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  • 1
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    Laminar heating in interior corners at Mach 19 (1970)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: Laminar heating in hypersonic vehicles interior corners, analyzing helium tunnel heat transfer data for various intersecting wedge corners
    Keywords: THERMODYNAMICS AND COMBUSTION
    Type: ; ACE(
    Format: text
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    NASA TECHNICAL REPORTS
  • 2
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    Effect of forward-facing jets on aerodynamic characteristics of blunt configurations at Mach 6. (1967)
    In:  Other Sources
    Details
    Publication Date: 2011-08-11
    Description: Forward facing jets effect on blunt configurations aerodynamic characteristics from wind tunnel tests at Mach 6, emphasizing drag increase
    Keywords: AERODYNAMICS
    Type: JOURNAL OF SPACECRAFT AND ROCKETS
    Format: text
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  • 3
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    Transition in compressible free shear layers. (1972)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: The transition Reynolds number for shear layers produced by interactions between weak and strong shock waves is determined on the basis of experiments performed in a 20-in. (Mach 6) and an 11-in. (Mach 6.9) hypersonic tunnel. A variable angle wedge was used to generate a planar shock wave which interacted with the bow wave of a blunt body. An average value of the transition length (defined as the length along the shear layer from the shock interaction to the point where turbulence became visible on schlieren photographs) was used to determine the transition Reynolds number.
    Keywords: AERODYNAMICS
    Type: Journal of Spacecraft and Rockets; 9; Aug. 197
    Format: text
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    NASA TECHNICAL REPORTS
  • 4
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    Correlations of peak heating in shock interference regions at hypersonic speeds. (1972)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: Correlations are given of measured pressure and heat-transfer peaks for shock/boundary-layer interactions and shear layer attachment on configurations with both two- and three-dimensional interactions. The peak values were obtained from an investigation of shock interference heating on hemispheres, a 30-deg included angle wedge, and a 2.54-cm-diam cylindrical leading-edge fin model. The investigation covers data for Mach numbers of 6 and 20 over freestream Reynolds numbers ranging from (3.3 to 25.6) million per meter, and specific heat ratios of 1.4 and 1.67.
    Keywords: AERODYNAMICS
    Type: Journal of Spacecraft and Rockets; 9; Aug. 197
    Format: text
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    NASA TECHNICAL REPORTS
  • 5
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    Shock interference peak heating measurements using phase change coatings (1976)
    In:  Other Sources
    Details
    Publication Date: 2011-08-16
    Description: The phase change coating technique is used to obtain peak heating measurements in shock interference flow regions with high surface shear and heating. This technique provides heat transfer coefficients which are determined by measuring the time for a point on the surface to reach the phase change temperature of the thin fusible coating. Measurements were conducted on a 5.08-cm diameter hemisphere-cylinder made of silica based epoxy at Mach 6 for free stream Reynolds numbers of 3.3 to 25.6 million per meter. A sketch of the shock interference pattern produced by a flat plate shock generator is included. Heating data obtained on a 5.08-cm diameter stainless steel hemispherical model instrumented with thermocouples is presented for the purpose of comparing the phase change technique with the thermocouple-calorimeter technique.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: Journal of Spacecraft and Rockets; 13; Jan. 197
    Format: text
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  • 6
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    Correlation of turbulent shear layer attachment peak heating near Mach 6 (1977)
    In:  Other Sources
    Details
    Publication Date: 2011-08-17
    Description: A correlation of new turbulent two-dimensional data and peak heating data for attaching free shear layers is presented for a 2.54-cm and 5.08-cm diam cylindrical leading-edge slab 25.4 cm long, and 7.62 and 10.16 cm wide. A 30.48 x 25.4 cm sharp leading-edge flat plate set at 15 and 20 deg is used to generate plane impinging shocks. The freestream Mach number is 6 and the freestream Reynolds number varies from 3,300,000 to 25,600,000/m. Peak heating is measured on silica-based epoxy models with a phase change coating technique. A comparison of the free shear layer data with the transition data of Birch and Keyes (1972) reveals that the shear layer data are turbulent at attachment. The trend of the data shows that peak heating is strongly affected by the state of development at attachment. As the free shear layers become more fully developed, the data approach the two-dimensional correlation. Persistence of transitional flow structures for supersonic free shear flows is pointed out.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 15; Dec. 197
