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  • Life and Medical Sciences  (1,428)
  • Cell & Developmental Biology  (1,172)
  • AERODYNAMICS  (851)
  • 2020-2020
  • 1980-1984  (2,004)
  • 1955-1959  (150)
  • 1935-1939  (125)
  • 1925-1929
  • 1983  (1,176)
  • 1980  (828)
  • 1958  (150)
  • 1939  (125)
Collection
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  • 2020-2020
  • 1980-1984  (2,004)
  • 1955-1959  (150)
  • 1935-1939  (125)
  • 1925-1929
Year
  • 101
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    Experimental and Theoretical Determination of Forces and Moments on a Store and on a Store- Pylon Combination Mounted on a 45 Deg Swept- Wing-Fuselage Configuration at a Mach Number of 1.61 (1958)
    In:  CASI
    Details
    Publication Date: 2019-05-24
    Description: Forces and moments of store-pylon combination mounting on swept wing-fuselage configuration in supersonic pressure tunnel
    Keywords: AERODYNAMICS
    Type: NACA-RM-L57K18
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  • 102
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    Performance of an isentropic, all-internal- contraction, axisymmetric inlet designed for mach 2.50 (1958)
    In:  CASI
    Details
    Publication Date: 2019-05-23
    Description: Performance of internal contraction, axisymmetric inlet with isentropic compression surfaces on cowl and centerbody at Mach 2.0 to 2.7
    Keywords: AERODYNAMICS
    Type: NACA-RM-E58E16
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  • 103
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    Investigation of Inlet Control Parameters for an External-internal-compression Inlet from Mach 2.1 to 3.0 (1958)
    In:  CASI
    Details
    Publication Date: 2019-05-23
    Description: Investigation of the control parameters of an external-internal compression inlet indicates that the cowl-lip shock provides a signal to position the spike and to start the inlet over a Mach number range from 2.1 to 3.0. Use of a single fixed probe position to control the spike over the range of conditions resulted in a 3.7-count loss in total-pressure recovery at Mach 3.0 and 0 deg angle of attack. Three separate shock-sensing-probe positions were required to set the spike for peak recovery from Mach 2.1 to 3.0 and angles of attack from 0 deg to 6 deg. When the inlet was unstarted, an erroneous signal was obtained from the normal-shock control through most of the starting cycle that prevented the inlet from starting. Therefore, it was necessary to over-ride the normal-shock control signal and not allow the control to position the terminal shock until the spike was positioned.
    Keywords: AERODYNAMICS
    Type: NACA-RM-E58G08
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  • 104
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    Exploratory wind-tunnel investigation to determine the lift effects of blowing over flaps from nacelles mounted above the wing (1958)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Keywords: AERODYNAMICS
    Type: NACA-TN-4298
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  • 105
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    Full potential aerodynamic prediction capability for complex configurations (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center, Theoretical Aerodynamics Contractors' Workshop, Volume 2; p 495-53
    Format: text
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  • 106
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    Transonic wing-body flow simulations. Upper/lower canted winglets and arbitrary planform pylons (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Keywords: AERODYNAMICS
    Type: NASA. Langley Research Center, Theoretical Aerodynamics Contractors' Workshop, Volume 2; p 407-43
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  • 107
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    Effect of surface waviness on a supercritical laminar-flow-control airfoil (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Calculations were made of the effects of surface waviness on the external pressure of a supercritical airfoil at design conditions. Wave parameters varied include amplitude, wavelength, phase, and number of cycles. Effects of single and multiple waves are calculated at various chordwise locations. General trends of surface waviness effects on pressure distribution are determined and these solutions are reported. Contour deviations are imposed on the upper surface of the airfoil. Results are presented in a manner designed to facilitate ready comparison with the ideal contour static pressure distribution.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85705 , NAS 1.15:85705
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  • 108
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    Computer programs for smoothing and scaling airfoil coordinates (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Detailed descriptions are given of the theoretical methods and associated computer codes of a program to smooth and a program to scale arbitrary airfoil coordinates. The smoothing program utilizes both least-squares polynomial and least-squares cubic spline techniques to smooth interatively the second derivatives of the y-axis airfoil coordinates with respect to a transformed x-axis system which unwraps the airfoil and stretches the nose and trailing-edge regions. The corresponding smooth airfoil coordinates are then determined by solving a tridiagonal matrix of simultaneous cubic-spline equations relating the y-axis coordinates and their corresponding second derivatives. A technique for computing the camber and thickness distribution of the smoothed airfoil is also discussed. The scaling program can then be used to scale the thickness distribution generated by the smoothing program to a specific maximum thickness which is then combined with the camber distribution to obtain the final scaled airfoil contour. Computer listings of the smoothing and scaling programs are included.
    Keywords: AERODYNAMICS
    Type: NASA-TM-84666 , NAS 1.15:84666
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  • 109
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    Low-speed tests of a high-aspect-ratio, supercritical-wing transport model equipped with a high-lift flap system in the Langley 4- by 7-meter and Ames 12-foot pressure tunnels (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The Ames 12-Foot Pressure Tunnel was used to determine the effects of Reynolds number on the static longitudinal aerodynamic characteristics of an advanced, high-aspect-ratio, supercritical wing transport model equipped with a full span, leading edge slat and part span, double slotted, trailing edge flaps. The model had a wing span of 7.5 ft and was tested through a free stream Reynolds number range from 1.3 to 6.0 x 10 to 6th power per foot at a Mach number of 0.20. Prior to the Ames tests, an investigation was also conducted in the Langley 4 by 7 Meter Tunnel at a Reynolds number of 1.3 x 10 to 6th power per foot with the model mounted on an Ames strut support system and on the Langley sting support system to determine strut interference corrections. The data obtained from the Langley tests were also used to compare the aerodynamic charactertistics of the rather stiff, 7.5-ft-span steel wing model tested during this investigation and the larger, and rather flexible, 12-ft-span aluminum-wing model tested during a previous investigation. During the tests in both the Langley and Ames tunnels, the model was tested with six basic wing configurations: (1) cruise; (2) climb (slats only extended); (3) 15 deg take-off flaps; (4) 30 deg take-off flaps; (5) 45 deg landing flaps; and (6) 60 deg landing flaps.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2097 , L-15484 , NAS 1.60:2097
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  • 110
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    Computer program for plotting and fairing wind-tunnel data (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A detailed description of the Langley computer program PLOTWD which plots and fairs experimental wind-tunnel data is presented. The program was written for use primarily on the Langley CDC computer and CALCOMP plotters. The fundamental operating features of the program are that the input data are read and written to a random-access file for use during program execution, that the data for a selected run can be sorted and edited to delete duplicate points, and that the data can be plotted and faired using tension splines, least-squares polynomial, or least-squares cubic-spline curves. The most noteworthy feature of the program is the simplicity of the user-supplied input requirements. Several subroutines are also included that can be used to draw grid lines, zero lines, axis scale values and lables, and legends. A detailed description of the program operational features and each sub-program are presented. The general application of the program is also discussed together with the input and output for two typical plot types. A listing of the program code, user-guide, and output description are presented in appendices. The program has been in use at Langley for several years and has proven to be both easy to use and versatile.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85648 , NAS 1.15:85648
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  • 111
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    Additional flow field studies of the GA(W)-1 airfoil with 30-percent chord Fowler flap including slot-gap variations and cove shape modifications (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Experimental measurements were made to determine the effects of slot gap opening and flap cove shape on flap and airfoil flow fields. Test model was the GA(W)-1 airfoil with 0.30c Fowler flap deflected 35 degrees. Tests were conducted with optimum, wide and narrow gaps, and with three cove shapes. Three test angles were selected, corresponding to pre-stall and post-stall conditions. Reynolds number was 2,200,000 and Mach number was 0.13. Force, surface pressure, total pressure, and split-film turbulence measurements were made. Results were compared with theory for those parameters for which theoretical values were available.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3687 , NAS 1.26:3687 , AR-79-3
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  • 112
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    Inviscid analysis of two supercritical laminar-flow-control airfoils at design and off-design conditions (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Inviscid transonic flow results are provided at design and off design conditions for two supercritical laminar flow control airfoils. The newer airfoil, with its lower suction requirements for full chord laminar flow, has a higher design Mach number, steeper pressure gradients, a more positive pressure level in the forward region of the lower surface, and a recovery to a less positive pressure at the trailing edge. The two dimensional design Mach numbers for the two airfoils are 0.755 and 0.730 at a common design lift coefficient of 0.60, and their thickness to chord ratios are 0.131 and 0.135, respectively. Off design shock formation characteristics are similar for the two airfoils over a range of Mach numbers between 0.6 and 0.8 and lift coefficients from 0.4 to 0.7. The newer airfoil is similar to the one used in a large chord swept model experiment designed for the Langley 8 Foot Transonic Pressure Tunnel.
