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  • 1
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    Investigation of a Supersonic-inlet - Turbojet-engine Combination at Mach 2.0 and Angles of Attack up to 6 Degrees (1957)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Type: NACA-RM-E57D18
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    Comparison of effect of a turbojet engine and three cold-flow configurations on the stability of a full-scale supersonicle inlet (1957)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Increasing the volume and length of the duct behind the inlet affected the inlet stability at Mach 2.0 and zero angle of attack. Close approximation of the inlet stability limit of the J34 engine-inlet configuration was obtained by a cold-pipe configuration having a length and volume approaching that measured to the engine turbine. Variation of these parameters had a small effect on the minimum subcritical stable mass flow below a cowl-lip-position parameter of 44 degrees and appeared to have a negligible effect on the inlet pressure-recovery - mass-flow curve. Initial buzz frequency and minimum cowl-lip-position parameter for complete buzz-free operation varied with configuration.
    Type: NACA-RM-E56K23
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    Effects of Free-stream Reynolds Number, Engine Installation, and Model Scale on Stability Characteristics of a Translating-spike Inlet at Mach 2.0 (1957)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Type: NACA-RM-E57D17
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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    Overexpanded Performance of Conical Nozzles with Area Ratios of 6 and 9 With and Without Supersonic External Flow (1959)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: An investigation of the thrust characteristics and internal pressure distributions of two convergent-divergent 15 deg. half-angle exhaust nozzles having area ratios of 6 and 9 was made in the NASA Lewis 10- by 10-foot supersonic wind tunnel. The tests were conducted at free-stream Mach numbers of 0, 2.0, 2.5, 3.0, and 3.5 over a range of nozzle pressure ratios from 3 to 105. Attempts were made to induce separation of the overexpanded nozzle flow using secondary airflow and a wedge. Nozzle flow expansion under all free-stream conditions followed one-dimensional theory until separation from the nozzle wall occurred. In quiescent air the nozzle flow expanded to a pressure approximately one-half the base pressure before separation. When the nozzles were tested with supersonic external flow at the same effective pressure ratios, the nozzle flow separated with negligible expansion below the base pressure. The effect of a supersonic stream on internal nozzle flow separation characteristics was well defined only at a free-stream Mach number of 2.0. Thrust data at supersonic free-stream conditions indicate that only a small percentage of the ideal nozzle thrust will be available at nozzle pressure ratios below design. However, the overexpanded primary nozzle thrust loss was decreased by injecting large quantities of secondary air near the nozzle exit. In most cases no net gain in thrust resulted from secondary-air injection when the nozzle thrust was compared with the ideal thrust of both the primary and secondary airflows.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: NASA-TM-X-83 , E-208
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Base Flow Characteristics for Several Four-Clustered Rocket Configurations at Mach Numbers from 2.0 to 3.5 (1961)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: A generalized study of base flow phenomena has been conducted with four 500-pound-thrust JP-4 fuel-liquid-oxygen rocket motors installed in the base of a 12-inch-diameter cylindrical model. Data were obtained over a Mach number and nozzle pressure ratio range of 2.0 to 3.5 and 340 to 600, respectively. Base heat flux, gas temperature, and pressure were highest in the center of the cluster core and decreased in a radial direction. Although a maximum heat flux of 93 Btu per square foot per second was measured within the cluster core, peripheral heat fluxes were low, averaging about 5 Btu per square foot per second for all configurations. Generally base heat flux was found to be independent of Mach number over the range investigated. Base heat flux within the cluster core was decreased by increasing motor spacing, motor extension, a combination of increasing nozzle area ratio and decreasing exit angle and gimbaling the two side engines. Small amounts of nitrogen injected within the cluster core sharply reduced core heat flux.
    Keywords: Fluid Mechanics and Heat Transfer
    Type: NASA-TN-D-1093 , E-1241
    Format: application/pdf
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