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  • AERODYNAMICS  (342)
  • 1980-1984  (342)
  • 1925-1929
  • 1980  (342)
  • 101
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    A flight investigation of blade section aerodynamics for a helicopter main rotor having NLR-1T airfoil sections (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A flight investigation was conducted using a teetering-rotor AH-1G helicopter to obtain data on the aerodynamic behavior of main-rotor blades with the NLR-1T blade section. The data system recorded blade-section aerodynamic pressures at 90 percent rotor radius as well as vehicle flight state, performance, and loads. The test envelope included hover, forward flight, and collective-fixed maneuvers. Data were obtained on apparent blade-vortex interactions, negative lift on the advancing blade in high-speed flight and wake interactions in hover. In many cases, good agreement was achieved between chordwise pressure distributions predicted by airfoil theory and flight data with no apparent indications of blade-vortex interactions.
    Keywords: AERODYNAMICS
    Type: NASA-TM-80166 , AVRADCOM-TR-80-B-2
    Format: application/pdf
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  • 102
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    Rotary balance data for a typical single-engine general aviation design for an angle-of-attack range of 8 deg to 90 deg. 1: Low-wing model A (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Aerodynamic characteristics obtained in a rotational flow environment utilizing a rotary balance are presented in plotted form for a 1/5 scale, single engine, low-wing, general aviation airplane model. The configuration tested included the basic airplane, various control deflections, tail designs, fuselage shapes, and wing leading edges. Data are presented without analysis for an angle of attack range of 8 to 90 deg and clockwise and counterclockwise rotations covering a range from 0 to 0.85.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3100
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  • 103
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    A calculation procedure for viscous flow in turbomachines, volume 2 (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Turbulent flow within turbomachines having arbitrary blade geometries is examined. Effects of turbulence are modeled using two equations, one expressing the development of the turbulence kinetic energy and the other its dissipation rate. To account for complicated blade geometries, the flow equations are formulated in terms of a nonorthogonal boundary fitted coordinate system. The analysis is applied to a radial inflow turbine. The solution obtained indicates the severity of the complex interaction mechanism that occurs between the different flow regimes (i.e., boundary layers, recirculating eddies, separation zones, etc.). Comparison with nonviscous flow solutions tend to justify strongly the inadequacy of using the latter with standard boundary layer techniques to obtain viscous flow details within turbomachine rotors. Capabilities and limitations of the present method of analysis are discussed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-159636
    Format: application/pdf
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  • 104
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    Full scale visualization of the wing tip vortices generated by a typical agricultural aircraft (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The trajectories of the wing tip vortices of a typical agricultural aircraft were experimentally determined by flight test. A flow visualization method, similar to the vapor screen method used in wind tunnels, was used to obtain trajectory data for a range of flight speeds, airplane configurations, and wing loadings. Detailed measurements of the spanwise surface pressure distribution were made for all test points. Further, a powered 1/8 scale model of the aircraft was designed, built, and used to obtain tip vortex trajectory data under conditions similar to that of the full scale test. The effects of light wind on the vortices were demonstrated, and the interaction of the flap vortex and the tip vortex was clearly shown in photographs and plotted trajectory data.
    Keywords: AERODYNAMICS
    Type: NASA-CR-162796 , MSSU-EIRS-ASE-80-2
    Format: application/pdf
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  • 105
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    Wind-tunnel/flight correlation study of aerodynamic characteristics of a large flexible supersonic cruise airplane (XB-70-1). 3: A comparison between characteristics predicted from wind-tunnel measurements and those measured in flight (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A program was undertaken by NASA to evaluate the accuracy of a method for predicting the aerodynamic characteristics of large supersonic cruise airplanes. This program compared predicted and flight-measured lift, drag, angle of attack, and control surface deflection for the XB-70-1 airplane for 14 flight conditions with a Mach number range from 0.76 to 2.56. The predictions were derived from the wind-tunnel test data of a 0.03-scale model of the XB-70-1 airplane fabricated to represent the aeroelastically deformed shape at a 2.5 Mach number cruise condition. Corrections for shape variations at the other Mach numbers were included in the prediction. For most cases, differences between predicted and measured values were within the accuracy of the comparison. However, there were significant differences at transonic Mach numbers. At a Mach number of 1.06 differences were as large as 27 percent in the drag coefficients and 20 deg in the elevator deflections. A brief analysis indicated that a significant part of the difference between drag coefficients was due to the incorrect prediction of the control surface deflection required to trim the airplane.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1516 , H-1079
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  • 106
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    An experimental investigation of two large annular diffusers with swirling and distorted inflow (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Two annular diffusers downstream of a nacelle-mounted fan were tested for aerodynamic performance, measured in terms of two static pressure recovery parameters (one near the diffuser exit plane and one about three diameters downstream in the settling duct) in the presence of several inflow conditions. The two diffusers each had an inlet diameter of 1.84 m, an area ratio of 2.3, and an equivalent cone angle of 11.5, but were distinguished by centerbodies of different lengths. The dependence of diffuser performance on various combinations of swirling, radially distorted, and/or azimuthally distorted inflow was examined. Swirling flow and distortions in the axial velocity profile in the annulus upstream of the diffuser inlet were caused by the intrinsic flow patterns downstream of a fan in a duct and by artificial intensification of the distortions. Azimuthal distortions or defects were generated by the addition of four artificial devices (screens and fences). Pressure recovery data indicated beneficial effects of both radial distortion (for a limited range of distortion levels) and inflow swirl. Small amounts of azimuthal distortion created by the artificial devices produced only small effects on diffuser performance. A large artificial distortion device was required to produce enough azimuthal flow distortion to significantly degrade the diffuser static pressure recovery.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1628 , AVRADCOM-TR-79-40 , A-7436
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  • 107
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    Prediction method for two-dimensional aerodynamic losses of cooled vanes using integral boundary-layer parameters (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A generalized analysis to predict the two-dimensional aerodynamic losses of film-cooled vanes by using integral boundary-layer parameters is presented. Heat-transfer and trailing-edge injection effects are included in the method. An approximate solution of the generalized equations is also included to show more clearly the effect of the different boundary-layer and cooling parameters on the losses. The analytical predictions agree well with the experimental results, indicating that available boundary-layer calculations for cooled vanes are of sufficient accuracy to use in the prediction method.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1623 , E-076
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  • 108
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    Application of numerical optimization to the design of wings with specified pressure distributions (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A practical procedure for the optimum design of transonic wings is demonstrated. The procedure uses an optimization program based on the method of feasible directions coupled with an aerodynamic analysis program which solves the three-dimensional potential equation for subsonic through transonic flow. Two new wings for the A-7 aircraft were designed by using the optimization procedure to achieve specified surface pressure distributions. The new wings, along with the existing A-7 wing, were tested in the Ames 11 ft transonic wind tunnel. The experimental data show that all of the performance goals were met. However, comparisons of the wind tunnel results with the theoretical predictions indicate some differences at conditions for which strong shock waves occur.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3238
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  • 109
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    Calculation of two-dimensional inlet flow fields in a supersonic free stream by an implicit marching code with nonorthogonal mesh generation: User's manual (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An implicit, shock-capturing finite-difference code which is used to calculate two-dimensional inlet flow fields in a supersonic free stream is explained. The Euler equations are subjected to general nonorthogonal transformation and a body-fitted coordinate system is employed. The mathematical formulation of the problem is given along with the numerical algorithm. Initial and boundary conditions, numerical stability, program limitations, and accuracy is discussed. An overall program logic as well as instructions for program use and operation are also furnished.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3222 , NEAR-TR-193
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  • 110
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    A direct-inverse technique for low speed high lift airfoil-field analysis (1980)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Keywords: AERODYNAMICS
    Type: NASA-CR-159313
    Format: text
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  • 111
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    General design method for three-dimensional potential flow fields. 1: Theory (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A general design method was developed for steady, three dimensional, potential, incompressible or subsonic-compressible flow. In this design method, the flow field, including the shape of its boundary, was determined for arbitrarily specified, continuous distributions of velocity as a function of arc length along the boundary streamlines. The method applied to the design of both internal and external flow fields, including, in both cases, fields with planar symmetry. The analytic problems associated with stagnation points, closure of bodies in external flow fields, and prediction of turning angles in three dimensional ducts were reviewed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3288
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  • 112
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    Water-tunnel and analytical investigation of the effect of strake design variables on strake vortex breakdown characteristics (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A systematic water-tunnel study was made to determine the vortex breakdown characteristics of 43 strakes. The strakes were mounted on a 1/2-scale model of a Langley Research Center general research fighter fuselage model with a 44deg leading-edge-sweep trapezoidal wing. The analytically designed strake shapes provided examples of the effects of the primary design parameters (size, span, and slenderness) on vortex breakdown characteristics. These effects were analyzed in relation to the respective strake leading-edge suction distributions. Included were examples of the effects of detailed strake planform shaping. It was concluded that, consistent with the design criterion, those strakes with leading-edge suction distributions which increase more rapidly near, and have a higher value at, the spanwise tip of the strake produce a more stable vortex.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1676 , L-13254
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  • 113
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    An approximate factorization solution of the Navier-Stokes equations for transonic flow using body-fitted coordinates with application to NACA 64A010 airfoils (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The implementation of the approximate factorization algorithm and its ability to efficiently and accurately describe transonic flow about an NACA 64A010 airfoil section is examined. The approximate factorization algorithm is developed from the nondimensional, conservative, vectorized Navier-Stokes equations expressed in curvilinear coordinates. Equations of state and transport coefficient relations appropriate to atmospheric air are appended to close the system of partial differential equations. An algebraic turbulence model is also incorporated into the equation set. This algorithm was verified by investigating the flow about an NACA 64A010 airfoil at 0, 2, and 3.5 deg angle of attack for free-stream conditions of 2,000,000 Reynolds number and 0.8 Mach number. Overall results were in good qualitative agreement with wind tunnel data sets. However, while nondimensional times of six were attained, numerical difficulties prevented any case from reaching a true steady state.