    Format: text
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  • 7
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    Turbulent heat transfer associated with control surface at Mach 6. (1968)
    In:  Other Sources
    Details
    Publication Date: 2011-08-12
    Description: Turbulent heat transfer associated with controls for attached or separated flow analyzed for design of hypersonic cruise aircraft
    Keywords: AERODYNAMICS
    Type: ; IVERSITET DRUZHBY NA
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  • 8
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    Noise reduction in a Mach 5 wind tunnel with a rectangular rod-wall sound shield (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A rod wall sound shield was tested over a range of Reynolds numbers of 0.5 x 10 to the 7th power to 8.0 x 10 to the 7th power per meter. The model consisted of a rectangular array of longitudinal rods with boundary-layer suction through gaps between the rods. Suitable measurement techniques were used to determine properties of the flow and acoustic disturbance in the shield and transition in the rod boundary layers. Measurements indicated that for a Reynolds number of 1.5 x 10 to the 9th power the noise in the shielded region was significantly reduced, but only when the flow is mostly laminar on the rods. Actual nozzle input noise measured on the nozzle centerline before reflection at the shield walls was attenuated only slightly even when the rod boundary layer were laminar. At a lower Reynolds number, nozzle input noise at noise levels in the shield were still too high for application to a quiet tunnel. At Reynolds numbers above 2.0 x 10 the the 7th power per meter, measured noise levels were generally higher than nozzle input levels, probably due to transition in the rod boundary layers. The small attenuation of nozzle input noise at intermediate Reynolds numbers for laminar rod layers at the acoustic origins is apparently due to high frequencies of noise.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-TP-1672 , L-13451
    Format: application/pdf
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  • 9
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    Experimental study of forward-located jet-interaction nozzles (1985)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The effects of forward-mounted sonic jet-interaction nozzles with one, three equal, or one large and two small orifices on the transitional or turbulent boundary-layer flow over a biconic (9-deg-25-arcmin/6-deg) model at angle of attack -10 to the 15 deg and Reynolds number 8.10 x 10 to the 6th/ft are investigated experimentally using pressure taps, force measurements, oil-flow visualization, and schlieren photography in the 20-in. Mach-6 wind tunel at NASA Langley. The injection flow rates are 0.055 and 0.026 lbm/s, with stagnation pressure 500 psia, stagnation temperature 1000 R, and dynamic pressure 8.7 psia. The results are presented in diagrams, graphs, and photographs and characterized. Moment amplification factors of 1.0 or better (maximum 1.8 at angle of attack 15 deg and flow rate 0.055 lbm/s) are observed for all nozzle types, but the configuration with three unequal orifices is found to give improvements of 10 percent over the single-orifice nozzle at a given flow rate.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 85-0454
    Format: text
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  • 10
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    Modifications to the nozzle test chamber to extend nozzle static-test capability (1985)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The nozzle test chamber was modified to provide a high-pressure-ratio nozzle static-test capability. Experiments were conducted to determine the range of the ratio of nozzle total pressure to chamber pressure and to make direct nozzle thrust measurements using a three-component strain-gage force balance. Pressure ratios from 3 to 285 were measured with several axisymmetric nozzles at a nozzle total pressure of 15 to 190 psia. Devices for measuring system mass flow were calibrated using standard axisymmetric convergent choked nozzles. System mass-flow rates up to 10 lbm/sec are measured. The measured thrust results of these nozzles are in good agreement with one-dimensional theoretical predictions for convergent nozzles.
    Keywords: AERODYNAMICS
    Type: NASA-TM-86368 , L-15674 , NAS 1.15:86368
    Format: application/pdf
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