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  • AERODYNAMICS  (484)
  • 1980-1984  (483)
  • 1945-1949  (1)
  • 1935-1939
  • 11
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    Technology developments for laminar boundary layer control on subsonic transport aircraft (1984)
    In:  CASI
    Details
    Publication Date: 2017-10-02
    Description: The development of laminar flow technology for commercial transport aircraft is discussed and illustrated in a review of studies undertaken in the NASA Aircraft Energy Efficiency (ACEE) program since 1976. The early history of laminar flow control (LFC) techniques and natural laminar flow (NLF) airfoil designs is traced, and the aims of ACEE are outlined. The application of slotted structures, composites, and electron beam perforated metals in supercritical LFC airfoils, wing panels, and leading edge systems is examined; wind tunnel and flight test results are summarized; studies of high altitude ice effects are described; and hybrid (LFC/NLF designs are characterized. Drawings and photographs are provided.
    Keywords: AERODYNAMICS
    Type: AGARD Improvement of Aerodynamic Performance Through Boundary Layer Control and High Lift Systems; 13 p
    Format: text
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    NASA TECHNICAL REPORTS
  • 12
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    Experimental Results for a Flapped Natural-laminar-flow Airfoil with High Lift/drag Ratio (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Experimental results have been obtained for a flapped natural-laminar-flow airfoil, NLF(1)-0414F, in the Langley Low-Turbulence Pressure Tunnel. The tests were conducted over a Mach number range from 0.05 to 0.40 and a chord Reynolds number range from about 3.0 x 10(6) to 22.0 x 10(6). The airfoil was designed for 0.70 chord laminar flow on both surfaces at a lift coefficient of 0.40, a Reynolds number of 10.0 x 10(6), and a Mach number of 0.40. A 0.125 chord simple flap was incorporated in the design to increase the low-drag, lift-coefficient range. Results were also obtained for a 0.20 chord split-flap deflected 60 deg.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85788 , NAS 1.15:85788
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 13
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    Boundary layer transition effects on flow separation around V/STOL engine inlets at high incidence (1984)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Numerical methods for calculating laminar and turbulent boundary layer development around vertical-short take off and landing engine inlets at high incidence angles are investigated. Various transition models were compared and evaluated in calculations off flow separation bound inside the inlet. Results of the transition effects on the boundary layer characteristics at onset of separation for two types of engine inlet geometries are presented. Some of the numerical results are compared with existing wind-tunnel test data for scaled inlet models to demonstrate the effects of transition models in the numerical scheme. The effects of transition modeling on the boundary layer development are illustrated for typical engine operating conditions.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0432
    Format: text
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    NASA TECHNICAL REPORTS
  • 14
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    Convair 990 transonic flow-field simulation about the forward fuselage (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A three-dimensional, viscous flow code was used to calculate the transonic flow about the forebody of the Convair CV-990 (Galileo II) research aircraft stationed at NASA Ames Research Center. The computations were used to determine the location for a differential pressure system. In addition, attitude sensor placements were verified. These instruments comprise a meteorological measurement system, which will be used for global determination of three-dimensional wind data. The code solves the thin layer form of the Reynolds-averaged Navier-Stokes equations using an implicit numerical procedure. The governing equations are written in a generalized, nonorthogonal coordinate system, and are cast in a strong conservation law form. Laminar boundary layer results are presented for free stream Mach number of 0.8 and angles of attack of zero and 2 deg. Use of this computational tool reduced the development time for the location of the sensors and aided in the optimal placement on the aircraft of these devices.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1785
    Format: text
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    NASA TECHNICAL REPORTS
  • 15
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    Calculation of a simulated 3-D high speed inlet using the Navier-Stokes equations (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A hybrid numerical algorithm, developed to solve the full three-dimensional Navier-Stokes equations, is applied to the computation of the flowfield in a simulated three-dimensional high speed aircraft inlet at a Mach number of 2.5 and Reynolds number of 1.4 x 10 to the 7th based on inlet length. The numerical algorithm incorporates a coordinate transformation in order to handle general flow geometries, and utilizes the algebraic turbulent eddy viscosity model of Baldwin and Lomax. The hybrid algorithm has been vectorized on the CDC CYBER 203 computer using the SL/1 vector programming language developed at NASA Langley. The computed results are compared with experimental measurements of the ramp and cowl static pressures, and boundary layer pitot profiles. The results are also compared with a previous two-dimensional Navier-Stokes computation of the same configuration. The agreement with the experimental data is generally good; however, additional improvements in turbulence modeling are needed.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1165
    Format: text
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    NASA TECHNICAL REPORTS
  • 16
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    Multiple steady states for characteristic initial value problems (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The time dependent, isentropic, quasi-one-dimensional equations of gas dynamics and other model equations are considered under the constraint of characteristic boundary conditions. Analysis of the time evolution shows how different initial data may lead to different steady states and how seemingly anamolous behavior of the solution may be resolved. Numerical experimentation using time consistent explicit algorithms verifies the conclusions of the analysis. The use of implicit schemes with very large time steps leads to erroneous results.