    Keywords: AERODYNAMICS
    Type: NASA-CR-163376
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  • 114
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    Phase 1 wind tunnel tests of the J-97 powered, external augmentor V/STOL model (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Test results are presented for a large scale, external augmentor V/STOL model in a 40 ft by 80 ft wind tunnel. The model was powered by a GE J97 engine and featured longitudinal ejectors alongside and external to the fuselage together with an augmentor flap on the low aspect ratio, double-delta wing. A static thrust augmentation ratio of 1.60 was measured for the fuselage augmentor at a nozzle pressure ratio of 3.0 and a nozzle exhaust gas temperature of 700 C. At forward speed the model showed a strong positive lift interference due to the augmentor flap, and a marked absence of negative lift interference due to the fuselage augmentor jet system. The nose-up moment of the fuselage augmentor inlet flow was approximately cancelled by a 60 deg deflection of the augmentor flap. An assessment of the thrust and drag components to allow the prediction of transition performance of aircraft designs based on the present conceptual model was made. Lateral tests showed strong but well ordered effects of power.
    Keywords: AERODYNAMICS
    Type: NASA-CR-152255 , DHC-DND-79-4
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  • 115
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    Calculation of Water Drop Trajectories to and About Arbitrary Three-Dimensional Bodies in Potential Airflow (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Calculations can be performed for any atmospheric conditions and for all water drop sizes, from the smallest cloud droplet to large raindrops. Any subsonic, external, non-lifting flow can be accommodated; flow into, but not through, inlets also can be simulated. Experimental water drop drag relations are used in the water drop equations of motion and effects of gravity settling are included. Seven codes are described: (1) a code used to debug and plot body surface description data; (2) a code that processes the body surface data to yield the potential flow field; (3) a code that computes flow velocities at arrays of points in space; (4) a code that computes water drop trajectories from an array of points in space; (5) a code that computes water drop trajectories and fluxes to arbitrary target points; (6) a code that computes water drop trajectories tangent to the body; and (7) a code that produces stereo pair plots which include both the body and trajectories. Code descriptions include operating instructions, card inputs and printouts for example problems, and listing of the FORTRAN codes. Accuracy of the calculations is discussed, and trajectory calculation results are compared with prior calculations and with experimental data.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3291
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  • 116
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    System for use in conducting wake investigation for a wing in flight (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A system supported by a wing in flight is described which has a reference total pressure port in spaced relation with a wake as the wake is generated by the wing, a reference static pressure port supported in spaced relation with the wake, and a probe adapted to be displaced along an accurate path through the wake including a total pressure port and static pressure ports. A differential pressure transducer and a pressure switching device are interposed between the ports and the transducer is provided for selectively connecting pairs of the ports to the transducer in opposed relation, whereby a single transducer is utilized to obtain differential pressure measurement for the wake with enhanced accuracy.
    Keywords: AERODYNAMICS
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  • 117
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    Three dimensional mean flow and turbulence characteristics of the near wake of a compressor rotor blade (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The investigation was carried out using the rotating hot wire technique. Measurements were taken inside the end wall boundary layer to discern the effect of annulus and hub wall boundary layer, secondary flow, and tip leakage on the wake structure. Static pressure gradients across the wake were measured using a static stagnation pressure probe insensitive to flow direction changes. The axial and the tangential velocity defects, the radial component of velocity, and turbulence intensities were found to be very large as compared to the near and far wake regions. The radial velocities in the trailing edge region exhibited characteristics prevalent in a trailing vortex system. Flow near the blade tips found to be highly complex due to interaction of the end wall boundary layers, secondary flows, and tip leakage flow with the wake. The streamwise curvature was found to be appreciable near the blade trailing edge. Flow properties in the trailing edge region are quite different compared to that in the near and far wake regions with respect to their decay characteristics, similarity, etc. Fourier decomposition of the rotor wake revealed that for a normalized wake only the first three coefficients are dominant.
    Keywords: AERODYNAMICS
    Type: NASA-CR-159518 , PSU-TURBO-R-80-4
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  • 118
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    Experimental studies of scale effects on oscillating airfoils at transonic speeds (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Experimental data are presented on the effect of Reynolds number on unsteady pressures induced by the pitching motion of an oscillating airfoil. Scale effects are discussed with reference to a conventional airfoil (NACA 64A010) and a supercritical airfoil (NLR 7301) at mean-flow conditions that support both weak and strong shock waves. During the experiment the Reynolds number was varied from 3,000,000 to 12,000,000 at a Mach number and incidence necessary to induce the required flow. Both fundamental frequency and complete time history data are presented over the range of reduced frequencies that is important in aeroelastic applications. The experimental data show that viscous effects are important in the case of the supercritical airfoil at all flow conditions and in the case of the conventional airfoil under strong shock-wave conditions. Some frequency-dependent viscous effects were also observed.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81216 , A-8259
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  • 119
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    Effect of Reynolds number on stability characteristics of a cruciform wing-body (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An experimental investigation was conducted to determine the effect of Reynolds number on the stability characteristics of a body with cruciform wings at large angles of attack. Pressure distributions and force and moment data (axial force not measured) are presented for Mach 1.60 and 2.70, Reynolds numbers based on body diameter from approximately 130,000 to 2,800,000, and angles of attack from 0 deg to 50 deg. In general, the data show only small effects of Reynolds number throughout the range of test condition. Also discussed are force balance and pressure data that suggest a direct relationship between wind choking and the onset of a nonlinear stability variaton with angle of attack.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1683 , L-13530
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  • 120
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    A calculation procedure for viscous flow in turbomachines, volume 3 (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A method for analyzing the nonadiabatic viscous flow through turbomachine blade passages was developed. The field analysis is based upon the numerical integration of the full incompressible Navier-Stokes equations, together with the energy equation on the blade-to-blade surface. A FORTRAN IV computer program was written based on this method. The numerical code used to solve the governing equations employs a nonorthogonal boundary fitted coordinate system. The flow may be axial, radial or mixed and there may be a change in stream channel thickness in the through-flow direction. The inputs required for two FORTRAN IV programs are presented. The first program considers laminar flows and the second can handle turbulent flows. Numerical examples are included to illustrate the use of the program, and to show the results that are obtained.
    Keywords: AERODYNAMICS
    Type: NASA-CR-159864
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  • 121
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    Steady, Oscillatory, and Unsteady Subsonic and Supersonic Aerodynamics, production version (SOUSSA-P 1.1). Volume 1: Theoretical manual (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Recent developments of the Green's function method and the computer program SOUSSA (Steady, Oscillatory, and Unsteady Subsonic and Supersonic Aerodynamics) are reviewed and summarized. Applying the Green's function method to the fully unsteady (transient) potential equation yields an integro-differential-delay equation. With spatial discretization by the finite-element method, this equation is approximated by a set of differential-delay equations in time. Time solution by Laplace transform yields a matrix relating the velocity potential to the normal wash. Premultiplying and postmultiplying by the matrices relating generalized forces to the potential and the normal wash to the generalized coordinates one obtains the matrix of the generalized aerodynamic forces. The frequency and mode-shape dependence of this matrix makes the program SOUSSA useful for multiple frequency and repeated mode-shape evaluations.
    Keywords: AERODYNAMICS
    Type: NASA-CR-159130 , ASI-TR-78-45-VOL-1
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  • 122
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    A computer program to generate two-dimensional grids about airfoils and other shapes by the use of Poisson's equation (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A method for generating two dimensional finite difference grids about airfoils and other shapes by the use of the Poisson differential equation is developed. The inhomogeneous terms are automatically chosen such that two important effects are imposed on the grid at both the inner and outer boundaries. The first effect is control of the spacing between mesh points along mesh lines intersecting the boundaries. The second effect is control of the angles with which mesh lines intersect the boundaries. A FORTRAN computer program has been written to use this method. A description of the program, a discussion of the control parameters, and a set of sample cases are included.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81198 , A-8178
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  • 123
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    An interaction algorithm for prediction of mean and fluctuating velocities in two-dimensional aerodynamic wake flows (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A theoretical analysis is presented yielding sets of partial differential equations for determination of turbulent aerodynamic flowfields in the vicinity of an airfoil trailing edge. A four phase interaction algorithm is derived to complete the analysis. Following input, the first computational phase is an elementary viscous corrected two dimensional potential flow solution yielding an estimate of the inviscid-flow induced pressure distribution. Phase C involves solution of the turbulent two dimensional boundary layer equations over the trailing edge, with transition to a two dimensional parabolic Navier-Stokes equation system describing the near-wake merging of the upper and lower surface boundary layers. An iteration provides refinement of the potential flow induced pressure coupling to the viscous flow solutions. The final phase is a complete two dimensional Navier-Stokes analysis of the wake flow in the vicinity of a blunt-bases airfoil. A finite element numerical algorithm is presented which is applicable to solution of all partial differential equation sets of inviscid-viscous aerodynamic interaction algorithm. Numerical results are discussed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3301 , COMCO-80-TR-1.0
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  • 124
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    Exploratory studies of the cruise performance of upper surface blown configurations. Experimental program: Test facilities, model design instrumentation, and lowspeed, high-lift tests (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The model hardware, test facilities and instrumentation utilized in an experimental study of upper surface blown configurations at cruise is described. The high speed (subsonic) experimental work, studying the aerodynamic effects of wing nacelle geometric variations, was conducted around semispan model configurations composed of diversified, interchangeable components. Power simulation was provided by high pressure air ducted through closed forebody nacelles. Nozzle geometry was varied across size, exit aspect ratio, exit position and boattail angle. Three dimensional force and two dimensional pressure measurements were obtained at cruise Mach numbers from 0.5 to 0.8 and at nozzle pressure ratios up to about 3.0. The experimental investigation was supported by an analytical synthesis of the system using a vortex lattice representation with first order power effects. Results are also presented from a compatibility study in which a short haul transport is designed on the basis of the aerodynamic findings in the experimental study as well as acoustical data obtained in a concurrent program. High lift test data are used to substantiate the projected performance of the selected transport design.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3192 , LG77FR0028
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  • 125
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    Development and validation of a combined rotor fuselage induced flow field computational method (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A potential-flow panel method was modified to calculate the effects of a rotor wake on the time-averaged surface pressure and velocity distributions on a helicopter fuselage. The rotor-induced velocities are calculated by using a vortex-tube wake model. The calculated pressure distributions are found to compare well with experimental data obtained from tests of a wind-tunnel model.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1656 , AVRADCOM-TR-80-B-3 , L-13363
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  • 126
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    Transonic Flow Field Analysis for Wing-Fuselage Configurations (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A computational method for simulating the aerodynamics of wing-fuselage configurations at transonic speeds is developed. The finite difference scheme is characterized by a multiple embedded mesh system coupled with a modified or extended small disturbance flow equation. This approach permits a high degree of computational resolution in addition to coordinate system flexibility for treating complex realistic aircraft shapes. To augment the analysis method and permit applications to a wide range of practical engineering design problems, an arbitrary fuselage geometry modeling system is incorporated as well as methodology for computing wing viscous effects. Configuration drag is broken down into its friction, wave, and lift induced components. Typical computed results for isolated bodies, isolated wings, and wing-body combinations are presented. The results are correlated with experimental data. A computer code which employs this methodology is described.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3243
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  • 127
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    Study of experiments on condensation of nitrogen by homogeneous nucleation at states modelling those on the national transonic facility (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A cryogenic wind tunnel is based on the twofold idea of lowering drive power and increasing Reynolds number by operating with nitrogen near its boiling point. There are two possible types of condensation problems involved in this mode of wind tunnel operation. They concern the expansion from the nozzle supply to the test section at relatively low cooling rates, and secondly the expansion around models in the test section. This secondary expansion involves higher cooling rates and shorter time scales. In addition to these two condensation problems it is not certain what purity of nitrogen can be achieved in a large facility. Therefore, one cannot rule out condensation processes other than those of homogeneous nucleation.