    Keywords: AERODYNAMICS
    Type: NASA-TM-84657 , L-15571 , NAS 1.15:84657
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  • 113
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    Nonlinear potential analysis techniques for supersonic-hypersonic configuration design (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Approximate nonlinear inviscid theoretical techniques for predicting aerodynamic characteristics and surface pressures for relatively slender vehicles at moderate hypersonic speeds were developed. Emphasis was placed on approaches that would be responsive to preliminary configuration design level of effort. Second order small disturbance and full potential theory was utilized to meet this objective. Numerical pilot codes were developed for relatively general three dimensional geometries to evaluate the capability of the approximate equations of motion considered. Results from the computations indicate good agreement with higher order solutions and experimental results for a variety of wing, body and wing-body shapes for values of the hypersonic similarity parameter M delta approaching one. Case computational times of a minute were achieved for practical aircraft arrangements.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166078 , NAS 1.26:166078 , NA-82-1170
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  • 114
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    TRANSEP: A program for high lift separated flow about airfoils (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A method and program called TRANSEP is presented that can be used for the analysis of the flow about a low speed airfoil under high lift, massive separation conditions. Since the present program is a modification of the direct-inverse TRANDES code, it can also be used for the design and analysis of transonic airfoils, including the effects of weak viscous interaction. Interactions on program usage, program modifications to convert TRANDES to TRANSEP, and sample cases and results are given.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3376 , NAS 1.26:3376
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  • 115
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    Study of transonic flow over swept wings (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The design of shockless airfoils that are appropriate for experimental work on a supersonic transport with an oblique wing are examined. A series of computer codes for the design and analysis of airfoils and wings in two dimensional and three dimensional transonic flow are studied. The oblique wing 3-D code was the forerunner of the later swept wing code. Techniques to incorporate the effect of an engine or fuselage in the inverse design code while using a minimum of computer resources are developed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-173799 , NAS 1.26:173799
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  • 116
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    Results of design studies and wind tunnel tests of high-aspect-ratio supercritical wings for an energy efficient transport (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: These basic characteristics of critical wings included wing area, aspect ratio, average thickness, and sweep as well as practical constraints on the planform and thickness near the wing root to allow for the landing gear. Within these constraints, a large matrix of wing designs was studied with spanwise variations in the types of airfoils and distribution of lift as well as some small planform changes. The criteria by which the five candidate wings were chosen for testing were the cruise and buffet characteristics in the transonic regime and the compatibility of the design with low speed (high-lift) requirements. Five wing-wide-body configurations were tested in the NASA Ames 11-foot transonic wind tunnel. Nacelles and pylons, flap support fairings, tail surfaces, and an outboard aileron were also tested on selected configurations.
    Keywords: AERODYNAMICS
    Type: NASA-CR-159332 , NAS 1.26:159332 , ACEE-06-FR-9894
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  • 117
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    Helicopter rotor wake geometry and its influence in forward flight. Volume 1: Generalized wake geometry and wake effect on rotor airloads and performance (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An analytic investigation to generalize wake geometry of a helicopter rotor in steady level forward flight and to demonstrate the influence of wake deformation in the prediction of rotor airloads and performance is described. Volume 1 presents a first level generalized wake model based on theoretically predicted tip vortex geometries for a selected representative blade design. The tip vortex distortions are generalized in equation form as displacements from the classical undistorted tip vortex geometry in terms of vortex age, blade azimuth, rotor advance ratio, thrust coefficient, and number of blades. These equations were programmed to provide distorted wake coordinates at very low cost for use in rotor airflow and airloads prediction analyses. The sensitivity of predicted rotor airloads, performance, and blade bending moments to the modeling of the tip vortex distortion are demonstrated for low to moderately high advance ratios for a representative rotor and the H-34 rotor. Comparisons with H-34 rotor test data demonstrate the effects of the classical, predicted distorted, and the newly developed generalized wake models on airloads and blade bending moments. Use of distorted wake models results in the occurrence of numerous blade-vortex interactions on the forward and lateral sides of the rotor disk. The significance of these interactions is related to the number and degree of proximity to the blades of the tip vortices. The correlation obtained with the distorted wake models (generalized and predicted) is encouraging.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3726 , NAS 1.26:3726 , AD-A135555 , UTRC/R83-912666-58-VOL-1
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  • 118
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    The COREL and W12SC3 computer programs for supersonic wing design and analysis (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Two computer codes useful in the supersonic aerodynamic design of wings, including the supersonic maneuver case are described. The nonlinear full potential equation COREL code performs an analysis of a spanwise section of the wing in the crossflow plane by assuming conical flow over the section. A subsequent approximate correction to the solution can be made in order to account for nonconical effects. In COREL, the flow-field is assumed to be irrotional (Mach numbers normal to shock waves less than about 1.3) and the full potential equation is solved to obtain detailed results for the leading edge expansion, supercritical crossflow, and any crossflow shockwaves. W12SC3 is a linear theory panel method which combines and extends elements of several of Woodward's codes, with emphasis on fighter applications. After a brief review of the aerodynamic theory used by each method, the use of the codes is illustrated with several examples, detailed input instructions and a sample case.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3676 , NAS 1.26:3676
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  • 119
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    Phase Averaged Measurements of the Coherent Structure of a Mach Number 0.6 Jet (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The existence of a large scale structure in a Mach number 0.6, axisymmetric jet of cold air was proven. In order to further characterize the coherent structure, phase averaged measurements of the axial mass velocity, radial velocity, and the product of the two were made. These measurements yield information about the percent of the total fluctuations contained in the coherent structure. These measured values were compared to the total fluctuation levels for each quantity and the result expressed as a percent of the total fluctuation level contained in the organized structure at a given frequency. These measurements were performed for five frequencies (St=0.16, 0.32, 0.474, 0.95, and 1.26). All of the phase averaged measurements required that the jet be artificially excited.
    Keywords: AERODYNAMICS
    Type: NASA-CR-175359 , NAS 1.26:175359
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  • 120
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    Family of airfoil shapes for rotating blades (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An airfoil which has particular application to the blade or blades of rotor aircraft such as helicopters and aircraft propellers is described. The airfoil thickness distribution and camber are shaped to maintain a near zero pitching moment coefficient over a wide range of lift coefficients and provide a zero pitching moment coefficient at section Mach numbers near 0.80 and to increase the drag divergence Mach number resulting in superior aircraft performance.
