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  • Aircraft Propulsion and Power  (10)
  • 1950-1954  (10)
  • 1935-1939
  • 1950  (10)
  • 1
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    Relation Between Inflammables and Ignition Sources in Aircraft Environments (1950)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A literature survey was conducted to determine the relation between aircraft ignition sources and inflammables. Available literature applicable to the problem of aircraft fire hazards is analyzed and, discussed herein. Data pertaining to the effect of many variables on ignition temperatures, minimum ignition pressures, and minimum spark-ignition energies of inflammables, quenching distances of electrode configurations, and size of openings incapable of flame propagation are presented and discussed. The ignition temperatures and the limits of inflammability of gasoline in air in different test environments, and the minimum ignition pressure and the minimum size of openings for flame propagation of gasoline - air mixtures are included. Inerting of gasoline - air mixtures is discussed.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-TN-2227
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    Investigation of the NACA 4-(3)(8)-045 Two-blade Propellers at Forward Mach Numbers to 0.725 to Determine the Effects of Compressibility and Solidity on Performance (1950)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: As part of a general investigation of propellers at high forward speeds, tests of two 2-blade propellers having the NACA 4-(3)(8)-03 and NACA 4-(3)(8)-45 blade designs have been made in the Langley 8-foot high-speed tunnel through a range of blade angle from 20 degrees to 60 degrees for forward Mach numbers from 0.165 to 0.725 to establish in detail the changes in propeller characteristics due to compressibility effects. These propellers differed primarily only in blade solidity, one propeller having 50 percent and more solidity than the other. Serious losses in propeller efficiency were found as the propeller tip Mach number exceeded 0.91, irrespective of forward speed or blade angle. The magnitude of the efficiency losses varied from 9 percent to 22 percent per 0.1 increase in tip Mach number above the critical value. The range of advance ratio for peak efficiency decreased markedly with increase of forward speed. The general form of the changes in thrust and power coefficients was found to be similar to the changes in airfoil lift coefficient with changes in Mach number. Efficiency losses due to compressibility effects decreased with increase of blade width. The results indicated that the high level of propeller efficiency obtained at low speeds could be maintained to forward sea-level speeds exceeding 500 miles per hour.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-TR-999
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    The Supersonic Axial-Flow Compressor (1950)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation has been made to explore the possibilities of axial-flow compressors operating with supersonic velocities into the blade rows. Preliminary calculations showed that very high pressure ratios across a stage, together with somewhat increased mass flows, were apparently possible with compressors which decelerated air through the speed of sound in their blading. The first phase of the investigation was the development of efficient supersonic diffusers to decelerate air through the speed of sound. The present report is largely a general discussion of some of the essential aerodynamics of single-stage supersonic axial-flow compressors. As an approach to the study of supersonic compressors, three possible velocity diagrams are discussed briefly. Because of the encouraging results of this study, an experimental single-stage supersonic compressor has been constructed and tested in Freon-12. In this compressor, air decelerates through the speed of sound in the rotor blading and enters the stators at subsonic speeds. A pressure ratio of about 1.8 at an efficiency of about 80 percent has been obtained.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-TR-974 , NACA-ACR-L6D02 , NACA-AR-36
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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    Theoretical Investigations on the Efficiency and the Conditions for the Realization of Jet Engines (1950)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: Contents: Preliminary notes on the efficiency of propulsion systems; Part I: Propulsion systems with direct axial reaction rockets and rockets with thrust augmentation; Part II: Helicoidal reaction propulsion systems; Appendix I: Steady flow of viscous gases; Appendix II: On the theory of viscous fluids in nozzles; and Appendix III: On the thrusts augmenters, and particularly of gas augmenters
    Keywords: Aircraft Propulsion and Power
    Type: NACA-TM-1259
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Performance of J33-A-27 Turbojet-Engine Compressor (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An investigation was conducted to determine the effects of water injection on the over-all performance of a modified J33-A-27 turbojet-engine compressor at the design equivalent speed of 11,800 rpm. The water-air ratio by weight was 0.05. With water injection the peak pressure ratio increased 9.0 per- cent, the maximum efficiency decreased 15 percent (actual numerical difference 0.12), and. the maximum total weight flow increased 9.3 percent.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE50F14
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 6
