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  • 1
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    Comparison of theoretical and flight-measured local flow aerodynamics for a low-aspect-ratio fin (1986)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Flight test and theoretical aerodynamic data were obtained for a flight test fixture mounted on the underside of an F-104G aircraft. The theoretical data were generated using two codes: a two-dimensional transonic code called code H, and a three-dimensional subsonic and supersonic code called wing-body. Pressure distributions generated by the codes for the flight test fixture, as well as compared with the flight-measured data. The two-dimensional code pressure distributions compared well except at the minimum pressure point and the trailing edge. Shock locations compared well except at high transonic speeds. However, the two-dimensional code did not adequately predict the displacement thickness of the flight test fixture. The three-dimensional code pressure distributions compared well except at the trailing edge of the flight test fixture.
    Keywords: AERODYNAMICS
    Type: NASA-TM-86806 , H-1336 , NAS 1.15:86806
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    Digital program for calculating static pressure position error (1987)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A computer program written to calculate the static pressure position error of airspeed systems contains five separate methods for determining position error, of which the user may select from one to five at a time. The program uses data from both the test aircraft and the ground-based radar to calculate the error. In addition, some of the methods require rawinsonde data or an atmospheric analysis, or both. The program output lists the corrections to Mach number, altitude, and static pressure that are due to position error. Reference values such as angle of attack, angle of sideslip, indicated Mach number, indicated pressure altitude, stagnation pressure, and total temperature are also listed.
    Keywords: AIRCRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA-TM-86726 , H-1284 , NAS 1.15:86726
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    Qualitative evaluation of a flush air data system at transonic speeds and high angles of attack (1987)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Flight tests were performed on an F-14 aircraft to evaluate the use of flush pressure orifices on the nose section for obtaining air data at transonic speeds over a large range of flow angles. This program was part of a flight test and wind tunnel program to assess the accuracies of such systems for general use on aircraft. It also provided data to validate algorithms developed for the shuttle entry air data system designed at NASA Langley. Data were obtained for Mach numbers between 0.60 and 1.60, for angles of attack up to 26.0 deg, and for sideslip angles up to 11.0 deg. With careful calibration, a flush air data system with all flush orifices can provide accurate air data information over a large range of flow angles. Several orificies on the nose cap were found to be suitable for determination of stagnation pressure. Other orifices on the nose section aft of the nose cap were shown to be suitable for determination of static pressure. Pairs of orifices on the nose cap provided the most sensitive measurements for determining angles of attack and sideslip, although orifices located farther aft on the nose section could also be used.
    Keywords: AIRCRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA-TP-2716 , H-1277 , NAS 1.60:2716
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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    Preliminary In-Flight Boundary Layer Transition Measurements on a 45 Deg Swept Wing at Mach Numbers Between 0.9 and 1.8 (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A preliminary flight experiment was flown to generate a full-scale supersonic data base to aid the assessment of computational codes, to improve instrumentation for measuring boundary layer transition at supersonic speeds, and to provide preliminary information for the definition of follow-on programs. The experiment was conducted using an F-15 aircraft modified with a small cleanup test section on the right wing. Results are presented for Mach (M) numbers from 0.9 to 1.8 at altitudes from 25,000 to 55,000 ft. At M greater than or = 1.2, transition occurred near or at the leading edge for the clean configuration. The furthest aft that transition was measured was 20 percent chord at M = 0.9 and M = 0.97. No change in transition location was observed after the addition of a notch-bump on the leading edge of the inboard side of the test section which was intended to minimize attachment line transition problems. Some flow visualization was attempted during the flight experiment with both subliming chemicals and liquid crystals. However, difficulties arose from the limited time the test aircraft was able to hold test conditions and the difficulty of positioning the photo chase aircraft during supersonic test points. Therefore, no supersonic transition results were obtained.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA-TM-100412 , H-1436 , NAS 1.15:100412
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Hot-Film Anemometer For Boundary-Flow Transitions (1988)
    In:  Other Sources
    Details
    Publication Date: 2019-07-12
    Description: Temperature-compensated instrument yields data at subsonic and supersonic speeds. Modifications in new anemometer include addition of temperature-compensation resistor and resistors Rs and Rp series and parallel with compensation device.
    Keywords: ELECTRONIC COMPONENTS AND CIRCUITS
    Type: ARC-11811 , NASA Tech Briefs (ISSN 0145-319X); 12; 7; P. 26
    Format: text
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    NASA TECHNICAL REPORTS
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