    Keywords: AERODYNAMICS
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  • 121
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    NASA wake interactions research and applications (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: This report describes the development, validation, and application of a computer program for predicting trajectories and ground deposit patterns for particle released in the wakes of airplanes or helicopters. The computer program accounts for the effects of atmospheric turbulence, crosswind, propeller slipstream, terrain variations, evaporation, and plant canopy density on the particle trajectories. In order to validate the prediction method, some comparisons are shown between experimental data and theoretical predictions. Possible applications of the code for spray pattern improvement and for mission operations analysis are illustrated. In addition, the effect of winglets on pattern uniformity, drift, and airplane aerodynamics are presented.
    Keywords: AERODYNAMICS
    Type: SAE PAPER 830764
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  • 122
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    Studies of light-twin wing-body interference (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The results of an analytical study of aerodynamic interference effects for a light twin aircraft are presented. The data presented concentrates on the influence of a wing on a body (the fuselage). Wind tunnel comparisons of three fillets are included, with corresponding computational analysis. Results indicate that potential flow analysis is useful to guide the design of intersection fairings, but experimental tuning is still required. While the study specifically addresses a light twin aircraft, the methods are applicable to a wide variety of aircraft.
    Keywords: AERODYNAMICS
    Type: SAE PAPER 830709
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  • 123
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    Comparison of supercritical airfoil flow calculations with wind-tunnel results (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Navier-Stokes calculations have been performed for a supercritical airfoil at a transonic design condition and at a subsonic condition. Wind-tunnel pressure-rail measurements were employed as boundary data in the calculations to account for wall-interference effects. A fine mesh was used so that most details of the flows were resolved, particular attention having been given to the trailing-edge region. Detailed comparisons are made with the experimental data. Good agreement was obtained on the airfoil except at the trailing edge where separation occurred. Flow details in the trailing-edge region are examined and differences are shown to be attributable to the turbulence model employed.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1688
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  • 124
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    Convergence characteristics of nonlinear vortex-lattice methods for configuration aerodynamics (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Nonlinear panel methods have no proof for the existence and uniqueness of their solutions. The convergence characteristics of an iterative, nonlinear vortex-lattice method are, therefore, carefully investigated. The effects of several parameters, including 1) the surface-paneling method, 2) the integration method of the trajectories of the wake vortices, 3) vortex-grid refinement , and 4) the initial conditions for the first iteration on the computed aerodynamic coefficients and on the flow field details are presented. The convergence of the iterative-solution procedure is usually rapid. The solution converges with grid refinement to a constant value, but the final value is not unique and varies with the wing surface-paneling and wake-discretization methods within some range in the vicinity of the experimental result.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1882
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  • 125
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    A zonal approach for the steady transonic simulation of inviscid rotational flow (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A finite difference zonal method is developed to compute steady inviscid transonic flow by coupling a semi-flux split form of the Euler equations in a vorticity producing zone with a zone of scalar and vector (i.e., dual) potential equations. The dual potential equations permit vorticity convection, but not production, and are efficiently solved as an iteratively decoupled set of scalar equations. Zonal results presented for a nonlifting biconvex airfoil on a stretched Cartesian grid show substantial savings in CPU time compared to solving the semi-flux split Euler equations alone. The dual potential equations also provide an alternate way of treating potential flows with circulation. This has been demonstrated by computing a subcritical flow over a lifting airfoil using generalized curvilinear coordinates.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1927
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  • 126
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    The aerodynamics of propellers and rotors using an acoustic formulation in the time domain (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The aerodynamics of propellers and rotors is especially complicated because of the highly three-dimensional and compressible nature of the flow field. However, in linearized theory the problem is governed by the wave equation, and a numerically-efficient integral formulation can be derived. This reduces the problem from one in space to one over a surface. Many such formulations exist in the aeroacoustics literature, but these become singular integral equations if one naively tries to use them to predict surface pressures, i.e., for aerodynamics. The present paper illustrates how one must interpret these equations in order to obtain nonambiguous results. After the regularized form of the integral equation is derived, a method for solving it numerically is described. This preliminary computer code uses Legendre-Gaussian quadrature to solve the equation. Numerical results are compared to experimental results for ellipsoids, wings, and rotors, including effects due to lift. Compressibility and the farfield boundary conditions are satisfied automatically using this method.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1821
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  • 127
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    Some recent applications of XTRAN3S (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A time-marching finite difference code, XTRAN3S, that solves the three-dimensional transonic small perturbation equation for flow over isolated wings has recently been developed. During initial applications of the program, problems were encountered in the prediction of unsteady forces. The use of a revised grid and force calculation scheme improved those predictions. Comparisons are made between predicted and experimental pressure data for a rectangular supercritical wing. Comparisons of steady and unsteady data at freestream Mach number = 0.700 show good agreement between calculated and experimental values. A comparison of steady data at freestream Mach number = 0.825 shows poor agreement between calculations and experiment. Program difficulties have been encountered with swept and tapered configurations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1811
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  • 128
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    Design and true Reynolds number 2-D testing of an advanced technology airfoil (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A NASA-industry program has been conducted to determine the accuracy of available 2-D airfoil analysis procedures over a wide range of Reynolds numbers. The program also served to develop and demonstrate effective wind tunnel model designs for use in a cryogenic environment. A Lockheed design, CRYO 12X, supercritical, shockfree airfoil was configured using a continuous curvature analytical definition of the ordinates. Test results show a very close ordinate tolerance was necessary to realize the intended pressure distribution. Correlation of test with Korn-Garabedian 2-D analysis pressure data were generally good. GRUMFOIL analysis with a sidewall correction gave a better correlation.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1792
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  • 129
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    Assessment of NASA and RAE viscous-inviscid interaction methods for predicting transonic flow over nozzle afterbodies (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The Langley Research Center of the National Aeronautics and Space Administration and the Royal Aircraft Establishment have undertaken a cooperative program to conduct an assessment of their patched viscous-inviscid interaction methods for predicting the transonic flow over nozzle afterbodies. The assessment was made by comparing the predictions of the two methods with experimental pressure distributions and boattail pressure drag for several convergent circular-arc nozzle configurations. Comparisons of the predictions of the two methods with the experimental data showed that both methods provided good predictions of the flow characteristics of nozzles with attached boundary layer flow. The RAE method also provided reasonable predictions of the pressure distributions and drag for the nozzles investigated that had separated boundary layers. The NASA method provided good predictions of the pressure distribution on separated flow nozzles that had relatively thin boundary layers. However, the NASA method was in poor agreement with experiment for separated nozzles with thick boundary layers due primarily to deficiencies in the method used to predict the separation location.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1789
