratio of 2.8 with an adiabatic efficiency of 83.9 percent. At the first rotor inlet, design flow per unit annulus area is 42 lbm/sec/sq ft (205 kg/sec/sq m), hub/tip ratio is 0.4 with a tip diameter of 31 inches (0.787 m), and design tip speed is 1450 ft/sec (441.96 m/sec). Other features include use of multiple-circular-arc airfoils, resettable stators, and split casings over the rotor tip sections for casing treatment tests.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-120859 , PWA-4148 , HQ-E-DAA-TN38870
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  • 83
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    Noise generated by quiet engine fans. 1: FanB (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Acoustical tests of full scale fans for jet engines are presented. The fans are described and some aerodynamic operating data are given. Far field noise around the fan was measured for a variety of configurations over a range of operating conditions. Complete results of one third octave band analysis are presented in tabular form. Power spectra and sideline perceived noise levels are included.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-2528 , E-6652
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  • 84
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    Effects of radial and circumferential inlet velocity profile distortions on performance of a short-length double-annular ram induction combustor (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Inlet air velocity profile tests were conducted on a full-scale short-length 102-centimeter-diameter annual combustor designed for advanced gas turbine engine applications. The inlet profiles studied include radial distortions that were center peaked, and tip peaked, as well as a circumferential distortion which was center peaked for one-third of the circumference and flat for the other two-thirds. An increase in combustor pressure loss was the most significant effect of the radial air velocity distortions. With the circumferential distortion, exit temperature pattern factor doubled when compared to a flat velocity profile.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-6706 , E-6464
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  • 85
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    Propulsion system studies for an advanced high subsonic, long range jet commercial transport aircraft (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Propulsion system characteristics for a long range, high subsonic (Mach 0.90 - 0.98), jet commercial transport aircraft are studied to identify the most desirable cycle and engine configuration and to assess the payoff of advanced engine technologies applicable to the time frame of the late 1970s to the mid 1980s. An engine parametric study phase examines major cycle trends on the basis of aircraft economics. This is followed by the preliminary design of two advanced mixed exhaust turbofan engines pointed at two different technology levels (1970 and 1985 commercial certification for engines No. 1 and No. 2, respectively). The economic penalties of environmental constraints - noise and exhaust emissions - are assessed. The highest specific thrust engine (lowest bypass ratio for a given core technology) achievable with a single-stage fan yields the best economics for a Mach 0.95 - 0.98 aircraft and can meet the noise objectives specified, but with significant economic penalties. Advanced technologies which would allow high temperature and cycle pressure ratios to be used effectively are shown to provide significant improvement in mission performance which can partially offset the economic penalties incurred to meet lower noise goals. Advanced technology needs are identified; and, in particular, the initiation of an integrated fan and inlet aero/acoustic program is recommended.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-121016 , R72AEG296
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  • 86
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    Study to adapt solar electric propulsion to the Pioneer F and G spacecraft (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The addition of an electric thrust subsystem to the spin-stabilized Pioneer F and G spacecraft to improve performance capability for certain missions is discussed. The evaluation was performed for the Atlas and Titan launch vehicles with Centaur and TE-364-4 stages and for electric thrust stages of 8- and 5-kw with three 30- and five 15-cm thrusters respectively. The combination of a spinning spacecraft with electric propulsion is a concept only recently evaluated and the penalty from spinning over three-axis stabilized is not as significant as might initally be thought. There are major gains in weight, cost, and reliability, the disadvantages being lower data rate during the thrust phase and less efficient pointing. A variety of missions were evaluated from a solar approach mission into 0.14 AU to a flyby mission of Neptune at approximately 30 AU. Performance improvements were present for all missions evaluated.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-114514 , TRW-22125-6001-R0-00
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  • 87
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    Development and in-flight performance of the Mariner 9 spacecraft propulsion system (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: On November 14, 1971, Mariner 9 was decelerated into orbit about Mars by a 1334-newton (300-lbf) liquid bipropellant propulsion system. The development and in-flight performance are described and summarized of this pressure-fed, nitrogen tetroxide/monomethyl hydrazine bipropellant system. The design of all Mariner propulsion subsystems has been predicated upon the premise that simplicity of approach, coupled with thorough qualification and margin-limits testing, is the key to cost-effective reliability. The qualification test program and analytical modeling of the Mariner 9 subsystem are discussed. Since the propulsion subsystem is modular in nature, it was completely checked, serviced, and tested independent of the spacecraft. Proper prediction of in-flight performance required the development of three significant modeling tools to predict and account for nitrogen saturation of the propellant during the six-month coast period and to predict and statistically analyze in-flight data. The flight performance of the subsystem was excellent, as were the performance prediction correlations. These correlations are presented.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-129097 , JPL-TM-33-574
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  • 88
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    Atomization and mixing characteristics of gas/liquid coaxial injector elements. (1972)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: (For abstract see issue 14, page 2291, Accession no. A71-30736)
    Keywords: PROPULSION SYSTEMS
    Type: AIAA PAPER 71-672
    Format: text
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  • 89
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    Numerical solutions of reactive fluid flows during postignition transients in hybrid rocket systems. (1972)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: A computational method has been developed for the study of the post-ignition transients in hybrid rocket systems. The particular system chosen consisted of a gaseous oxidizer flowing within a tube of solid fuel, resulting in heterogeneous combustion. With the appropriate assumptions, two-dimensional, time-dependent conservation equations were derived for the reacting gas phase, and for the solid phase, in a cylindrical coordinate system. These were then programmed for numerical computation, using two implicit finite-difference schemes, the Lax-Wendroff scheme for the gas phase, and the Crank-Nicolson scheme for the solid phase. Appropriate initial and boundary conditions were represented, including heat and mass conservation at the interface between gas and solid. Initially, no attempt was made to relate the recession rate at the surface to the surface temperature, or to include heat transfer by radiation. A simple case was selected for preliminary calculations, with aluminum and oxygen as fuel and oxidizer, and aluminum oxide as the product.