    Keywords: AERODYNAMICS
    Type: NASA-CR-163217
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  • 128
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    Further investigation of a finite difference procedure for analyzing the transonic flow about harmonically oscillating airfoils and wings (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Analytical and empirical studies of a finite difference method for the solution of the transonic flow about harmonically oscillating wings and airfoils are presented. The procedure is based on separating the velocity potential into steady and unsteady parts and linearizing the resulting unsteady equations for small disturbances. The steady velocity potential is obtained first from the well-known nonlinear equation for steady transonic flow. The unsteady velocity potential is then obtained from a linear differential equation in complex form with spatially varying coefficients. Since sinusoidal motion is assumed, the unsteady equation is independent of time. An out-of-core direct solution procedure was developed and applied to two-dimensional sections. Results are presented for a section of vanishing thickness in subsonic flow and an NACA 64A006 airfoil in supersonic flow. Good correlation is obtained in the first case at values of Mach number and reduced frequency of direct interest in flutter analyses. Reasonable results are obtained in the second case. Comparisons of two-dimensional finite difference solutions with exact analytic solutions indicate that the accuracy of the difference solution is dependent on the boundary conditions used on the outer boundaries. Homogeneous boundary conditions on the mesh edges that yield complex eigenvalues give the most accurate finite difference solutions. The plane outgoing wave boundary conditions meet these requirements.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3195 , D-48851
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    Simple turbulence models and their application to boundary layer separation (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Measurements in the boundary layer and wake of a stalled airfoil are presented in two coordinate systems, one aligned with the airfoil chord, the other being conventional boundary layer coordinates. The NACA 4412 airfoil is studied at a single angle of attack corresponding to maximum lift, the Reynolds number based on chord being 1.5 x 10 to the 6th power. Turbulent boundary layer separation occurred at the 85 percent chord position. The two-dimensionality of the flow was documented and the momentum integral equation studied to illustrate the importance of turbulence contributions as separation is approached. The assumptions of simple eddy-viscosity and mixing-length turbulence models are checked directly against experiment. Curvature effects are found to be important as separation is approached.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3283
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  • 130
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    A general panel method for the analysis and design of arbitrary configurations in incompressible flows (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A method for solving the linear integral equations of incompressible potential flow in three dimensions is presented. Both analysis (Neumann) and design (Dirichlet) boundary conditions are treated in a unified approach to the general flow problem. The method is an influence coefficient scheme which employs source and doublet panels as boundary surfaces. Curved panels possessing singularity strengths, which vary as polynomials are used, and all influence coefficients are derived in closed form. These and other features combine to produce an efficient scheme which is not only versatile but eminently suited to the practical realities of a user-oriented environment. A wide variety of numerical results demonstrating the method is presented.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3079 , D6-43808
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  • 131
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    A comparison of calculated and experimental lift and pressure distributions for several helicopter rotor sections (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The use of computational techniques in predicting lift coefficients and pressure distributions of two dimenstional airfoil sections was studied. The computer code FL06/IBL was used to solve the compressible, two dimensional flow about four different airfoil sections. The lift coefficients of the airfoils were calculated at various angles of attack at subsonic Mach numbers and compared with experimental data.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81160 , A-8029
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  • 132
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    Wing flapping with minimum energy (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: For slow flapping motions it is found that the minimum energy loss occurs when the vortex wake moves as a rigid surface that rotates about the wing root - a condition analogous to that determined for a slow-turning propeller. The optimum circulation distribution determined by this condition differs from the elliptic distribution, showing a greater concentration of lift toward the tips. It appears that very high propulsive efficiencies are obtained by flapping.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81174 , A-8076
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  • 133
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    Development of panel methods for subsonic analysis and design (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Two computer programs, developed for subsonic inviscid analysis and design are described. The first solves arbitrary mixed analysis design problems for multielement airfoils in two dimensional flow. The second calculates the pressure distribution for arbitrary lifting or nonlifting three dimensional configurations. In each program, inviscid flow is modelled by using distributed source doublet singularities on configuration surface panels. Numerical formulations and representative solutions are presented for the programs.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3234
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  • 134
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    The quasi-vortex-lattice method for wings with edge vortex separation (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The aerodynamic characteristics of wings with leading-edge vortex separation were predicted using a method based on a flow model with free vortex elements which are allowed to merge into a concentrated core. The calculated pressure distribution is more accurate than that predicted by methods with discrete vortex filaments alone. In addition, the computer time is reduced approximately by half.
    Keywords: AERODYNAMICS
    Type: NASA-CR-162530 , CRINC-FRL-385-1
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  • 135
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    A three dimensional vortex wake model for missiles at high angles on attack (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A three dimensional model for the steady flow past missile and aircraft nose shaped bodies is presented based on augmenting a potential solution with a wake composed of vortex filaments. The vortex positions are determined by the requirement that they, in some sense, align with the flow. The aerodynamic loads on the body are compared with experimental values and used to evaluate the model. The vortex positions compare well with flow visualization results for slender bodies at high angles of attack. The approximations in the wake near the body cause peaks in the force distributions more severe than in the measured values. For given vortex strengths and body attachment points multiple steady vortex positions were not found.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3208 , TRW-30584-6003-RU-00
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  • 136
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    A simplified theory of oscillating aerofoils in transonic flow - Review and extension (1980)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A simplified theory of the dynamic motion of aerofoils of finite thickness in transonic flow is presented which excludes the effect of shock waves on the aerofoil itself and, thus, is restricted to free stream Mach numbers equal to unity or above. Numerical examples are analyzed for two-dimensional steady and unsteady (including transient) aerofoil motion and three-dimensional steady and unsteady flow over delta wings. The effects of flow separation and improvements in Bernoulli's equation and the surface boundary condition are briefly discussed.
    Keywords: AERODYNAMICS
    Type: Aeronautical Quarterly; 31; Nov. 198
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  • 137
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    An alternating-direction-implicit algorithm for the unsteady potential equation in conservation form (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An implicit finite difference scheme for an efficient computation of unsteady potential flow about airfoils is presented. The formulation uses density and velocity potential as dependent variables, and is cast in conservation form to assure the theoretically correct determination of shockwave location and speed. To enable boundary conditions to be imposed directly on the airfoil surface, a time varying sheared rectilinear coordinate transformation is employed. Calculated time history solutions on a pulsating airfoil are compared with the results of another unsteady transonic code. It is concluded that the method has excellent numerical stability and gives accurate solutions with sharply resolved shocks.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166152
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  • 138
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    An experimental investigation of three dimensional low speed minimum interference wind tunnel for high lift wings (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: As a means to achieve a minimum interference correction wind tunnel, a partially actively controlled test section was experimentally examined. A jet flapped wing with 0.91 m (36 in) span and R = 4.05 was used as a model to create moderately high lift coefficients. The partially controlled test section was simulated using an insert, a rectangular box 0.96 x 1.44 m (3.14 x 4.71 ft) open on both ends in the direction of the tunnel air flow, placed in the University of Washington Aeronautical Laboratories (UWAL) 2.44 x 3.66 m (8 x 12 ft) wind tunnel. A tail located three chords behind the wing was used to measure the downwash at the tail region. The experimental data indicates that, within the range of momentum coefficient examined, it appears to be unnecessary to actively control all four sides of the test section walls in order to achieve the near interference free flow field environment in a small wind tunnel. The remaining wall interference can be satisfactorily corrected by the vortex lattice method.