    Keywords: AERODYNAMICS
    Type: NASA-CR-172486 , ICASE-84-57 , NAS 1.26:172486
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 17
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    High Reynolds number tests of the cast 10-2/DOA 2 airfoil in the Langley 0.3-meter transonic cryogenic tunnel, phase 2 (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Wind tunnel tests of an advanced technology airfoil, the CAST 10-2/DOA 2, were conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT). This was the third of a series of tests conducted in a cooperative airfoil research program between the National Aeronautics and Space Administration and the Deutsche Forschungsund Versuchsanstalt fur Luft- und Raumfahrt e. V. For these tests, temperature was varied from 270 K to 110 K at pressures from 1.5 to 5.75 atmospheres. Mach number was varied from 0.60 to 0.80, and the Reynolds number (based on airfoil chord) was varied from 2 to 20 million. The aerodynamic data for the 7.62 cm chord airfoil model used in these tests is presented without analysis. Descriptions of the 0.3-m TCT, the airfoil model, the test instrumentation, and the testing procedures are included.
    Keywords: AERODYNAMICS
    Type: NASA-TM-86273 , NAS 1.15:86273
    Format: application/pdf
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  • 18
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    Unsteady transonic small disturbance theory with strong shock waves (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A theory to correct the transonic small disturbance (TSD) equation to treat strong shock waves in unsteady flow is developed. The technique involves the addition of higher order terms, which are formally of negligible magnitude, to the low frequency TSD equation. These terms are then chosen such that any shock waves in the flow have strengths approximately equal to the appropriate Rankine-Hugoniot shock strength. Two correcting approaches are investigated. The first is to derive a correction for the mean steady flow and then simply use this corrected form for oscillatory flows. The second is to derive a correction for both steady and oscillatory parts of the flow. This second development is the most satisfactory and comparisons of the present results with Euler equation results are generally favorable, particularly regarding shock location, although there are some discrepancies in the pressure distribution in the leading edge region.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166157 , NEAR-TR-230
    Format: application/pdf
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  • 19
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    Numerical lifting line theory applied to drooped leading-edge wings below and above stall (1980)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A numerical iterative solution to the classical Prandtl lifting-line theory, suitably modified for poststall behavior, is used to study the aerodynamic characteristics of straight rectangular finite wings with and without leading-edge droop. This study is prompted by the use of such leading-edge modifications to inhibit stall/spins in light general aviation aircraft. The results indicate that lifting-line solutions at high angle of attack can be obtained that agree with experimental data to within 20%, and much closer for many cases. Therefore, such solutions give reasonable preliminary engineering results for both drooped and undrooped wings in the poststall region. However, as predicted by von Karman, the lifting-line solutions are not unique when sectional negative lift slopes are encountered. In addition, the present numerical results always yield symmetrical lift distributions along the span, in contrast to the asymmetrical solutions observed by Schairer in the late 1930's. Finally, a series of parametric tests at low angle of attack indicate that the effect of drooped leading edges on aircraft cruise performance is minimal.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 17; Dec. 198
    Format: text
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  • 20
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    Combined linear theory/impact theory method for analysis and design of high speed configurations (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Pressure distributions on a wing body at Mach 4.63 are calculated. The combined theory is shown to give improved predictions over either linear theory or impact theory alone. The combined theory is also applied in the inverse design mode to calculate optimum camber slopes at Mach 4.63. Comparisons with optimum camber slopes obtained from unmodified linear theory show large differences. Analysis of the results indicate that the combined theory correctly predicts the effect of thickness on the loading distributions at high Mach numbers, and that finite thickness wings optimized at high Mach numbers using unmodified linear theory will not achieve the minimum drag characteristics for which they are designed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3314
    Format: application/pdf
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