    Keywords: PROPULSION SYSTEMS
    Type: Journal of Computational Physics; 9; Apr. 197
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  • 90
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    Technology development and demonstration of a low thrust resistojet thruster (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Three thrusters were fabricated to definitized thruster drawings using new rhenium vapor deposition technology. Two of the thrusters were operated using ammonia as propellant and one was operated using hydrogen propellant for performance determination. All demonstrated consistent operational specific impulse performance while demonstrating thermal performance better than the development units from which they evolved. Two of the thrusters were subjected to environmental structural testing including vibration, acceleration and shock loading to specifications. Both of the thrusters subjected to the environmental tests passed all required tests. The third, spare, thruster was introduced into the life test portion of the program. Two thrusters were then subjected to a life cycling test program under typical spacecraft operating power levels. During the life test sequence, the hydrogen thruster accrued 720 operating life test cycles, more than 370 on-off cycles and 365 hours of powered up time. The ammonia accrued approximately 380 on-off cycles and 392.2 on time hours of operation during the 720 cycling hour test. Both thrusters completed the scheduled operational life test in reasonably good condition, structurally integral and capable of indefinite further operation.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-112160
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  • 91
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    Effect of jet convergence angle on the performance of annular nozzles with semitoroidal concave plugs at Mach numbers up to 1.82 (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Results of an investigation at static conditions and at Mach numbers up to 1.82 are presented for 12 nozzle configurations which have jet convergence angle and jet throat area as geometric parameters. The variation of jet convergence angle from 15 to 40 deg had little effect on the performance of the nozzles having the large value of primary throat area; however, increasing jet convergence angle generally had an adverse effect on performance of the nozzles having the smaller value of primary throat area. The performance of the nozzle configurations with the larger primary throat area is competitive with nozzles designed for operation over the Mach number range.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-6897 , L-8440
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  • 92
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    TF-34 turbofan quiet engine study (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A study of high bypass turbofan engines in heavily sound-suppressed nacelles based on the TF-34 engine. The four-engine noise objective was 95 PNdb at four locations typical of takeoff and landing. Three engines were studied; these had fan pressure ratios, bypass ratios, and fan tip speeds respectively of 1.48/6.5/404 m/sec (1327 ft/sec), 1.25/13/305 (1000), 1.25/13/366(1200). The bypass 13 engines had a variable pitch fan, direct- and gear-driven. Noise suppressive treatment was identified which met 95 PNdb objective except for sideline liftoff at 6.5 bypass, full power, which was 2 PNdb noisier; at 90% power, 95 PNdb was achieved.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-120914
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  • 93
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    A single axis electrostatic beam deflection system for a 5-cm diameter ion thruster (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A single-axis electrostatic beam deflection system has been tested on a 5-cm diameter mercury ion thruster at a thrust level of about 0.43 mlb (25 mA beam current at 1400 volts). The accelerator voltage was 500 volts. Beam deflection capability of plus or minus 10 deg was demonstrated. A life test of 1367 hours was run at the above conditions. Results of the test indicated that the system could possibly perform for upwards of 10,000 hours.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-68133
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    Effective specific impulse of external nuclear pulse propulsion systems (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An investigation of a simple self-similar flow model for an external nuclear pulse propulsion system indicates that to achieve the high effective specific impulse of such a system three principal factors are required. The are (1) attaining pulses of optimum energy, (2) attaining good propellant collimation, and (3) using an ablative material for the pusher surface which has high absorptivity for radiant energy at the propellant stagnation temperature.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TN-D-6984 , E-7013
    Format: application/pdf
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  • 95
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    Drive turbine system for 20-inch turbofan simulators. 2: Core turbine design (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Drive-turbines for a given set of 20-inch turbo-fan simulators are described. The simulators had both single-stage and two-stage fans that had design pressure ratios as low as 1.25 and as high as 3.0. The desired objective of the study was to be able to drive all of the single-stage fans with one core turbine and to drive all of the two-stage fans with this same core turbine in combination with a duct turbine. The core turbine is described. Included are the design operating conditions, design velocity diagram and a power-speed envelope determined by an off-design performance procedure. Also discussed is the adaption and scaling of an existing turbine design to this particular application.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-68130 , E-7121