    Keywords: AERODYNAMICS
    Type: NASA-CR-164439
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  • 139
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    Data base for the prediction of inlet external drag (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Results are presented from a study to define and evaluate the data base for predicting an airframe/propulsion system interference effect shown to be of considerable importance, inlet external drag. The study is focused on supersonic tactical aircraft with highly integrated jet propulsion systems, although some information is included for supersonic strategic aircraft and for transport aircraft designed for high subsonic or low supersonic cruise. The data base for inlet external drag is considered to consist of the theoretical and empirical prediction methods as well as the experimental data identified in an extensive literature search. The state of the art in the subsonic and transonic speed regimes is evaluated. The experimental data base is organized and presented in a series of tables in which the test article, the quantities measured and the ranges of test conditions covered are described for each set of data; in this way, the breadth of coverage and gaps in the existing experimental data are evident. Prediction methods are categorized by method of solution, type of inlet and speed range to which they apply, major features are given, and their accuracy is assessed by means of comparison to experimental data.
    Keywords: AERODYNAMICS
    Type: AGARD Subsonic(Transonic Configuration Aerodyn.; 23 p
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  • 140
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    Leading-edge deflection optimization for a highly swept arrow wing configuration (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Tests were also conducted to determine the sensitivity of the lateral stability derivative C sub l sub beta to geometric anhedral. The optimized leading edge deflection was developed by aligning the leading edge with the incoming flow along the entire span. Owing to the spanwise variation of upwash, the resulting optimized leading edge was a smooth, continuously warped surface. For the particular configuration studied, levels of leading edge suction on the order of 90 percent were achieved with the smooth, continuously warped leading edge contour. The results of tests conducted to determine the sensitivity of C sub l sub beta to geometric anhedral indicate values of delta C sub l sub beta/delta T which are in reasonable agreement with estimates provided by simple vortex lattice theories.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1777 , L-13820
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  • 141
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    Aerodynamic side-force alleviator means (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An apparatus for alleviating high angle of attack side force on slender pointed cylindrical forebodies such as fighter aircraft, missiles and the like is described. A symmetrical pair of helical separation trips was employed to disrupt the leeside vortices normally attained. The symmetrical pair of trips starts at either a common point or at space points on the upper surface of the forebody and extends along separate helical paths along the circumference of the forebody.
    Keywords: AERODYNAMICS
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  • 142
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    Rotary balance data for a single-engine agricultural airplane configuration for an angle-of-attack range of 8 deg to 90 deg (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Aerodynamic characteristics obtained in a helical flow environment utilizing a rotary balance located in the Langley spin tunnel are presented in plotted form for a 1/10 scale single engine agricultural airplane model. The configurations tested include the basic airplane, various wing leading edge and wing tip devices, elevator, aileron, and rudder control settings, and other modifications. Data are presented without analysis for an angle of attack range of 8 deg to 90 deg, and clockwise and counter-clockwise rotations covering a spin coefficient range from 0 to .9.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3311
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  • 143
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    A computer program for the design and analysis of low-speed airfoils, supplement (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Three new options were incorporated into an existing computer program for the design and analysis of low speed airfoils. These options permit the analysis of airfoils having variable chord (variable geometry), a boundary layer displacement iteration, and the analysis of the effect of single roughness elements. All three options are described in detail and are included in the FORTRAN IV computer program.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81862
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  • 144
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    Some observations on supersonic wing design (1980)
    In:  Other Sources
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    Publication Date: 2019-07-27
    Description: The paper presents a brief review on the development of supersonic wing design. Attention is given to linearized aerodynamic theory, emphasizing equations for drag and ratios of slopes and Mach lines. Diagrams that depict conditions for minimum drag as well as the effects of fore-and-aft dimension of wings and Mach numbers on areas of lateral entrainment are presented.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-3040 , The evolution of aircraft wing design; March 18, 19, 1980; Dayton, OH
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  • 145
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    Reformulation of Possio's kernel with application to unsteady wind tunnel interference (1980)
    In:  Other Sources
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    Publication Date: 2019-06-27
    Description: An efficient method for computing the Possio kernel has remained elusive up to the present time. In this paper the Possio is reformulated so that it can be computed accurately using existing high precision numerical quadrature techniques. Convergence to the correct values is demonstrated and optimization of the integration procedures is discussed. Since more general kernels such as those associated with unsteady flows in ventilated wind tunnels are analytic perturbations of the Possio free air kernel, a more accurate evaluation of their collocation matrices results with an exponential improvement in convergence. An application to predicting frequency response of an airfoil-trailing edge control system in a wind tunnel compared with that in free air is given showing strong interference effects.
    Keywords: AERODYNAMICS
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  • 146
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    Aerodynamic coefficients in generalized unsteady thin airfoil theory (1980)
    In:  Other Sources
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    Publication Date: 2019-06-27
    Description: Two cases are considered: (1) rigid body motion of an airfoil-flap combination consisting of vertical translation of given amplitude, rotation of given amplitude about a specified axis, and rotation of given amplitude of the control surface alone about its hinge; the upwash for this problem is defined mathematically; and (2) sinusoidal gust of given amplitude and wave number, for which the upwash is defined mathematically. Simple universal formulas are presented for the most important aerodynamic coefficients in unsteady thin airfoil theory. The lift and moment induced by a generalized gust are evaluated explicitly in terms of the gust wavelength. Similarly, in the control surface problem, the lift, moment, and hinge moments are given as explicit algebraic functions of hinge location. These results can be used together with any of the standard numerical inversion routines for the elementary loads (pitch and heave).
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 18; July 198
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  • 147
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    Interaction between a normal shock wave and a turbulent boundary layer at high transonic speeds. II - Wall shear stress (1980)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: Asymptotic methods are used to calculate the shear stress at the wall for the interaction between a normal shock wave and a turbulent boundary layer on a flat plate. A mixing length model is used for the eddy viscosity. The shock wave is taken to be strong enough that the sonic line is deep in the boundary layer and the upstream influence is thus very small. It is shown that unlike the result found for laminar flow an asymptotic criterion for separation is not found; however, conditions for incipient separation are computed numerically using the derived solution for the shear stress at the wall. Results are compared with available experimental measurements.
    Keywords: AERODYNAMICS
    Type: Zeitschrift fuer angewandte Mathematik und Physik; 31; Mar. 25
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  • 148
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    Interaction between a normal shock wave and a turbulent boundary layer at high transonic speeds. I - Pressure distribution (1980)
    In:  Other Sources
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    Publication Date: 2019-06-27
    Description: Asymptotic solutions are derived for the pressure distribution in the interaction of a weak normal shock wave with a turbulent boundary layer. The undisturbed boundary layer is characterized by the law of the wall and the law of the wake for compressible flow. In the limiting case considered, for 'high' transonic speeds, the sonic line is very close to the wall. Comparisons with experiment are shown, with corrections included for the effect of longitudinal wall curvature and for the boundary-layer displacement effect in a circular pipe.
    Keywords: AERODYNAMICS
    Type: Zeitschrift fuer angewandte Mathematik und Physik; 31; Mar. 25
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  • 149
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    Analysis of two-dimensional incompressible flows by a subsurface panel method (1980)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: A new approach to panel methods is explored for two-dimensional steady incompressible flows. The method uses linear distributions of sources and vortices on straight-line panels, but satisfies boundary conditions on the actual body surface, at nodes that are also end points of the panels. The result is continuity in body-surface velocity distribution, without recourse to numerical quadrature for the velocity influence coefficients. The method is unusually sensitive to the distribution of the nodes. For example, it almost always fails to give acceptable results when the nodes are distributed randomly. However, the continuity of the velocity distribution makes possible a unique node redistribution scheme, which may be iterated to give accurate results reliably.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 18; May 1980
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  • 150
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    Low-frequency pressure fluctuations in axisymmetric turbulent boundary layers (1980)
    In:  Other Sources
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    Publication Date: 2019-06-27
    Description: Measurements were made of wall pressure fluctuations under a turbulent boundary layer on the fuselage of a sailplane. Experiments with the sailplane offered a noise-free flow with a low free-stream turbulence level. In this environment the wall-pressure spectrum of a turbulent boundary layer with natural transition was found to drop off at low frequencies. Correlations between several wall-mounted microphones revealed that the large-scale motions contribute about 35% to the mean square pressure. Velocity fluctuations at several positions within and outside the boundary layer were measured and correlated with the wall pressure. It seems that the irrotational motions in the turbulent region are primarily responsible for the large-scale wall-pressure fluctuations. A time-lagged conditional correlation of the pressure was introduced to gain further insight into the pressure-producing motions.