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    Wall-function boundary conditions in the solution of the Navier-Stokes equations for complex compressible flows (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: To make computer codes for two-dimensional compressible flows more robust and economical, wall functions for these flows, under adiabatic conditions, have been developed and tested. These wall functions have been applied to three two-equation models of turbulence. The tests consist of comparisons of calculated and experimental results for transonic and supersonic flow over a flat plate and for two-dimensional and axisymmetrical transonic shock-wave/boundary-layer interaction flows with and without separation. The calculations are performed with an implicit algorithm that solves the Reynolds-averaged Navier-Stokes equations. It is shown that results obtained agree very well with the data for the complex compressible flows tested, provided criteria for use of the wall functions are followed. The expected savings in cost of the computations and improved robustness of the code were achieved.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1694
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  • 131
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    Surface pressures induced on a flat plate with in-line and side-by-side dual jet configurations (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A jet in a crossflow is of interest in numerous practical situations including jet-powered VTOL aircraft. Two aspects of the problem have received little prior study. First is the effect of the angle of the jet to the crossflow. Second is the performance of dual-jet configurations both in-line and side-by-side. The test plan for this work was designed to address these two aspects. The experiments were conducted in the 7 x 10 tunnel at NASA Ames at velocities from 14.5 - 35.8 m/sec (47.6 - 117.4 ft/sec.). Detailed pressure distributions are presented for single and dual jets over a range of velocity ratios from 3 to 8, spacings from 2 to 6 diameters and injection angles of 90, 75, and 105 degrees. The effects of the various parameters and the differences between the axisymmetric and flat plate geometries on the nature, size, shape, and strength of the interaction regions on the body surfaces are shown.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1849
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  • 132
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    PAN AIR modeling studies (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: PAN AIR is a computer program that predicts subsonic or supersonic linear potential flow about arbitrary configurations. The code's versatility and generality afford numerous possibilities for modeling flow problems. Although this generality provides great flexibility, it also means that studies are required to establish the dos and don'ts of modeling. The purpose of this paper is to describe and evaluate a variety of methods for modeling flows with PAN AIR. The areas discussed are effects of panel density, internal flow modeling, forebody modeling in subsonic flow, propeller slipstream modeling, effect of wake length, wing-tail-wake interaction, effect of trailing-edge paneling on the Kutta condition, well- and ill-posed boundary-value problems, and induced-drag calculations. These nine topics address problems that are of practical interest to the users of PAN AIR.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1830
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  • 133
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    An investigation of wing leading-edge vortices at supersonic speeds (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Studies at subsonic and transonic speeds of the fundamental vortex behavior on the leeward surface of wings have led to the design of several unique and novel leading-edge devices commonly referred to as 'vortex flaps'. The present investigation has the objective to provide some fundamental vortex-flow results obtained at supersonic speeds. Experimental studies were performed in which pressure data and several types of flow visualization data were obtained on the leeward surface of a series of flat delta-wing models to identify the various flow mechanisms which can occur and to determine the effect of leading-edge sweep, Mach number, and angle of attack on the vortex strength and location. The reported investigation forms part of a study which is to explore the use of wing leading-edge vortex technology as a supersonic wing-design tool. The obtained results indicate that the procedure of distributing the vortex force as a pressure variation about a vortex action line is a promising concept.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1816
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  • 134
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    Application of a full potential method for predicting supersonic flow fields and aerodynamic characteristics (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A nonlinear aerodynamic prediction technique which solves the conservative full potential equation has been applied to the analysis of three waverider configurations. This technique was selected based on its capability to analyze the off-design characteristics of the waveriders. Very good correlations were achieved with surface pressure data for both the Mach 4 elliptic cone waverider and the Mach 6 caret-wing derivative. Off-design Mach number and angle-of-attack pressure correlations were very good for the elliptic cone waverider. The range of correlation with data exceeded that expected based on the theory limitations. A surface pressure integration routine was demonstrated and agreement between predicted aerodynamic forces and experimental force data for the Mach 4 waverider was excellent. Analysis of a nonconical waverider configuration was initiated where a discrete input option is used to achieve the computational gridding. Preliminary analysis of this configuration indicates the correct shock location will be predicted.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1802
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  • 135
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    Quasi-three-dimensional turbomachinery flow calculations on multiple hub-to-shroud stream surfaces (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Results are presented from a quasi-three-dimensional calculation of steady (relative), inviscid, adiabatic, subsonic/shock-free transonic flow on multiple hub-to-shroud stream surfaces through turbomachinery blade rows. The quasi-three-dimensional technique incorporates some three-dimensional effects while retaining much of the simplicity of two-dimensional computational methods. Three typical turbomachinery flowfield calculations are presented including an axial-flow compressor rotor, a turbine stator vane cascade, and a radial-inflow turbine rotor. The calculations were performed using quasi-three-dimensional extensions of existing two-dimensional methods. The current results represent an intermediate step in the complete quasi-three-dimensional solution process. However, the results demonstrate the usefulness of the quasi-three-dimensional technique in complementing and extending the applicability of the two-dimensional methods of turbomachinery flow analysis.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1820
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  • 136
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    Convair 990 transonic flow-field simulation about the forward fuselage (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A three-dimensional, viscous flow code was used to calculate the transonic flow about the forebody of the Convair CV-990 (Galileo II) research aircraft stationed at NASA Ames Research Center. The computations were used to determine the location for a differential pressure system. In addition, attitude sensor placements were verified. These instruments comprise a meteorological measurement system, which will be used for global determination of three-dimensional wind data. The code solves the thin layer form of the Reynolds-averaged Navier-Stokes equations using an implicit numerical procedure. The governing equations are written in a generalized, nonorthogonal coordinate system, and are cast in a strong conservation law form. Laminar boundary layer results are presented for free stream Mach number of 0.8 and angles of attack of zero and 2 deg. Use of this computational tool reduced the development time for the location of the sensors and aided in the optimal placement on the aircraft of these devices.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1785
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  • 137
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    Advanced technology exhaust nozzle development (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The requirement for greater tactical aircraft operational capabilities has led to increasing research emphasis on the refinement of engine/airframe integration methods and exhaust nozzle flexibility. A major prospective advancement in the development of these capabilities takes the form of multifunctional exhaust nozzle systems with thrust reversal and thrust vectoring features, whose operation will be shared by both airframe and powerplant control systems. Attention is presently given to the two-dimensional convergent-divergent and single expansion ramp nozzle designs, with emphasis on the variable geometry mechanical systems by which they assume cruising flight, vectoring, and thrust reversal operations. The nozzles have been wind tunnel model-tested for the cases of the F-18 fighter and a supersonic cruise configuration concept.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1286
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  • 138
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    The effect of a normal shock on the aeroelastic stability of a panel (1983)
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    Publication Date: 2019-06-28
    Description: The effect of a standing shock wave on the static and dynamic aeroelastic stability of a flexible panel is investigated using a linear structural and aerodynamic theoretical model. It is found that the shock is generally stabilizing. The lowest critical dynamic pressures are associated with shock positions downstream from the panel, where the panel is uninfluenced by the shock.