    Format: application/pdf
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  • 96
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    Oxidation resistant porous material for transpiration cooled vanes (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Porous metal sheet with controlled permeability was made by space winding and diffusion bonding fine wire. Two iron-chromium-aluminum alloys and three-chromium alloys were used: GE 1541 (Fe-Cr-Al-Y), H 875 (Fe-Cr-Al-Si), TD Ni Cr, DH 245 (Ni-Cr-Al-Si) and DH 242 (Ni-Cr-Si-Cb). GE 1541 and H 875 were shown in initial tests to have greater oxidation resistance than the other candidate alloys and were therefore tested more extensively. These two materials were cyclic furnace oxidation tested in air at 1800 and 2000 F for accumulated exposure times of 4, 16, 64, 100, 200, 300, 400, 500, and and 600 hours. Oxidation weight gain, permeability change and mechanical properties were determined after exposure. Metallographic examination was performed to determine effects of exposure on the porous metal and electron beam weld joints of porous sheet to IN 100 strut material. Hundred hour stress rupture life and tensile tests were performed at 1800 F. Both alloys had excellent oxidation resistance and retention of mechanical properties and appear suitable for use as transpiration cooling materials in high temperature gas turbine engines.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-1999
    Format: application/pdf
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  • 97
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    Evaluation of range and distortion tolerance for high Mach number transonic fan stages. Task 2: Performance of a 1500-foot-per-second tip speed transonic fan stage with variable geometry inlet guide vanes and stator (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A 0.5 hub/tip radius ratio compressor stage consisting of a 1500 ft/sec tip speed rotor, a variable camber inlet guide vane and a variable stagger stator was designed and tested with undistorted inlet flow, flow with tip radial distortion, and flow with 90 degrees, one-per-rev, circumferential distortion. At the design speed and design IGV and stator setting the design stage pressure ratio was achieved at a weight within 1% of the design flow. Analytical results on rotor tip shock structure, deviation angle and part-span shroud losses at different operating conditions are presented. The variable geometry blading enabled efficient operation with adequate stall margin at the design condition and at 70% speed. Closing the inlet guide vanes to 40 degrees changed the speed-versus-weight flow relationship along the stall line and thus provided the flexibility of operation at off-design conditions. Inlet flow distortion caused considerable losses in peak efficiency, efficiency on a constant throttle line through design pressure ratio at design speed, stall pressure ratio, and stall margin at the 0 degrees IGV setting and high rotative speeds. The use of the 40 degrees inlet guide vane setting enabled partial recovery of the stall margin over the standard constant throttle line.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-72880 , R71AEG195
    Format: application/pdf
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  • 98
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    Procedure for estimating effects of ion beam interaction with spacecraft (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Procedure for calculating various interactions of neutral and charged particle efflux from ion thrusters with spacecraft surfaces is outlined. Calculation details are referenced. Details of charge-exchange calculation in the near-field of the thruster are presented.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-68043 , E-6880
    Format: application/pdf
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  • 99
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    Optimization of engines for a commercial Mach 0.98 transport using advanced turbine cooling methods (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A study was made of an advanced technology airplane using supercritical aerodynamics. Cruise Mach number was 0.98 at 40,000 feet altitude with a payload of 60,000 pounds and a range of 3000 nautical miles. Separate-flow turbofans were examined parametrically to determine the effect of sea-level-static design turbine-inlet-temperature and noise on takeoff gross weight (TOGW) assuming full-film turbine cooling. The optimum turbine inlet temperature was 2650 F. Two-stage-fan engines, with cruise fan pressure ratio of 2.25, achieved a noise goal of 103.5 EPNdB with todays noise technology while one-stage-fan engines, achieved a noise goal of 98 EPNdB. The take-off gross weight penalty to use the one-stage fan was 6.2 percent.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-68031 , E-6848
    Format: application/pdf
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  • 100
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    Feasibility study of a pressure-fed engine for a water recoverable space shuttle booster. Volume 1: Executive summary (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The activities leading to a tentative concept selection for a pressure-fed engine and propulsion support are outlined. Multiple engine concepts were evaluted through parallel engine major component and system analyses. Booster vehicle coordination, tradeoffs, and technology/development aspects are included. The concept selected for further evaluation has a regeneratively cooled combustion chamber and nozzle in conjuction with an impinging element injector. The propellants chosen are LOX/RP-1, and combustion stabilizing baffles are used to assure dynamic combustion stability.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-123538
    Format: application/pdf
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