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  • 1
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    Vacuum system for the Boiling Liquid-Metal Heat Transfer Facility (1966)
    In:  CASI
    Details
    Publication Date: 2006-10-26
    Description: Vacuum system for boiling liquid metal heat transfer facility
    Keywords: FACILITIES, RESEARCH, AND SUPPORT
    Format: text
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    NASA TECHNICAL REPORTS
  • 2
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    Conventional profile coaxial jet noise prediction (1983)
    In:  Other Sources
    Details
    Publication Date: 2011-08-18
    Description: (Previously cited in issue 24, p. 4249, Accession no. A81-49743)
    Keywords: ACOUSTICS
    Format: text
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    NASA TECHNICAL REPORTS
  • 3
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    Progress in the Lewis Research Center Altitude Wind Tunnel (AWT) modeling program (1986)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The rehabilitation of the AWT at the NASA Lewis Research Center is under study with the goal of providing a modern subsonic wind tunnel for conducting propulsion system/airframe integration, isolated propulsion system, propulsion acoustics and adverse weather tests. Because of the increased Mach number capability (from Mach 0.6 to 0.9 plus) and the incorporation of acoustic and adverse weather capabilities into an existing tunnel, the AWT rehabilitation represents a significant technical challenge. In order to reduce the risk associated with such an undertaking, an extensive AWT modeling program is being conducted to guide and verify the tunnel design. Significant findings and progress in this modeling program are the subject of this paper.
    Keywords: RESEARCH AND SUPPORT FACILITIES (AIR)
    Type: AIAA PAPER 86-0757
    Format: text
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    NASA TECHNICAL REPORTS
  • 4
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    Evaluation of hydrogen fuel in a full-scale afterburner (1957)
    In:  CASI
    Details
    Publication Date: 2019-05-23
    Description: Combustion efficiency of hydrogen fuel for varying afterburner configurations
    Keywords: PROPELLANTS
    Type: NACA-RM-E57H06
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Low frequency noise in a quiet, clean, general aviation turbofan engine (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A quiet, clean, general aviation, turbofan engine was instrumented to measure the fluctuating pressures in the combustor, turbine exit duct, engine nozzle and the far field. Both a separate flow nozzle and an internal mixer nozzle were tested. The fluctuating pressure data are presented in overall pressure and power levels and in spectral plots. The combustor data are compared to recent theory and found to be in excellent agreement. The results indicate that microphone correction procedures for elevated mean pressures are questionable. Ordinary coherence function analysis suggests the presence of an additional low frequency noise source downstream of the turbine that is due to the turbine itself. Low frequency narrowband data and coherence function analysis are presented.
    Keywords: ACOUSTICS
    Type: NASA-TM-83520 , E-1879 , NAS 1.15:83520
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 6
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    Geometric factors affecting noise suppresion and thrust loss of divergent-lobe supersonic jet noise suppressor (1973)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The thrust loss and noise suppression of a divergent-lobe supersonic jet noise suppressor were experimentally determined over a range of nozzle pressure ratios of 1.5 to 4.0. These small-scale cold flow tests were made to determine the effect on thrust and noise of: suppressor length, rearward facing step height, suppressor divergence angle, and ejector shroud length and location. Noise suppression was achieved at nozzle pressure ratios of 2.5 and greater. Maximum lobe jet noise attenuation of 15 db with thrust loss differences of 1.5 percent compared to the convergent nozzle were obtained at a nozzle pressure ratio of 3.5 with an ejector shroud two nozzle diameters long. Without the ejector the attenuation was 13 db with thrust loss differences of 11 percent. Short suppressors approximately one primary nozzle throat diameter long performed as well as longer suppressors. Rearward facing step height had a significant effect on noise suppression. Ejector shrouds two nozzle diameters in length are feasible.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-2820 , E-7393
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 7
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    Peak axial-velocity decay with multi-element rectangular and triangular nozzles (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The aircraft noise created by the impingement of engine exhaust jet of STOL aircraft with externally blown flaps is discussed. It was determined that the jet-flap interaction noise can be lowered by reducing the impinging velocity of the jet. The reduction must occur at a specific distance from the flap to be effective. The peak axial-velocity decay obtained with rectangular and triangular single element mixer nozzles is presented. Equations are developed for estimating the peak axial velocity decay curves for a wide range of nozzle configurations.
    Keywords: AIRCRAFT
    Type: NASA-TM-X-68047
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 8
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    Mixer nozzle-externally blown flap noise tests (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A summary is given of the noise suppression tests conducted for the STOL aircraft. The tests were made using a large scale mixer nozzle and externally blown flap model. Data were obtained over a range of nozzle exhaust velocities (172 to 284 m/sec) and flap angles. Comparisons were made between the results of the mixer nozzle and those obtained with a standard single convergent nozzle. The resulting conclusions show that a reduction in noise level did occur using the mixer nozzle system.
    Keywords: AIRCRAFT
    Type: NASA-TM-X-68021 , E-6827
    Format: application/pdf
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  • 9
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    JT150 1/2-scale nozzle jet noise experiment and comparison with prediction (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: As part of a program to study flight effects on the exhaust noise of a full scale JT15D engine, static half scale model jet noise experiments were conducted. Acoustic data were recorded for microphone angles of 45 deg to 155 deg with jet conditions for the model scale nozzle corresponding closely to those at 55, 73 and 97 percent of corrected rated speed for the full scale engine. These data are useful for determining the relative importance of jet and core noise in the static full scale engine test data and will in turn allow for a proper evaluation of flight effects on the exhaust noise results. The model scale data are also compared with the coaxial jet noise prediction. Above 1000 Hz, the prediction is nominally 0 to 3 dB higher than the data. The arithmetic mean of the differences between the experimental OASPL and the predicted OASPL for all angles for each run ranged from 0 to -3.2 dB. The standard deviation of all the OASPL differences is 2.2 dB. The discrepancies are greatest at low primary jet velocities and appear to be due to inadequacy in the variable jet density exponent incorporated in the prediction procedure.
    Keywords: ACOUSTICS
    Type: NASA-TM-83370 , E-1636 , NAS 1.15:83370
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 10
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    Splitting supersonic nozzle flow into separate jets by overexpansion into a multilobed divergent nozzle (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Air flowing from a convergent nozzle at pressure ratios greater than 2.5 has been split into eight separate jets by overexpansion of the flow into a divergent, eight-lobed passage. The splitting of the flow is accompanied by a decrease in the nozzle axial centerline Mach number. This in part is due to the radial inflow of secondary air between the lobes toward the nozzle centerline. Each of the smaller jets is partially split after it leaves the end of the divergent lobed section of the nozzle, thus creating a velocity profile having 16 peaks. At a pressure ratio of 3.5 the flow decelerates to Mach 1 in three convergent nozzle throat diameters. Convergent nozzle flow normally requires 12 diameters to reach Mach 1. The nozzle has a sound attenuation of 12 decibels with a thrust loss of 9 percent for the best configuration tested.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-6667 , E-6640
    Format: application/pdf
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