    Keywords: AERODYNAMICS
    Type: ASME PAPER 83-APM-28 , (ISSN 0021-8936)
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  • 139
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    An experimental and numerical investigation of the impingement of an oblique shock wave on a body of revolution (1983)
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    Publication Date: 2019-06-28
    Description: Attention is given to the experimental study and numerical simulation of the impingement of oblique shock wave on a cylinder, in order both to document the complex three-dimensional shock wave and boundary layer interaction occurring in practical problems (such as stores carriage interference in a supersonic tactical aircraft) and to conduct a critical comparison of experimental measurements and numerical computations for such complex flows. A thin layer approximation of the Navier-Stokes equations was solved by means of a mixed explicit-implicit scheme. Experimental measurements reveal a highly complex flow field with two distinct adjacent separation zones, regions of high cross flow, and multiply reflected shocks and expansion fans.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1757
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  • 140
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    Conical similarity of shock/boundary layer interactions generated by swept fins (1983)
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    Publication Date: 2019-06-28
    Description: A parametric experimental study has been made of the class of 3D shock wave/turbulent boundary layer interactions generated by swept-leading-edge fins. The fin sweepback angles ranged from 0 to 65 deg at angles of attack of 5, 9, and 15 deg. Two equilibrium 2D turbulent boundary layers with a free-stream Mach number of 2.95 and a Reynolds number of 6.3 x 10 to the 7th/m were used as incoming flow conditions. All the resulting interactions were found to possess conical symmetry of surface pressures and skin friction lines beyond an initial inception zone. Further, these interactions revealed a simple similarity based on inviscid shock strength irrespective of fin sweepback or angle of attack.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1756
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  • 141
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    Numerical study of the effect of an embedded surface-heat source on the separation bubble of supersonic flow (1983)
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    Publication Date: 2019-06-28
    Description: A numerical study of the conjugated problem of a separated supersonic flow field and a conductive solid wall with an embedded heat source is presented. Implicit finite-difference schemes were used to solve the two-dimensional time-dependent compressible Navier-Stokes equations and the time-dependent heat-conduction equation for the solid in both general coordinate systems. A detailed comparison between the thin-layer and Navier-Stokes models was made for steady and unsteady supersonic flow and showed insignificant differences. Steady-state and transient cases were computed and the results show that a temperature pulse at the solid-fluid interface can be used to detect the flow direction near the wall in the vicinity of separation without significant distortion of the flow field.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1753
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  • 142
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    An experimental investigation of multiple shock wave/turbulent boundary layer interactions in a circular duct (1983)
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    Publication Date: 2019-06-28
    Description: Detailed pitot, static and wall pressure measurements have been obtained for multiple shock wave/turbulent boundary layer interactions in a circular duct at a free-stream Mach number of 1.49 and at a unit Reynolds number of 4.90 x 10 to the 6th per meter. The details of the flow field show the formation of a series of normal shock waves with successively decreasing strength and with decreasing distance between the successive shock waves. The overall pressure recovery is much lower than the single normal shock pressure recovery at the same free-stream Mach number. A one-dimensional flow model based on the boundary layer displacement buildup is postulated to explain the formation of a series of normal shock waves.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1744
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  • 143
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    Measurements of compressible secondary flow in a circular S-duct (1983)
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    Publication Date: 2019-06-28
    Description: This paper presents the results of an experimental study of secondary flow in a circular cross section 30 deg - 30 deg S-duct with entrance Mach number of 0.6. Local flow velocity vectors have been measured along the length of the duct at six stations. These measurements have been made using a five-port cone probe. Static and total pressure profiles in the transverse planes are obtained from the cone probe measurements. Wall static pressure measurements along three azimuth angles of 0 deg, 90 deg, and 180 deg along the duct are also made. Contour plots presenting the three dimensional velocity field as well as the total- and static-pressure fields are obtained. Surface oil flow visualization technique has been used to provide details of the flow on the S-duct boundaries. The experimental observations have been compared with typical computational results.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1739
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  • 144
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    Contouring tunnel walls to achieve free-air flow over a transonic swept wing (1983)
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    Publication Date: 2019-06-28
    Description: The effects of wind-tunnel walls on the flow over a swept wing were greatly reduced by wall contouring. Significant reductions in spanwise pressure gradients were achieved by shaping all of the walls to conform to the streamlines over the model in free air. Surface pressure and oil-flow data were used to evaluate the effects of Mach and Reynolds numbers on the design. Comparisons of these data with inviscid calculations indicate that free-air flow is established at a Mach number of 0.74 and at Reynolds numbers above 4.7 million.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1725
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  • 145
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    Calculation of afterbody flows with a composite velocity formulation (1983)
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    Publication Date: 2019-06-28
    Description: A recently developed technique for numerical solution of the Navier-Stokes equations for subsonic, laminar flows is investigated. It is extended here to allow for the computation of transonic and turbulent flows. The basic approach involves a multiplicative composite of the appropriate velocity representations for the inviscid and viscous flow regions. The resulting equations are structured so that far from the surface of the body the momentum equations lead to the Bernoulli equation for the pressure, while the continuity equation reduces to the familiar potential equation. Close to the body surface, the governing equations and solution techniques are characteristic of those describing interacting boundary layers. The velocity components are computed with a coupled strongly implicity procedure. For transonic flows the artificial compressibility method is used to treat supersonic regions. Calculations are made for both laminar and turbulent flows over axisymmetric afterbody configurations. Present results compare favorably with other numerical solutions and/or experimental data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1736
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  • 146
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    Energetics and optimum motion of oscillating lifting surfaces (1983)
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    Publication Date: 2019-06-28
    Description: Low-frequency, unsteady, lifting-line theory is used to characterize the energetics and optimum motion of an unswept rigid wing oscillating harmonically in an inviscid, incompressible flow. The energetics calculations account for the leading edge suction force, the power absorbed in the wing oscillations, and the energy loss rate produced by vortex shedding. Optimization is achieved by minimizing the average energy loss rate in relation to a given thrust, and a unique solution is found in the three dimensional case for low, reduced frequencies. The two-dimensional solution is nonunique, a condition which is examined in terms of the normal modes of the energy loss rate matrix. An invisible mode with a hydrodynamic efficiency of 100 pct is obtained in the two-dimensional case, causing the nonuniqueness of the solution by yielding no fixed positive thrust through perfect unsteady feathering.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1710
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  • 147
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    Navier-Stokes calculations for the vortex wake of a rotor in hover (1983)
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    Publication Date: 2019-06-28
    Description: An efficient finite-difference scheme for the solution of the incompressible Navier-Stokes equation is used to study the vortex wake of a rotor in hover. The solution procedure uses a vorticity-stream function formulation and incorporates an asymptotic far-field boundary condition enabling the size of the computational domain to be reduced in comparison to other methods. The results from the present method are compared with experimental data obtained by smoke flow visualization and hot-wire measurements for several rotor blade configurations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1676
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  • 148
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    An experimental study of surface curvature effects on a supersonic turbulent boundary layer (1983)
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    Publication Date: 2019-06-28