    Keywords: AERODYNAMICS
    Type: Journal of Fluid Mechanics; 97; Mar. 25
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  • 151
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    A simple method for converting frequency domain aerodynamics to the time domain (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A simple, direct procedure was developed for converting frequency domain aerodynamics into indicial aerodynamics. The data required for aerodynamic forces in the frequency domain may be obtained from any available (linear) theory. The method retains flexibility for the analyst and is based upon the particular character of the frequency domain results. An evaluation of the method was made for incompressible, subsonic, and transonic two dimensional flows.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81844 , L-13789
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  • 152
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    Rotary balance data for a typical single-engine general aviation design for an angle of attack range of 8 deg to 90 deg. 1: Low wing model C (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Aerodynamic characteristics obtained in a helical flow environment utilizing a rotary balance located in the Langley spin tunnel are presented in plotted form for a 1/6 scale, single engine, low wing, general aviation model (model C). The configurations tested included the basic airplane and control deflections, wing leading edge and fuselage modification devices, tail designs and airplane components. Data are presented without analysis for an angle of attack range of 8 deg to 90 deg and clockwise and counter clockwise rotations covering an omega b/2v range from 0 to .9.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3200
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  • 153
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    WIND: Computer program for calculation of three dimensional potential compressible flow about wind turbine rotor blades (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A computer program is presented which numerically solves an exact, full potential equation (FPE) for three dimensional, steady, inviscid flow through an isolated wind turbine rotor. The program automatically generates a three dimensional, boundary conforming grid and iteratively solves the FPE while fully accounting for both the rotating cascade and Coriolis effects. The numerical techniques incorporated involve rotated, type dependent finite differencing, a finite volume method, artificial viscosity in conservative form, and a successive line overrelaxation combined with the sequential grid refinement procedure to accelerate the iterative convergence rate. Consequently, the WIND program is capable of accurately analyzing incompressible and compressible flows, including those that are locally transonic and terminated by weak shocks. The program can also be used to analyze the flow around isolated aircraft propellers and helicopter rotors in hover as long as the total relative Mach number of the oncoming flow is subsonic.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1729 , E-474
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  • 154
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    A flight investigation of performance and loads for a helicopter with 10-64C main rotor blade sections (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A flight investigation produced data on performance and rotor loads for a teetering rotor, AH-1G helicopter flown with a main rotor that had the NLR-1T airfoil as the blade section contour. The test envelope included hover, forward flight speeds from 34 to 83 m/sec (65 to 162 knots), and collective fixed maneuvers at about 0.25 tip speed ratio. The data set for each test point describes vehicle flight state, control positions, rotor loads, power requirements, and blade motions. Rotor loads are reviewed primarily in terms of peak to peak and harmonic content. Lower frequency components predominated for most loads and generally increased with increased airspeed, but not necessarily with increased maneuver load factor. Detailed data for an advanced airfoil on an AH-1G are presented.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81871 , AVRADCOM-TM-80-B-2
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  • 155
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    Experimental unsteady aerodynamics of conventional and supercritical airfoils (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Experimental data on the unsteady aerodynamics of oscillating airfoils in transonic flow are presented. Two 0.5 m-chord airfoil models - an NACA 64A010 and an NLR 7301 - were tested in the NASA-Ames 11 by 11 foot Transonic Wind Tunnel at Mach numbers to 0.85, at chord Reynolds numbers to 12 million and at mean angles of attack to 4 deg. The airfoils were subjected to both pitching and plunging motions at reduced frequencies to 0.3 (physical frequencies to 53 Hz). The new hardware and the extensive use of computer-experiment integration developed for this test are described. The geometrical configuration of the model and the test arrangement are described in detail. Mean and first harmonic data are presented in both tabular and graphical form to aid in comparisons with other data and with numerical computations.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81221 , A-8294
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  • 156
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    Full-scale wind tunnel-investigation of the Advanced Technology Light Twin-Engine airplane (ATLIT) (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An investigation was conducted to evaluate the aerodynamic performance, stability, and control characteristics of the Advanced Technology Light Twin Engine airplane (ATLIT). Data were measured over an angle of attack range from -4 deg to 20 deg for various angles of sideslip between -5 deg and 15 deg at Reynolds numbers of 0.0000023 and 0.0000035 for various settings of power and flap deflection. Measurements were also made by means of special thrust torque balances to determine the installed propeller characteristics. Part of the investigation was devoted to drag cleanup of the basic airplane and to the evaluation of the effect of winglets on drag and stability.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1591 , L-13135
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  • 157
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    Viscous-inviscid calculations of jet entrainment effects on the subsonic flow over nozzle afterbodies (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A viscous-inviscid interaction model was developed to account for jet entrainment effects in the prediction of the subsonic flow over nozzle afterbodies. The model is based on the concept of a weakly interacting shear layer in which the local streamline deflections due to entrainment are accounted for by a displacement-thickness type of correction to the inviscid plume boundary. The entire flow field is solved in an iterative manner to account for the effects on the inviscid external flow of the turbulent boundary layer, turbulent mixing and chemical reactions in the shear layer, and the inviscid jet exhaust flow. The components of the computational model are described, and numerical results are presented to illustrate the interactive effects of entrainment on the overall flow structure. The validity of the model is assessed by comparisons with data obtained form flow-field measurements on cold-air jet exhausts. Numerical results and experimental data are also given to show the entrainment effects on nozzle boattail drag under various jet exhaust and free-stream flow conditions.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1626 , L-13362
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    Theoretical prediction of nonlinear propagation effects on noise signatures generated by subsonic or supersonic propeller or rotor-blade tips (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The nonlinear propagation equations for sound generated by a constant speed blade tip are presented. Propagation from a subsonic tip is treated as well as the various cases that can occur at supersonic speeds. Some computed examples indicate that the nonlinear theory correlates with experimental results better than linear theory for large amplitude waves. For swept tips that generate a wave with large amplitude leading expansion, the nonlinear theory predicts a cancellation effect that results in a significant reduction of both amplitude and impulse.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1660 , L-13388
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  • 159
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    Experimental Study of Airfoil Trailing Edge Noise: Instrumentation, Methodology and Initial Results (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The mechanisms of aerodynamic noise generation at the trailing edge of an airfoil is investigated. Instrumentation was designed, a miniature semiconductor strain-gauge pressure transducer and associated electronic amplifier circuitry were designed and tested and digital signal analysis techniques applied to gain insight into the relationship between the dynamic pressure close to the trailing edge and the sound in the acoustic far-field. Attempts are made to verify some trailing-edge noise generation characteristics as theoretically predicted by several contemporary acousticians. It is found that the noise detected in the far-field is comprised of the sum of many uncorrelated emissions radiating from the vicinity of the trailing edge. These emissions appear to be the result of acoustic energy radiation which has been converted by the trailing-edge noise mechanism from the dynamic fluid energy of independent streamwise 'strips' of the turbulent boundary layer flow.
    Keywords: AERODYNAMICS
    Type: NASA-CR-163007
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  • 160
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    An experimental evaluation of a helicopter rotor section designed by numerical optimization (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The wind tunnel performance of a 10-percent thick helicopter rotor section design by numerical optimization is presented. The model was tested at Mach number from 0.2 to 0.84 with Reynolds number ranging from 1,900,000 at Mach 0.2 to 4,000,000 at Mach numbers above 0.5. The airfoil section exhibited maximum lift coefficients greater than 1.3 at Mach numbers below 0.45 and a drag divergence Mach number of 0.82 for lift coefficients near 0. A moderate 'drag creep' is observed at low lift coefficients for Mach numbers greater than 0.6.
    Keywords: AERODYNAMICS
    Type: NASA-TM-78622 , AVRADCOM-TR-79-44 , A-7956
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  • 161
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    USSAERO computer program development, versions B and C (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Versions B and C of the unified subsonic and supersonic aerodynamic analysis program, USSAERO, are described. Version B incorporates a new symmetrical singularity method to provide improved surface pressure distributions on wings in subsonic flow. Version C extends the range of application of the program to include the analysis of multiple engine nacelles or finned external stores. In addition, nonlinear compressibility effects in high subsonic and supersonic flows are approximated using a correction based on the local Mach number at panel control points. Several examples are presented comparing the results of these programs with other panel methods and experimental data.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3227
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  • 162
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    Comparison of predicted and experimental real-gas pressure distributions on space shuttle orbiter nose for shuttle entry air data system (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An experimental investigation of inviscid real-gas effects on the pressure distribution along the Space Shuttle Orbiter nose center line up to an angle of attack of 32 deg was performed in support of the Shuttle Entry Air Data System (SEADS). Free-stream velocities from 4.8 to 6.6 kn/s were generated at hypersonic conditions with helium, air, and CO2, resulting in normal-shock density ratios from 3.7 to 18.4. The experimental results for pressure distribution agreed closely with numerical results. Modified Newtonian theory deviates from both experiment and the numerical results as angle of attack increases or shock density ratio decreases. An evaluation of the use of modified Newtonian theory for predicting SEADS pressure distributions in actual flight conditions was made through comparison with numerical predictions.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1627 , L-13341
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  • 163
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    System and method for refurbishing and processing parachutes (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A system and method for refurbishing and processing parachutes is disclosed including an overhead monorail conveyor system on which the parachute is suspended for horizontal conveyance. The parachute is first suspended in a partially opened tented configuration wherein open inspection of the canopy is permitted to remove debris and inspect all areas. Following inspection, the parachute is transported by the monorail conveyor to a washing and drying station with the parachute canopy mounted on the conveyor in a systematic arrangement which permits water and air to pass through the ribbon-like materials of the canopy. Following drying, the chute is conveyed into an interior space where it is finally inspected and removed from the monorial conveyor for folding. The chute is once again mounted on the conveyor and conveyed to a packing area.
    Keywords: AERODYNAMICS
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  • 164
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    A flight investigation of performance and loads for a helicopter with RC-SC2 main-rotor blade sections (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The test envelope included hover, forward-flight speed sweeps from 33 to 74 m/sec (65 to 144 knots), and collective-fixed maneuvers at about 0.25 tip-speed ratio. The data set for each test point describes vehicle flight states, control positions, rotor loads, power requirements and blade motions. Rotor loads were reviewed primarily in terms of peak-to-peak and harmonic content. Lower frequency components predominated for most loads and generally increased with increased airspeed, but not necessarily with increased maneuver load factor.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81898 , AVRADCOM-TM-81-B-1
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  • 165
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    Effects of sidewall geometry on the installed performance of nonaxisymmetric convergent-divergent exhaust nozzles (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The investigation was conducted at static conditions and over a Mach number range from 0.6 to 1.2. Angle of attack was held constant at 0 deg. High pressure air was used to simulate jet exhaust flow at ratios of jet total pressure to free-stream static pressure from 1 (jet off) to approximately 10. Sidewall cutback appears to be a viable way of reducing nozzle weight and cooling requirements without compromising installed performance.