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    Investigation of Performance of Axial-Flow Compressor of XT-46 Turbine-Propeller Engine. II - Performance of Revised Compressor at Design Equivalent Speed (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: The compressor from the XT-46 turbine-propeller engine was revised by removing the last two rows of stator blades and by eliminating the interstage leakage paths described in a previous report. With the revised compressor, the flow choking point shifted upstream into the last rotor-blade row but the maximum weight flow was not increased over that of the original compressor. The flow range of the revised compressor was reduced to about two-thirds that obtained with the original compressor. The later stages of the compressor did not produce the design static-pressure increase probably because of excessive boundary-layer build-up in this region. Measurements obtained in the ninth-stage stator showed that the performance up to this station was promising but that the last three stages of the compressor were limiting the useful operating range of the preceding stages. Some modifications in flow-passage geometry and blade settings are believed to be necessary, however, before any major improvements in over-all compressor performance can be obtained.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE50J10
    Format: application/pdf
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  • 7
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    Summary of Investigations of Mark 25 Aerial-Torpedo Turbine (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: The power plant from a Mark 25 aerial torpedo was investigated both as a two-stage turbine and as a single-stage modified turbine to determine the effect on overall performance of nozzle size and shape, first-stage rotor-blade configuration, and axial nozzle-rotor running clearance. Performance was evaluated in terms of brake, rotor, and blade efficiencies. All the performance data were obtained for inlet total to outlet static pressure ratios of 8, 15 (design), and 20 with inlet conditions maintained constant at 95 pounds per square inch gage and 1000 F for rotor speeds from approximately 6000 to 18,000 rpm.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE50D12
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 8
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    Preliminary Performance Data on General Electric Integrated Electronic Control Operating on J47 RX1-3 Turbojet Engine in NACA Altitude Wind Tunnel (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: Performance data obtained with recording oscillographs are presented to show the transient response of the General Electric Integrated Electronic Control operating on the J47 RXl-3 turbo-Jet engine over a range of altitudes from 10,000 to 45,000 feet and at ram pressure ratios of 1.03 and 1.4. These data represent the performance of the final control configuration developed after an investigation of the engine transient behavior in the NACA altitude wind tunnel. Oscillograph traces of controlled accelerations (throttle bursts),oontrolled decelerations (throttle chops), and controlled altitude starts are presented.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE50G12
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 9
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    Performance of J33-A-27 Turbojet-Engine Compressor (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: A modified J33-A-27 compressor was operated over a range of equivalent impeller speeds from 6100 to 13,250 rpm in order to obtain the over-all compressor performance. At the equivalent design speed of 11,800 rpm, the maximum efficiency of 0.764 and peak pressure ratio of 4.56 occurred at an equivalent weight flow of 104.07 pounds per second. At the highest equivalent speed (13,250 rpm) a maximum efficiency of 0.711, a maximum equivalent weight flow of 123.00 pounds per second, and a peak pressure ratio of 5.76 were obtained.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE50D25
    Format: application/pdf
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  • 10
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    Investigation of Performance of Axial-Flow Compressor of XT-46 Turbine-Propeller Engine. I - Preliminary Investigation at 50-,70-, and 100-Percent Design Equivalent Speed (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: An investigation is being conducted to determine the performance of the 12-stage axial-flow compressor of the XT-46 turbine-propeller engine. This compressor was designed to produce a pressure ratio of 9 at an adiabatic efficiency of 0.86. The design pressure ratios per stage were considerably greater than any employed in current aircraft gas-turbine engines using this type of compressor. The compressor performance was evaluated at two stations. The station near the entrance section of the combustors indicated a peak pressure ratio of 6.3 at an adiabatic efficiency of 0.63 for a corrected weight flow of 23.1 pounds per second. The other, located one blade-chord downstream of the last stator row, indicated a peak pressure ratio of 6.97 at an adiabatic efficiency of 0.81 for a corrected weight flow of 30.4 pounds per second. The difference in performance obtained at the two stations is attributed to shock waves in the vicinity of the last stator row. These shock waves and the accompanying flow choking, together with interstage circulatory flows, shift the compressor operating curves into the region where surge would normally occur. The inability of the compressor to meet design pressure ratio is probably due to boundary-layer buildup in the last stages, which cause axial velocities greater than design values that, in turn, adversely affect the angles of attack and turning angles in these blade rows.
    Keywords: Aircraft Propulsion and Power
    Type: NACA-RM-SE50E22
    Format: application/pdf
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