    Description: The mean flow results of an experimental study of compressible turbulent boundary layers in an adverse pressure gradient with and without surface curvature effects are presented. The test was conducted in an axisymmetric flow facility. The upstream Reynolds number based on boundary layer momentum thickness was 5884 and the boundary layer thickness was 0.90 cm. The curvature effects were examined by studying two flows with essentially identical adverse pressure gradients. One flow was along a concave compression surface test section, while the other was along a straight-walled test section. Mean flow measurements included wall static pressure distributions, wall temperatures, pitot pressure profiles and total temperature profiles. The mean flow results indicated that the surface curvature resulted in a definite increase of turbulent mixing in the boundary layer.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1672
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  • 149
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    Experimental studies of the boundary layer on an airfoil at low Reynolds numbers (1983)
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    Publication Date: 2019-06-28
    Description: An experimental wind tunnel investigation was carried out to study the effect of laminar separation bubbles on a NACA 66(3)-018 airfoil for Reynolds numbers less than 4.0 x 10 to the 5th. Leading edge laminar separation bubbles formed for angles of attack of approximately 7 to 12 deg. To study the leading edge separation bubble more closely, hotwire anemometer measurements were made in the airfoil a Reynolds number of 8.0 x 10 to the 4th. Velocity and turbulence intensity profiles were obtained and boundary layer parameters were calculated. Frequency spectra were also calculated at key points in the airfoil boundary layer for this case. Correlation of the anemometry data with static pressure distributions, and flow visualization data provided insight into laminar separation bubble behavior at low Reynolds numbers.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1671
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  • 150
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    Calculation of viscous supersonic flows over finned bodies (1983)
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    Publication Date: 2019-06-28
    Description: The parabolized Navier-Stokes (PNS) equations are used to calculate the viscous, supersonic flow fields about a six-finned projectile and a generic four-finned missile at angles of attack. Since current computer speeds and storage preclude a fully three-dimensional calculation using the unsteady, Reynolds-averaged, Navier-Stokes equations, the applicability of the PNS equations to the above flow fields is of considerable interest. Two important aspects of the calculation are grid generation and the type of smoothing used to prevent nonphysical solutions. This paper includes a description of the grid-generation process. Results in the form of density contours and velocity vector plots are presented for the two configurations. The applicability of the PNS equations to the complicated flow fields considered is successfully demonstrated.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1667
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  • 151
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    Topology of vortices in conical flow (1983)
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    Publication Date: 2019-06-28
    Description: The possibility of the existence of a viscous vortex in a conical flow is discussed. Existence of a viscous vortex is shown to be consistent with a flow-field model where only the velocity components and the total enthalpy behave conically. Existence of a viscous vortex is not consistent with the fully conical, flow-field model; however, far from the nose tip, the physical flow quantities tend to behave asymptotically as a full conical flow. Such a viscous vortex, when it exists, must spiral inwards to the focal point. It can be used as a reasonable model to start numerical solutions of various kinds.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1664
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  • 152
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    Hypersonic flows over biconics using a variable-effective-gamma, Parabolized-Navier-Stokes code (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Hypersonic flows over straight and bent biconics are calculated for a range of freestream conditions in which the gas behind the shock is treated as either perfect or real. The Parabolized-Navier-Stokes (PNS) equations form the basis of the approximation scheme. Good comparisons with experimental data for pressures, forces and moments, heat transfer, and oil-flow patterns serve to validate the perfect-gas version of the code. Circumferential velocity vector plots further aid in the interpretation of leeside oil-flow patterns. A variable-effective-gamma (VEG) option is implemented for the real-gas calculations. Gamma, now defined as the ratio of enthalpy to internal energy, is determined from a locally valid linear relation in enthalpy and pressure at every mesh point which in turn is calculated from a benchmark equilibrium code. The VEG option is easily incorporated into a host PNS code because it uses the underlying perfect-gas structure of the code. Comparisons to experimental data for heat transfer and shock shape in high enthalpy air have been obtained using the VEG option.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1666
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  • 153
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    Three-dimensional transonic nacelle/inlet flowfield computations using an efficient approximate factorization algorithm (1983)
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    Publication Date: 2019-06-28
    Description: A highly efficient computer analysis has been developed for predicting transonic nacelle/inlet flowfields. This algorithm can compute the three-dimensional transonic flowfield about axisymmetric or asymmetric nacelle/inlet configurations at zero or nonzero incidence. The flowfield is determined by solving the full-potential equation in conservative form on a body-fitted curvilinear computational mesh. The difference equations are solved using the AF2 approximate factorization scheme. The effects of boundary layer viscous entrainment are approximated in the inviscid algorithm by applying a surface transpiration velocity which is determined from the calculated boundary layer growth. Computed results and correlations with existing methods and experiment are presented to illustrate application of the analysis.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1417
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  • 154
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    Three-dimensional compressible viscous analysis of mixer nozzles (1983)
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    Publication Date: 2019-06-28
    Description: An analysis-based design procedure for compound-mixer exhaust nozzles is presented and compared to test data. The design approach is based on two numerical solutions to the 3-D viscous compressible Navier-Stokes equations: an equation splitting technique used for the analysis of the core and bypass flow, and a parabolic marching scheme used in the analysis of the mixing duct. The selection of the analytical methods through test data comparisons and their coupling into an integrated design system are discussed. NASA test data is used to demonstrate the validity of the computations from the exhaust system rating station, upstream of the mixer lobe, to the nozzle throat. An estimate is made of the savings in development time and cost utilizing the new procedure.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1401
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  • 155
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    An integral analysis of transonic normal shock wave/turbulent boundary layer interactions in internal flow (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An approximate integral viscous-inviscid interaction method is presented for calculating the development of a turbulent boundary layer subjected to a normal shock wave induced adverse pressure gradient in an internal axisymmetric flow. The inflow conditions and the downstream pressure are provided for the computation. In the supersonic region of shock pressure rise, the Prandtl-Meyer function is used to couple the viscous and inviscid flows. An analytical model for the coupling process is postulated and appropriate equations are defined. Downstream of the sonic point, one-dimensional inviscid flow is assumed for coupling with the viscous flow. The turbulent boundary layer is calculated using Green's integral lag-entrainment method. Comparisons of the solutions with the experimental data are made for interactions which are unseparated, near separation and separated. For comparison purposes, solutions to the time-dependent, mass-averaged, Navier-Stokes equations incorporating a two-equation, Wilcox-Rubesin turbulence model are also shown. The computed results from the integral method show good agreement with experimental data for unseparated interactions and reasonable agreement with the trend of the viscous effects when the interaction becomes increasingly separated.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1402
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  • 156
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    Progress toward the analysis of supersonic inlet flows (1983)
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    Publication Date: 2019-06-28
    Description: A zonal flow analysis procedure was developed to predict the flow through the supersonic diffuser of an axisymmetric mixed compression inlet at angle-of-attack. In this analysis, the inlet flow is divided into three types of regions, each with different dominant flow phenomena. These are the inviscid supersonic core, boundary layer, and shock/boundary layer interaction flows. An appropriate analysis was selected or developed for the three-dimensional flow in each type of region. Procedures were developed to interface these analyses for the overall inlet flow analysis. This analysis was applied to an inlet operating at M = 2.58 at several angle-of-attack conditions. Comparisons are presented between computed and measured flow properties for the inlet and for the component analysis flows. Extensions of the present procedure to include the terminal shock and subsonic diffuser flows are recommended. Desirable experiments for evaluation of the inlet analysis procedure or the component analyses and to support improved modeling or extension of the inlet analysis are defined and recommended.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1371