    Keywords: AERODYNAMICS
    Type: NASA-TP-1771 , L-13826
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  • 166
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    Effect of wing location and strakes on stability and control characteristics of a monoplanar circular missile with low-profile tail fins at supersonic speeds (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A wind tunnel test was conducted at Mach numbers from 1.70 to 2.86 to extend the aerodynamic data base for wing tail effect on stability and control characteristics of monoplanar missiles. The results are summarized to show the effects of tail fin dihedral angle, wing location, and nose body strakes. The results indicate that an increase in tail fin dihedral angle produces positive increments in directional stability that allow greater trimmed lift coefficient values (maneuver potential) to be obtained. An increase in wing tail gap for the Mach number range reduces the aerodynamic center travel and produces reductions in directional stability at the lower angles of attack. A change in wing height (vertical location) strongly influences the angle of attack at which pitch up and the most directional stability occur. The addition of strakes to the baseline configuration increases directional stability, which allows a significant increase in stable trimmed maneuver capability. The tail fins of the baseline configuration are effective in producing roll and yaw control that are accompanied by favorable yaw and roll, respectively.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81878 , L-13852
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  • 167
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    The prediction of pressure distributions on an arrow-wing configuration including the effect of camber, twist, and a wing fin (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Wind tunnel tests of an arrow wing body configuration consisting of flat, twisted, and cambered twisted wings were conducted at Mach numbers from 0.40 to 2.50 to provide an experimental data base for comparison with theoretical methods. A variety of leading and trailing edge control surface deflections were included in these tests, and in addition, the cambered twisted wing was tested with an outboard vertical fin to determine its effect on wing and control surface loads. Theory experiment comparisons show that current state of the art linear and nonlinear attached flow methods were adequate at small angles of attack typical of cruise conditions. The incremental effects of outboard fin, wing twist, and wing camber are most accurately predicted by the advanced panel method PANAIR. Results of the advanced panel separated flow method, obtained with an early version of the program, show promise that accurate detailed pressure predictions may soon be possible for an aeroelasticity deformed wing at high angles of attack.
    Keywords: AERODYNAMICS
    Type: Supersonic Cruise Res. 1979, Pt. 1; p 59-115
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  • 168
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    Blade planform for a quiet helicopter (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The effects of blade planform and tip speed on noise and performance for a Hughes 500 C rotor system were studied. A cursory examination of the effects of such planform shapes as regular, inverse, and no taper on the noise and performance of the rotor was conducted. It was found that a constant width wide chord planform at tower tip speed provided the best performance and lowest noise. The tapered planforms had lower performance figures due to the reduced solidity. However, some noise reductions were achieved.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166256 , T-35584
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  • 169
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    Transonic steady- and unsteady-pressure measurements on a high-aspect-ratio supercritical-wing model with oscillating control surfaces (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A supercritical wing with an aspect ratio of 10.76 and with two trailing-edge oscillating control surfaces is described. The semispan wing is instrumented with 252 static orifices and 164 in situ dynamic-pressure gages for studying the effects of control-surface position and motion on steady- and unsteady-pressures at transonic speeds. Results from initial tests conducted in the Langley Transonic Dynamics Tunnel at two Reynolds numbers are presented in tabular form.
    Keywords: AERODYNAMICS
    Type: NASA-TM-81888 , L-13964 , NAS 1.15:81888
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  • 170
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    A vectorized code for calculating laminar and turbulent hypersonic flows about blunt axisymmetric bodies at zero and small angles of attack (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A user's guide is provided for a computer code which calculates the laminar and turbulent hypersonic flows about blunt axisymmetric bodies, such as spherically blunted cones, hyperboloids, etc., at zero and small angles of attack. The code is written in STAR FORTRAN language for the CDC-STAR-100 computer. Time-dependent, viscous-shock-layer-type equations are used to describe the flow field. These equations are solved by an explicit, two-step, time asymptotic, finite-difference method. For the turbulent flow, a two-layer, eddy-viscosity model is used. The code provides complete flow-field properties including shock location, surface pressure distribution, surface heating rates, and skin-friction coefficients. This report contains descriptions of the input and output, the listing of the program, and a sample flow-field solution.
    Keywords: AERODYNAMICS
    Type: NASA-TM-80202 , NAS 1.15:80202
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  • 171
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    Long duct nacelle aerodynamic development for DC-10 derivatives (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The results are presented of a wind tunnel test utilizing a 4.7-percent-scale semispan model of the DC-10 in the Calspan 8-foot transonic wind tunnel. The effect of a revised long-duct nacelle shape on the channel velocities, the incremental drag relative to the baseline long-duct nacelle, and channel velocities for the baseline long-duct nacelle were determined and compared with data obtained at Ames. The baseline and the revised long-duct nacelles are representative of a CF6-50 mixed-flow configuration and were evaluated on a model of a proposed DC-10 stretched-fuselage configuration. The results showed that the revised long-duct nacelle has an appreciable effect on the inboard channel velocities, resulting in an increased channel Mach number. However, the pressure recovery on the nacelle afterbody was about the same for both nacelles. The lift curves for both long-duct nacelle configurations were the same. The channel pressures measured at Calspan were in good agreement with those measured at Ames for the baseline long-duct nacelle. The incremental drag for the revised nacelle was measured as two to four counts (three counts is approximately equal to one percent of the airplane drag) higher than that of the baseline long-duct nacelle.
    Keywords: AERODYNAMICS
    Type: NASA-CR-159271 , NAS 1.26:159271 , ACEE-17-FR-9005
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  • 172
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    Conditional sampling technique to test the applicability of the Taylor hypothesis for the large-scale coherent structures (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: Comparisons of the distributions of large scale structures in turbulent flow with distributions based on time dependent signals from stationary probes and the Taylor hypothesis are presented. The study investigated an area in the near field of a 7.62 cm circular air jet at a Re of 32,000, specifically having coherent structures through small-amplitude controlled excitation and stable vortex pairing in the jet column mode. Hot-wire and X-wire anemometry were employed to establish phase averaged spatial distributions of longitudinal and lateral velocities, coherent Reynolds stress and vorticity, background turbulent intensities, streamlines and pseudo-stream functions. The Taylor hypothesis was used to calculate spatial distributions of the phase-averaged properties, with results indicating that the usage of the local time-average velocity or streamwise velocity produces large distortions.
    Keywords: AERODYNAMICS
    Type: In: EUROMECH 132; Colloquium on Hot-Wire, Hot-Film Anemometry and Conditional Measurements; Jul 02, 1980 - Jul 04, 1980; Rhone; France
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  • 173
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    Large-eddy breakup scheme for turbulent viscous drag reduction (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: A concept for reducing turbulent skin friction drag by altering/controlling the large coherent eddy structures within the turbulent boundary layer is proposed. Results of an ongoing experimental and numerical investigation to develop large-eddy breakup devices (LEBU devices) are presented and indicate that the average skin friction drag downstream of the LEBU devices is reduced by up to 24% compared to 'undisturbed' flat plate levels; device drag requires further reduction before net drag reductions can be realized. Future work is discussed and will focus on reducing device drag by taking advantage of the unsteady 'freestream' ahead of the LEBU devices.
    Keywords: AERODYNAMICS
    Type: Symposium on Viscous flow drag reduction; Nov 07, 1979 - Nov 08, 1979; Dallas, TX
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  • 174
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    Computation of supersonic turbulent flows over an inclined ogive-cylinder-flare (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: A supersonic turbulent flow over an ogive-cylinder-flare has been solved numerically. Initially, the parabolized Navier-Stokes equations are solved for the ogive cylinder back to a location upstream of the shock-wave and boundary-layer interaction. Then, the time-dependent Navier-Stokes equations with a thin-layer approximation are solved for the remaining cylinder-flare portion. Results for a Mach number of 2.9 and a unit Reynolds number of 11.42 x 10 to the 6th/m are obtained for angles of attack alpha = 0, 4, and 8 deg. Good agreement has been found between computed and experimental results of the surface pressure on the ogive-cylinder portion, and for the interaction region at alpha = 0 and 4 deg. The role of circumferential communication in a three-dimensional shock-wave and boundary-layer interaction flow field is discussed.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-1410 , Fluid and Plasma Dynamics Conference; Jul 14, 1980 - Jul 16, 1980; Snowmass, CO
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  • 175
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    Nonreflecting far-field boundary conditions for unsteady transonic flow computation (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: The approximate nonreflecting far-field boundary condition, as proposed by Engquist and Majda, is implemented in the computer code LTRAN2. This code solves the implicit finite-difference representation of the small disturbance equations for unsteady transonic flows about airfoils. The nonreflecting boundary condition and the description of the algorithm for implementing these conditions in LTRAN2 are discussed. Various cases are computed and compared with results from the older, more conventional procedures. One concludes that the nonreflecting far-field boundary approximation allows the far-field boundary to be located closer to the airfoil; this permits a decrease in the computer time required to obtain the solution through the use of fewer mesh points.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-1393 , Fluid and Plasma Dynamics Conference; Jul 14, 1980 - Jul 16, 1980; Snowmass, CO
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  • 176
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    Investigation of transonic flow in a cascade using an adaptive mesh (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: The solution of two-dimensional full potential equation for the analysis of steady transonic flow through cascades is investigated. Finite element method is employed in the analysis. Accuracy and efficiency of the obtained numerical solutions are discussed in terms of the employed computational grid. Accurate modeling of subsonic and supersonic flow regions together with the shock is discussed. The choice of artificial viscosity and relaxation factors are examined and related to the design of a computational grid. Shock capturing and shock fitting procedures are compared for improved accuracy and efficiency. Numerical results include cascades of Gostelow and NACA 0012 airfoils.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-1430 , Fluid and Plasma Dynamics Conference; Jul 14, 1980 - Jul 16, 1980; Snowmass, CO
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  • 177
