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  • 1
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    Comparison of the noise characteristics of two low pressure ratio fans with a high throat Mach number inlet (1978)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Acoustics data obtained in experiments with two low pressure ratio 50.8 cm (20 in.) diameter model fans differing in design tip speed were compared. Determination of the average throat Mach number used to compare high Mach inlet noise reduction characteristics was based on a correlation of inlet wall static pressure measurements with a flow field calculation. The largest noise reductions were generally obtained with the higher tip speed fan. At a throat Mach number of 0.79, the difference in noise reduction was about 3.5 db with static test conditions. Although the noise reduction increased for the lower tip speed fan with a simulated flight velocity of 41 m/sec (80 knots), it was still about 2 db less than that of the high tip speed fan which was only tested at the static condition. However, variations in acoustic performance could not be absolutely attributed to the different fan designs because of differences in inlet lip contours which resulted in small variations of peak wall Mach number and axial extend of supersonic and near-sonic flow.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: NASA-TM-73880 , E-9489
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    Simulated flight effects on noise characteristics of a fan inlet with high throat Mach number (1978)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An anechoic wind tunnel experiment was conducted to determine the effects of simulated flight on the noise characteristics of a high throat Mach number fan inlet. Comparisons were made with the performance of a conventional low throat Mach number inlet with the same 50.8 cm fan noise source. Simulated forward velocity of 41 m/sec reduced perceived noise levels for both inlets, the largest effect being more than 3 db for the high throat Mach number inlet. The high throat Mach number inlet was as much as 7.5 db quieter than the low throat Mach number inlet with tunnel airflow and about 6 db quieter without tunnel airflow. Effects of inlet flow angles up to 30 deg were seemingly irregular and difficult to characterize because of the complex flow fields and generally small noise variations. Some modifications of tones and directivity at blade passage harmonics resulting from inlet flow angle variation were noted.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: NASA-TP-1199 , E-9253
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    Boundary-layer measurements in accelerated flows near Mach 1 (1967)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Velocity and temperature profile measurements for adiabatic boundary layers on cylindrical bodies in transonic parallel accelerated flows
    Keywords: FLUID MECHANICS
    Type: NASA-TN-D-3882
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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    Effect of entry-lip design on aerodynamics and acoustics of high throat Mach number inlets for the quiet, clean, short-haul experimental engine (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Results of scale model tests of high-throat-Mach-number inlets designed to suppress inlet-emitted engine machinery noise produced in a V/STOL wind tunnel are presented. A vacuum system was used to induce inlet airflow with a siren as a noise source. Inlet mass flow was 11.68 kilograms (25.75 lb. min) per second at a throat Mach number of 0.79. The effect of entry-lip design (contraction ratio and diameter ratio) on inlet total-pressure recovery, steady-state pressure distortion, performance at high incidence angles, and noise suppression was determined. With proper entry-lip design, total-pressure recovery in excess of 0.988 could be obtained statically at an average throat Mach number of 0.79. Total-pressure distortion was 5 percent. The reduction in the siren tone sound pressure level transmitted through the inlet was 10 to 14 db relative to that measured at throat Mach 0.6.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-3222 , E-8160
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Low-speed wind tunnel tests of a 50.8-centimeter (20-in.) 1.15-pressure-ratio fan engine model (1974)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: At a typical STOL aircraft takeoff and landing velocity, wind tunnel aerodynamic and acoustic measurements demonstrated that an inlet lip-area contraction ratio of 1.35 was superior to a ratio of 1.26 at high incidence angles. A 17 percent reduction in net thrust and an increase of 9 decibels in sound pressure level at the blade passing frequency resulted from inlet flow separation at an incidence angle of 50 deg with the 1.26-contraction-ratio inlet. Reverse-thrust forces obtained with blade rotation through the feathered angle were 1.8 times larger than with blade rotation through the flat angle. Reverse-thrust force was reduced from 30 to 50 percent and sound pressure level increased from 3 to 7 decibels at the blade passing frequency between the wind-tunnel-off condition and a typical STOL aircraft landing velocity.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-TM-X-3062 , E-7844
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 6
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    Boundary layer measurements in an accelerated flow with and without heat transfer (1971)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Boundary layer measurements in accelerated flow with and without heat transfer
    Keywords: FLUID MECHANICS
    Type: NASA-TN-D-7030 , E-5674
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 7
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    Wind tunnel tests of a 20 inch diameter 1.15 pressure ratio fan engine model (1973)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: Aerodynamic and acoustic measurements at a typical STOL aircraft takeoff and landing velocity demonstrated that a 1.35 inlet lip area contraction ratio was superior to a 1.26 ratio at high nacelle incidence angles. Reverse thrust, obtained with a variable pitch rotor, was lower at the landing velocity, and the noise level higher, than at the static condition. High speed tests showed that, for the design cruise Mach number of 0.75, internal losses and external drag were 27 per cent of the ideal fan net thrust, and propulsive efficiency was estimated to be 59 per cent for an 85 per cent efficient fan stage. For comparison, a similar 1.55 pressure ratio fan system would have a propulsive efficiency of 62 per cent.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 73-1216 , Propulsion Conference; Nov 05, 1973 - Nov 07, 1973; Las Vegas, NV; US
    Format: text
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    NASA TECHNICAL REPORTS
  • 8
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    Wind tunnel tests of a 20 inch diameter 1.15 pressure ratio fan engine model (1973)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Aerodynamic and acoustic measurements at a typical STOL aircraft takeoff and landing velocity demonstrated that a 1.35 inlet lip area contraction ratio was superior to a 1.26 ratio at high nacelle incidence angles. Reverse thrust, obtained with a variable pitch rotor, was lower at the landing velocity, and the noise level higher, than at the static condition. High speed tests showed that, for the design cruise Mach number of 0.75, internal losses and external drag were 27 percent of the ideal fan net thrust, and propulsive efficiency was estimated to be 59 percent for an 85 percent efficient fan stage.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-71445 , E-7705 , Propulsion Joint Specialist Conf.; Nov 05, 1973 - Nov 07, 1973; Las Vegas, NV; United States
    Format: application/pdf
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    NASA TECHNICAL REPORTS
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