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  • 157
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    Progress toward the analysis of complex propulsion installation flow phenomenon (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A trend toward replacement of parametric model testing with parametric analysis for the design of aircraft is driven by the rapidly escalating cost of wind tunnel testing, the increasing availability of large fast computers, and powerful numerical flow algorithms. In connection with the complex flow phenomena characteristic of propulsion installations, it is now necessary to employ both parametric analysis and testing for design procedures. Powerful flow analysis techniques are available to predict local flow phenomena. However, the employment of these techniques is very expensive. It is, therefore, necessary to link these analyses with less powerful and less expensive procedures for an accurate analysis of propulsion installation flowfields. However, the interfacing and coupling processes needed are not available. The present investigation is concerned with progress made regarding the development of suitable linking methods. Attention is given to methods of analysis for predicting the flow around a nacelle coupled to a highly swept wing.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1367
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  • 158
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    A static investigation of yaw vectoring concepts on two-dimensional convergent-divergent nozzles (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The flow-turning capability and nozzle internal performance of yaw-vectoring nozzle geometries were tested in the NASA Langley 16-ft Transonic wind tunnel. The concept was investigated as a means of enhancing fighter jet performance. Five two-dimensional convergent-divergent nozzles were equipped for yaw-vectoring and examined. The configurations included a translating left sidewall, left and right sidewall flaps downstream of the nozzle throat, left sidewall flaps or port located upstream of the nozzle throat, and a powered rudder. Trials were also run with 20 deg of pitch thrust vectoring added. The feasibility of providing yaw-thrust vectoring was demonstrated, with the largest yaw vector angles being obtained with sidewall flaps downstream of the nozzle primary throat. It was concluded that yaw vector designs that scoop or capture internal nozzle flow provide the largest yaw-vector capability, but decrease the thrust the most.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1288
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  • 159
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    STOL wind tunnel test results for a tactical supercruiser (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Attention is given to the NASA Langley Research Center's testing of a 10.5 percent-scale supersonic cruiser (supercruiser) aircraft model in its V/STOL wind tunnel, in order to investigate the low speed aerodynamic characteristics of STOL enhancement devices. The STOL devices employed by the supercruiser configuration are high vector angle ramp nozzles, working in conjunction with a remote augmented lift system (RALS), in addition to a canard trim system. Also investigated were thrust reverser/ground plane interaction effects, for the evaluation of landing characteristics. It is noted that STOL approach thrust management requires the use of a partially reversing RALS nozzle which develops approximately 31 percent of main nozzle thrust, and that strong nose-up interactions during ground roll, with reverser operation, may limit dry power engine thrust for braking assistance to about 50 percent of maximum dry power.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1224
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  • 160
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    Method for calculating effects of a propfan on aircraft aerodynamics at subsonic speeds (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A subsonic-flow panel code has been modified to handle the effects of a propeller wake. The effects of the propeller were modelled by a system of ring vortices of constant strength. Principles based on the blade element theory and the momentum theory were used to evaluate the swirl velocity and the pressure increase, across the propeller. Theoretical calculations are compared to experimental results at a Mach number of 0.50. The discrepancies between the theory and the experimental results are analysed. Suggestions for improvements to enhance the accuracy of the theoretical prediction are indicated.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1216
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  • 161
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    Calculation of a simulated 3-D high speed inlet using the Navier-Stokes equations (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A hybrid numerical algorithm, developed to solve the full three-dimensional Navier-Stokes equations, is applied to the computation of the flowfield in a simulated three-dimensional high speed aircraft inlet at a Mach number of 2.5 and Reynolds number of 1.4 x 10 to the 7th based on inlet length. The numerical algorithm incorporates a coordinate transformation in order to handle general flow geometries, and utilizes the algebraic turbulent eddy viscosity model of Baldwin and Lomax. The hybrid algorithm has been vectorized on the CDC CYBER 203 computer using the SL/1 vector programming language developed at NASA Langley. The computed results are compared with experimental measurements of the ramp and cowl static pressures, and boundary layer pitot profiles. The results are also compared with a previous two-dimensional Navier-Stokes computation of the same configuration. The agreement with the experimental data is generally good; however, additional improvements in turbulence modeling are needed.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1165
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  • 162
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    Empennage/afterbody integration for single and twin-engine fighter aircraft (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An extensive experimental program to determine the effects of empennage surfaces on single and twin-engine afterbody/nozzle drag has been conducted by the Propulsion Aerodynamics Branch at the NASA Langley Research Center. Empennage interference drag was obtained by using experimental values of afterbody/nozzle drag and computed values of empennage drag. The effects of tail location, span, number (single versus twin), toe angle, cant angle, camber and root chord length are discussed. The magnitude of empennage interference drag on single and twin engine configurations is examined.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1126
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  • 163
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    Three-dimensional flow studies on a slotted transonic wind tunnel wall (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Previously cited in issue 06, p. 799, Accession no. A82-17855
    Keywords: AERODYNAMICS
    Type: (ISSN 0001-1452)
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  • 164
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    Improved relaxation schemes for transonic potential calculations (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A block relaxation scheme, grouped in a red-black ordering, is applied to transonic airfoil calculations using body fitted coordinates. The scheme is simple and is easily vectorizable. Detailed comparisons with Approximate Factorization Method (AF2) are presented and it is shown that the improved relaxation scheme is competitive in all cases considered. Transonic results, of engineering accuracy, on an 0-type grid of 149 x 30 points, are ususally obtained within two hundred iterations (approximately 40 seconds on Cyber 175).
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-0372
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  • 165
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    Entrainment and mixing in thrust augmenting ejectors (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An experimental investigation of two-dimensional thrust augmenting ejector flows has been conducted. Measurements of the shroud surface pressure distribution, mean velocity, turbulent intensities and Reynolds stresses were made in two shroud geometries at various primary nozzle pressure ratios. The effects of shroud geometry and primary nozzle pressure ratio on the shroud surface pressure distribution, mean flow field and turbulent field were determined. From these measurements the evolution of mixing within the shroud of the primary flow and entrained fluid was obtained. The relationship between the mean flow field, the turbulent field and the shroud surface pressure distribution is discussed.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-0172
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  • 166
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    Transonic flow calculations using the Euler equations (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Previously cited in issue 06, p. 796, Accession no. A82-17786
    Keywords: AERODYNAMICS
    Type: (ISSN 0001-1452)
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  • 167
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    Aerodynamic theory for wing with side edge passing subsonically through a gust (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An approximate solution for the unsteady loading near the square-shape tip of a wing passing through an oblique gust is obtained in closed form. The aerodynamic theory developed can be used to predict airloads felt by a helicopter blade experiencing a blade/vortex interaction for high blade tip speed and/or for small vertical blade/vortex separation. Under these conditions one can show that the blade's trailing edge has little influence on the character of the chordwise loading at all spanwise sections; thus, the chord may be allowed to extend to infinity in the downstream direction. Therefore, the model considered here is that of a quarter-infinite flat plate wing with side edge passing subsonically through an oblique gust.