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    Prediction and measurement of turbulent aerodynamic trailing edge flows (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: A viscous-inviscid interaction algorithm is developed for prediction of two-dimensional mean and fluctuating velocity distributions in the wake immediately downstream of an airfoil trailing edge. A composite pressure field is defined, and a Poisson equation solved for transverse pressure variations. A parabolized form of the time-averaged steady Navier-Stokes equations are solved in conjunction with a viscous-augmented two-dimensional inviscid potential flow analysis. A tensor constitutive equation is employed to predict Reynolds stress distributions from solutions of a turbulence kinetic energy two equation closure model. Numerical predictions compared favorably with detailed experimental data for mean and fluctuating velocities, and Reynolds shear stress distributions, in the trailing edge region of a NACA 63-012 airfoil.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-1395 , Fluid and Plasma Dynamics Conference; Jul 14, 1980 - Jul 16, 1980; Snowmass, CO
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  • 178
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    Analysis of nonlifting and lifting airfoils in transonic flow by parametric differentiation (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: A method is presented which allows one to solve nonlinear transonic flow problems by analyzing a sequence of linear equations. The small disturbance formulation of steady transonic flow over airfoils is linearized by considering the perturbations due to small changes in airfoil thickness ratio and angle of attack. Repeatedly incrementing those parameters results in a series of nonlinear solutions and cumulatively determines the effects of large changes in airfoil geometry. Successive line overrelaxation is used to solve the associated linear equations and is coupled with predictor-corrector methods to yield series of nonlinear solutions. Computed pressure distributions on biconvex airfoils show good agreement with experimental data and other transonic prediction methods. Possible extensions to unsteady and/or three-dimensional transonic flow problems are briefly discussed.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-1394 , Fluid and Plasma Dynamics Conference; Jul 14, 1980 - Jul 16, 1980; Snowmass, CO
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  • 179
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    Complete viscous flowfield solutions about a blunt parabolic body in a supersonic stream (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: A hybrid computational technique which splits the flowfield into inviscid and viscous regions is used to investigate the complete flowfield about axisymmetric parabolic blunt bodies in a supersonic stream. The solutions are carried out on the CDC CYBER-203 computer which, with its extensive memory, allows for the use of a large number of finite-difference mesh points, allowing resolution of important flowfield features. A range of freestream Mach number of 2-5 and a range of Re number based on nose radius of 500-125,000 was run for a sonic corner body. Contour plots of density, pressure, and Mach number, velocity vector plots, and surface distributions of pressure, heat transfer, and shear stress are presented. Also, correlations of the downstream extent of the base recirculation region with Re number based on nose radius are given.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-1351 , Fluid and Plasma Dynamics Conference; Jul 14, 1980 - Jul 16, 1980; Snowmass, CO
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  • 180
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    Zero-length, slotted-lip inlet for subsonic military aircraft (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: Zero-length, slotted-lip inlet performance and associated fan blade stresses were determined during model tests using a 20-inch diameter fan simulator in the NASA-LeRC 9- by 15-foot low-speed wind tunnel. The model configuration variables consisted of inlet contraction ratio, slot width, circumferential extent of slot fillers, and length of a constant area section between the inlet throat and fan face. Inlet configurations having contraction ratios of 1.2 and 1.3 satisfied all critical low-speed inlet operating requirements for a fixed horizontal nacelle and tilt-nacelle-type subsonic V/STOL aircraft, respectively. Relative to a conventional axisymmetric tilt-nacelle inlet, the zero-length, slotted-lip inlet has a 27-percent smaller inlet lip contraction ratio, an 83-percent shorter total length, and a 5-percent smaller maximum cowl diameter.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-1245 , Joint Propulsion Conference; Jun 30, 1980 - Jul 02, 1980; Hartford, CT
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  • 181
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    Study of cooling air inlet and exit geometries for horizontally opposed piston aircraft engines (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: A semispan wing and nacelle of a typical general aviation twin-engine aircraft was tested to evaluate the cooling capability and drag of several nacelle shapes; the nacelle shapes included cooling air inlet and exit variations. The tests were conducted in the Ames Research Center's 40- by 80-Foot Wind Tunnel. It was found that the cooling air inlet geometry of opposed piston engine installations has a major effect on inlet pressure recovery, but only a minor effect on drag. Exit location showed a large effect on drag, especially for those locations on the sides of the nacelle where the suction characteristics were based on interaction with the wing surface pressures.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-1242 , Joint Propulsion Conference; Jun 30, 1980 - Jul 02, 1980; Hartford, CT
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  • 182
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    An experimental investigation of endwall profiling in a turbine vane cascade (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: Measurements of surface static pressures, flow total pressure loss, and exit air angle were obtained for two linear cascades to establish the effects of endwall profiling. Testing was conducted at an isentropic exit Mach number of 0.85. One cascade was fabricated with planar endwalls while the other had one planar and one profiled endwall. Both cascades utilized the same high pressure turbine inlet guide vane section. It was found that in terms of full passage loss the profiled endwall cascade has the superior performance. The secondary loss results obtained are reasonably well predicted by correlations developed from incompressible flow testing of similar configurations. Inviscid flow and boundary layer calculations are compared with the test data, and overall, the agreement is found to be good. Use of the results for design purposes is briefly discussed.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-1089 , Joint Propulsion Conference; Jun 30, 1980 - Jul 02, 1980; Hartford, CT
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  • 183
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    Comparison between optical measurements and a numerical solution of the flow field within a transonic axial-flow compressor rotor (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: A comparison between numerical and experimental results is presented for the flowfield within a transonic axial-flow compressor rotor. The rotor was tested at design speed and a wide open throttle discharge condition. The relative tip Mach number was 1.4. A laser anemometer system was used to measure velocity and flow angle upstream, within, and downstream of the rotor. A holographic interferometer was used to visualize the rotor shock system near the tip. The computational procedure solves the full three-dimensional Euler equations using a time-marching technique. Shock location and shape determined from the two optical systems are compared. Calculated relative Mach number and flow angle contours, shock locations, and shock strength are compared to values measured with the laser anemometer.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-1078 , Joint Propulsion Conference; Jun 30, 1980 - Jul 02, 1980; Hartford, CT
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  • 184
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    Modern fluid dynamics of subsonic and transonic flight (1980)
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    Publication Date: 2019-07-13
    Description: The paper discusses a number of factors, termed research drivers, which are expected to provide much of the stimulus for research in the subsonic and transonic flight regimes in the coming decade. The research drivers discussed comprise the need for energy efficiency, new and improved facilities, better instrumentation, more capable and efficient computers, theoretical methodology refinements, increased use of optimization techniques, and military requirements. Illustrations of advances in aircraft aerodynamics at subsonic and transonic speeds are presented, along with a discussion of future research opportunities and trends. Particular attention is given to airfoil and basic fluids research designed to reduce skin-friction drag.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-0861 , International Meeting and Technical Display on Global Technology 2000; May 06, 1980 - May 08, 1980; Baltimore, MD
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  • 185
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    Control of forebody vortex orientation to alleviate side forces (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: The paper deals with the salient phenomena of three-dimensional symmetric and asymmetric separated flows about typical forbodies at high angles of attack. Particular consideration is given to pressure, forces, and laser vapor screen measurements carried out on a 5-deg semiangle cone in a Mach 0.6 flow under turbulent conditions and supportive tests using a 16-deg semiangle tangent ogive.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-0183 , Aerospace Sciences Meeting; Jan 14, 1980 - Jan 16, 1980; Pasadena, CA
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  • 186
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    Computation of three-dimensional viscous supersonic flow in inlets (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: A new approach has been developed for the computation of the three-dimensional viscous supersonic flow with embedded subsonic regions adjacent to solid boundaries and is applied to a mixed-compression supersonic inlet typical of current designs. The approach uses a reduced form of the three-dimensional Navier-Stokes equations so that the resultant equations can be treated as an initial boundary value problem and thus be solved by non-iterative forward marching in space. The numerical procedure utilizes an efficient consistently-split linearized block implicit technique to solve the finite difference analogues to the set of governing partial differential equations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-0194 , Aerospace Sciences Meeting; Jan 14, 1980 - Jan 16, 1980; Pasadena, CA
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  • 187
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    Influence of nonequilibrium radiation on heating of an ablating Jovian entry probe (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: The influence of non-local thermodynamic equilibrium (NLTE) radiative transfer on the entire shock-layer flow phenomena around a Jovian entry body is investigated. The flow in the shock layer is assumed to be viscous, axisymmetric, laminar, and in chemical equilibrium. The entry body considered is a 35-deg hyperboloid and the results have been obtained for the peak heating entry conditions. The results indicate that the radiative heating of the entry body is significantly higher under NLTE conditions.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-0356 , Aerospace Sciences Meeting; Jan 14, 1980 - Jan 16, 1980; Pasadena, CA
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  • 188
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    Experimental investigation of the asymmetric body vortex wake (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: An experimental investigation of the asymmetric body vortex wake of a circular cylinder in high subsonic flow is presented. Laser velocimeter, force and moment, and surface hot wire measurements were obtained for a freestream Mach number of 0.6 and Reynolds number (based on body diameter) of 0.62 x 10 to the 6th. Two component laser velocimeter measurements were made at three body cross-flow planes, x/d = 4, 8, and 12, and angles of attack of 25, 35, and 45 deg. Laser vapor screen photographs were also obtained at these body stations and angles of attack. Surface hot wire measurements were used to determine if any vortex switching occurred at various angles of attack of the body. The laser velocimeter measurements are related to the vapor screen photographs and side force measurements. These results show that more than one asymmetric body vortex wake configuration can exist for the same angle of attack and body roll angle.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-0174 , Aerospace Sciences Meeting; Jan 14, 1980 - Jan 16, 1980; Pasadena, CA
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  • 189
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    Implicit computations of unsteady transonic flow governed by the full-potential equation in conservation form (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: An alternating-direction implicit algorithm is presented for solving the conservative, full-potential equation for unsteady, transonic flow. A new development is the time-linearization of the density function. This linearization reduces the solution process from one of solving a system of two equations at each mesh point to one of solving a single equation. Two sample cases are computed. First, a one-dimensional traveling shock wave is computed and compared with the analytic solution. Second, a two-dimensional case is computed of a flow field that results from a thickening and subsequently thinning airfoil. The resulting flow field, which includes a traveling shock wave, is compared to the flow field obtained from the low-frequency, small-disturbance, transonic equation.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-0150 , Aerospace Sciences Meeting; Jan 14, 1980 - Jan 16, 1980; Pasadena, CA
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  • 190
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    Simulated transonic flows for aircraft with nacelles, pylons, and winglets (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: A computational method which simulates transonic flow about wing-fuselage configurations has been extended to include the treatment of multiple body and non-planar wing surfaces. The finite difference relaxation scheme is characterized by a modified small disturbance flow equation and multiple embedded grid system. Wing-body combinations with as many as four nacelles/pods, four pylons, and wing-tip-mounted winglets can be analyzed. A scheme for modeling inlet spillage and engine exhaust interference effects has been included. Computed results are correlated with experimental data for three transport configurations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-0130 , Aerospace Sciences Meeting; Jan 14, 1980 - Jan 16, 1980; Pasadena, CA
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  • 191
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    Calculation of inlet flow fields by an implicit technique (1980)
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    Publication Date: 2019-07-13
    Description: Two-dimensional inlet flow fields in a supersonic free stream are calculated by an implicit, shock-capturing, finite-difference method. The Euler equations are subjected to a general curvilinear transformation and a body-fitted coordinate system is employed. The method is used to solve supercritical, critical, and subcritical flow fields which are simulated by prescribing appropriate conditions at the inlet outflow boundary. Results are presented for a drooped-cowl inlet.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-0031 , Aerospace Sciences Meeting; Jan 14, 1980 - Jan 16, 1980; Pasadena, CA
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  • 192
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    Flight-measured effects of boattail angle and Mach number on the nozzle afterbody flow of a twin-jet fighter (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: The paper presents the flight-measured nozzle afterbody surface pressures and engine exhaust nozzle pressure-area integrated axial force coefficients on a twin-jet fighter for varying boattail angles. The objective of the tests was to contribute to a full-scale flight data base applicable to the nozzle afterbody drag of advanced tactical fighter concepts. The data were acquired during the NASA F-15 Propulsion/Airframe Interactions Flight Research Program. Nozzle boattail angles from 7.7 deg to 18.1 deg were investigated. Results are presented for cruise angle of attack at Mach numbers from 0.6 to 2.0 at altitudes from 20,000 to 45,000 feet. The data show the nozle axial force coefficients to be a strong function of nozzle boattail angle and Mach number.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-0110 , Aerospace Sciences Meeting; Jan 14, 1980 - Jan 16, 1980; Pasadena, CA
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  • 193
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    Experimental and analytical comparison of aerodynamic characteristics of a forward-control missile (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: Comparisons of analytical and experimental aerodynamic data for canard controlled missile configurations are presented. Recently, techniques to estimate the longitudinal, directional and lateral aerodynamic characteristics for cruciform missiles have been developed. Nielsen Engineering and Research, Inc. (NEAR, Inc.), supported by various governmental agencies, has been the originator of many of these new computational techniques. Two of these are major computer programs currently being implemented by several research organizations. Predicted data from these two programs are compared with experimental data recently obtained at the NASA Langley Research Center Unitary Plan wind tunnel facility. Comparisons cover the supersonic Mach number regime of 1.60 to 3.50, angles-of-attack from 0 to 20 degrees and roll angles of 0, 26.57 and 45 degrees. Major emphasis is on the roll characteristics due to aileron with limited longitudinal and directional characteristics addressed.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-0374 , Aerospace Sciences Meeting; Jan 14, 1980 - Jan 16, 1980; Pasadena, CA
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  • 194
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    Transonic swept-wing analysis using asymptotic and other numerical methods (1980)
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    Publication Date: 2019-07-13
    Description: The paper presents asymptotic methods for high-aspect-ratio wings in transonic flow developed for straight unyawed wings and for oblique wings. They show that the three-dimensional mixed-flow calculations may be reduced to solving a set of two-dimensional problems at each span station; the development of this theory and the related computational studies are reviewed. Differences between the piloted (oblique) wing, the swept-back wing, and the swept-forward-wing in the induced upwash are discussed; examples of similarity solutions are demonstrated for high subcritical and slightly supercritical component flows, and comparisons made with relaxation solutions of a full potential equation. The examples include oblique and symmetric swept wings, and the adequacy of the existing full-potential computer code is examined.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-0342 , Aerospace Sciences Meeting; Jan 14, 1980 - Jan 16, 1980; Pasadena, CA
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  • 195
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    Blade loading and spanwise effects on the near and far wake characteristics of a compressor rotor blade (1980)
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    Publication Date: 2019-07-13
    Description: An experimental investigation of the blade loading and spanwise effects on the rotor wake is presented. The investigation was limited to a study of a low subsonic and incompressible wake flow found downstream of a lightly loaded rotor. Measurements were made with a tri-axial hot wire probe mounted in the stationary frame of reference at six radial and nine axial positions. At each measurement location, the rotor was run at different operating conditions to discern the effects of blade loading on the wake. Near and far wake measurements are given, including mean velocity and turbulence intensity characteristics. The loading and spanwise effects on rotor wake characteristics were found to be substantial.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-0201 , Aerospace Sciences Meeting; Jan 14, 1980 - Jan 16, 1980; Pasadena, CA
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  • 196
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    Boundary-layer transition on a 10-deg cone - Wind tunnel/flight correlation (1980)
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    Publication Date: 2019-07-13
    Description: Boundary-layer transition location measurements were made on a 10-deg sharp cone in 23 wind tunnels of the US and Europe and in flight. The data were acquired at subsonic, transonic, and supersonic Mach numbers over a range of unit Reynolds numbers to obtain an improved understanding of wind tunnel flow quality influence. Cone surface microphone measurements showed Tollmien-Schlichting waves present. Transition location defined by pitot probe measurements showed transition Reynolds number to be correlatable to cone surface disturbance amplitude within + or - 20 percent for the majority of tunnel and flight data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-0154 , Aerospace Sciences Meeting; Jan 14, 1980 - Jan 16, 1980; Pasadena, CA
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  • 197
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    Propeller slipstream/wing interaction in the transonic regime (1980)
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    Publication Date: 2019-07-13
    Description: An inviscid model for the interaction between a thin wing and a nearly uniform propeller slipstream is presented. The model allows the perturbation velocities due to the interaction to be potential although the undisturbed slipstream velocity is rotational. A finite difference scheme is used to solve the governing equation. Numerical examples indicate that the slipstream has a strong effect on the aerodynamic properties of the wing section within the slipstream and lesser effects elsewhere. The slipstream swirling motion strongly affects the wing load distribution, however, its effect on the wing's total lift and wave drag is small. The axial velocity increment in the slipstream has a small effect on the wing lift, however, it causes a large increase in wave drag.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-0125 , Aerospace Sciences Meeting; Jan 14, 1980 - Jan 16, 1980; Pasadena, CA
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  • 198
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    Forebody and base region real-gas flow in severe planetary entry by a factored implicit numerical method. I - Computational fluid dynamics (1980)
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    Publication Date: 2019-07-13
    Description: A new code for the simulation of full (forebody and base region) flowfields about bluff bodies in the hypersonic regime of severe planetary entry is described. The present 'maximally conservative, maximally differenced' formulation of the unsteady compressible Navier-Stokes equations for 2-D axisymmetric 3-D flow is contrasted for stability with previous formulations of Viviand, Kutler, et al, and Thomas and Lombard. Discrete metric relations peculiar to the axisymmetric finite volume formulation are presented along with a general discussion of their relations to and consequences of failure to close computational cells. A computational mesh of curvilinear coordinate topology singular in the flow regime is presented that permits aligned capturing of the major physical features of the complex flowfield.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-0065 , Aerospace Sciences Meeting; Jan 14, 1980 - Jan 16, 1980; Pasadena, CA
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  • 199
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    Noise generation by a lifting wing/flap combination at Reynolds numbers to 2.8 x 10 to the 6th (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: Measurements relating to the noise source location and intensity within various frequency bands were made for an 0.75 m-chord wing/flap model installed in the Ames 7 x 10-foot wind tunnel. A directional microphone system, located outside the open-wall tunnel was scanned in a two-dimensional array of aiming points about the positive-pressure side of the model to determine the principal locations of noise production, and the intensity of each of these. It was found for the case of the flaps being differentially deflected (0 deg, 35 deg) at the half-span station that noise production was concentrated in the immediate region of the resultant surface discontinuity. For equal deflection of the halves (0 deg, 0 deg or 35 deg, 35 deg), noise was produced uniformly along the length of the gap between the wing and the flap. Simulated flap-mounting brackets generated considerable noise in certain cases, but reduced the noise in others. Trailing edge noise did not appear to be important in comparison with other sources.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-0035 , Aerospace Sciences Meeting; Jan 14, 1980 - Jan 16, 1980; Pasadena, CA
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  • 200
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    An analytical and experimental study of a short S-shaped subsonic diffuser of a supersonic inlet (1980)
    In:  Other Sources
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    Publication Date: 2019-07-13
    Description: An experimental investigation of a subscale HiMat forebody and inlet was conducted over a range of Mach numbers to 1.4. The inlet exhibited a transitory separation within the diffuser but steady state data indicated reattachment at the diffuser exit. A finite difference procedure for turbulent compressible flow in axisymmetric ducts was used to successfully model the HiMAT duct flow. The analysis technique was further used to estimate the initiation of separation and delineate the steady and unsteady flow regimes in similar S-shaped ducts.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 80-0386 , Aerospace Sciences Meeting; Jan 14, 1980 - Jan 16, 1980; Pasadena, CA
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