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 21; June 198
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  • 168
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    Analysis of aerothermal loads on spherical dome protuberances (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Hypersonic flow over spherical dome protuberances was investigated to determine increased pressure and heating loads to the surface. The configuration was mathematically modeled in a time-dependant three-dimensional analysis of the conservation of mass, momentum (Navier-Stokes), and energy equations. A boundary mapping techique was used to obtain a rectangular parallelepiped computational domain, and a MacCormack explicit time-split predictor-corrector finite difference algorithm was used to obtain solutions. Results show local pressures and heating rates for domes one-half, one, and two boundary layer thicknesses high were increased by factors on the order of 1.4, 2, and 6, respectively. However, because lee-side pressure and thermal loads were reduced the two lower height domes did not experience any net increase in total loads. The total loads on the higher dome were increased by twenty-five percent. Flow over the lower dome was everywhere attached while flow over the intermediate dome had small windward and leeside separations. The higher dome had an unsteady windward separation region and a large leeside separation region. Trailing vortices form on all domes with intensity increasing with dome height. Discussions of applying the results to a thermally bowed thermal protection system are presented.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1557
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  • 169
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    Aerodynamic characteristics of generalized bent biconic bodies for aero-assisted, orbital-transfer vehicles (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A method was developed to generate the surface coordinates of body shapes suitable for aero-assisted, orbital-transfer vehicles (AOTVs) by extending bent biconic geometries. Lift, drag, and longitudinal moments were calculated for the bodies using Newtonian flow theory. These techniques were applied to symmetric and asymmetric aerobraking vehicles, and to an aeromaneuvering vehicle with high L/D. Results for aerobraking applications indicate that a 70-deg, fore half cone angle with spherically blunted nose, rounded edges, and a slight asymmetry would be appropriate. Moreover, results show that an aeromaneuvering vehicle with L/D greater than 2.0, and with sufficient stability, is feasible.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1512
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  • 170
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    Viscous real gas flowfields about three dimensional configurations (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Laminar, real gas hypersonic flowfields over a three dimensional configuration are computed using an unsteady, factored implicit scheme. Local chemical and thermodynamic properties are evaluated by an equilibrium composition method. Transport properties are obtained from individual species properties and application of a mixture rule. Numerical solutions are presented for an ideal gas and equilibrium air for free-stream Mach numbers of 13 and 15 and at various angles of attack. The effect of real gas is to decrease the shock-layer thickness resulting from decreased shock-layer temperatures and corresponding increased density. The combined effects of viscosity and real gas are to increase the subsonic layer near the wall.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1511
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  • 171
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    Low L/D aerobrake test at Mach 10 (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An aerobraking Orbital transfer vehicle may be used to increase the Space Shuttle mission capability to and from high orbits. A Mach 10 wind-tunnel test was performed for a low lift-drag aerobrake, to define a preliminary aerothermal environment for this candidate concept. Test hardware simulated the ribs and stretched fabric of conceptual flight hardware. Pressures, paint-melting histories, and oilflow data were measured on the brake. Pressure and thermocouple heating rate data were measured on the payload. Brake peak heating is at the edge at all angles of attack, although the stagnation point is not outboard of 75 percent radius even at 20 degrees angle of attack. Brake ribs show slightly higher heating than flats, although pressures are essentially constant. Payload peak heating occurs near 12 degrees angle of attack, and is 30 percent of the sphere stagnation point heating (for a sphere of brake diameter). Payload pressure distributions follow the heating pattern. Reynolds number effects are small on the brake and large on the payload, for the range of test conditions: 0.4-1.0 million/foot.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1509
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  • 172
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    Aerothermal environment in control surface gaps in hypersonic flow - An overview (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Experimental and analytical research by the NASA Langley Research Center to develop an understanding of the fluid and thermal environment in control surface gaps such as the spanwise gap of the wing elevon and chordwise gap of a split elevon configuration typical of the Space Shuttle are summarized. Although the experimental and analytical studies were initiated too late to significantly impact the basic Shuttle design they do provide a fundamental understanding of the basic fluid/thermal environment in control surface gaps and help to establish a firm data base for future vehicle design.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1483
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  • 173
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    Computation of nonlinear supersonic potential flow over three-dimensional surfaces (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Previously cited in issue 12, p. 1850, Accession no. A82-27090
    Keywords: AERODYNAMICS
    Type: (ISSN 0021-8669)
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  • 174
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    Aerodynamic design using numerical optimization (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The procedure of using numerical optimization methods coupled with computational fluid dynamic (CFD) codes for the development of an aerodynamic design is examined. Several approaches that replace wind tunnel tests, develop pressure distributions and derive designs, or fulfill preset design criteria are presented. The method of Aerodynamic Design by Numerical Optimization (ADNO) is described and illustrated with examples.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85550 , NAS 1.15:85550 , CFDL-TR-83-2
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  • 175
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    Reynolds number effects on pressure loss and turbulence characteristics of four tube-bundle heat exchangers (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The aerodynamic characteristics of pressure loss and turbulence on four tube-bundle configurations representing heat-exchanger geometries with nominally the same heat capacity were measured as a function of Reynolds numbers from about 4000 to 400,000 based on tube hydraulic diameter. Two configurations had elliptical tubes, the other two had round tubes, and all four had plate fins. The elliptical-tube configurations had lower pressure loss and turbulence characteristics than the round-tube configurations over the entire Reynolds number range.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85807 , L-15721 , NAS 1.15:85807
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  • 176
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    Unsteady laminar boundary-layer calculations on oscillating configurations including backflow. Part 2: Airfoil in high-amplitude pitching motion. Dynamic stall (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A previously developed finite-difference procedure for calculating unsteady, incompressible, laminar boundary layers on an oscillating flat plate is applied to a wing section undergoing high-amplitude pitching oscillations about various mean incidences. To start the entire boundary-layer calculation, appropriate initial conditions and outer boundary conditions are specified, using a stagnation-point fixed frame of reference. The breakdown of the numerical calculation procedure in the x,t-domain is interpreted to coincide with unsteady separation. Details of the boundary-layer behavior in the vicinity of separation are investigated, and a close analogy between the present results and those for a three-dimensional steady separation is found.
    Keywords: AERODYNAMICS
    Type: NASA-TM-84319-PT-2 , A-9403 , NAS 1.15:84319-PT-2
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  • 177
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    Demonstration of the coast-down technique for determining train resistances (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Full-scale measurement or validation of the various factors of train running resistance is an essential step in decreasing train energy consumption. Such a measurement capability would enable railroads to evaluate the cost benefits of operational and train consistent configuration changes, and new vehicle and truck designs for decreasing aerodynamic drag and rolling resistance. A decrease in the rolling resistance affects more than just a decrease in energy consumption; it also will result in decreased mechanical wear, hence less wheel and rail maintenance and replacement costs. A demonstration of a simple coast-down technique (based on computer-reduction of distance history) was accomplished using specially configured trains on main line rail provided by the Atchison, Topeka and Sante Fe Railway Co. This demonstration test shows that this distance-history coast-down technique for trains is easy to execute in the field. The total running resistance history was accurately determined and subsequently separated into rolling resistance (mechanical friction) and aerodynamic drag.
    Keywords: AERODYNAMICS
    Type: NASA-CR-173468 , JPL-PUB-83-85 , NAS 1.26:173468
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  • 178
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    Experimental study of the separating confluent boundary-layer. Volume 2: Experimental data (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental low speed study of the separating confluent boundary layer on a NASA GAW-1 high lift airfoil is described. The airfoil was tested in a variety of high lift configurations comprised of leading edge slat and trailing edge flap combinations. The primary test instrumentation was a two dimensional laser velocimeter (LV) system operating in a backscatter mode. Surface pressures and corresponding LV derived boundary layer profiles are given in terms of velocity components, turbulence intensities and Reynolds shear stresses as characterizing confluent boundary layer behavior up to and beyond stall. LV derived profiles and associated boundary layer parameters and those obtained from more conventional instrumentation such as pitot static transverse, Preston tube measurements and hot-wire surveys are compared.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166018 , NAS 1.26:166018 , LG82ER0184
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  • 179
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    Numerical solution of three-dimensional unsteady transonic flow over wings including inviscid/viscous interactions (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A numerical procedure is presented for computing the unsteady transonic flow field about three dimensional swept wings undergoing general time dependent motion. The outer inviscid portion of the flow is assumed to be governed by the modified unsteady transonic small disturbance potential equation which is integrated in the time domain by means of an efficient alternating direction implicit approximate factorization algorithm. Gross dominant effects of the shock boundary layer interaction are accounted for by a simple empirically defined model. Viscous flow regions adjacent to the wing surface and in the trailing wake are described by a set of integral equations appropriate for compressible turbulent shear layers. The two dimensional boundary layer equations are applied quasi-statically stripwise across the span. Coupling with the outer inviscid flow is implemented through use of the displacement thickness concept within the limitations of small disturbance theory. Validity of the assumptions underlying the method is established by comparison with experimental data for the flow about a high aspect ratio transport wing having an advanced airfoil section.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166561 , NAS 1.26:166561
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  • 180
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    A wing concept for supersonic maneuvering (1983)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: A theoretical and experimental program in which a wing concept for supersonic maneuvering was developed and then demonstrated experimentally in a series of wind tunnel tests is described. For the typical fighter wing, the problem of obtaining efficient lift at supersonic maneuvering C sub 's occurs due to development of a strong crossflow shock, and boundary layer separation. A natural means of achieving efficient supersonic maneuvering is based on controlling the non-linear inviscid crossflow on the wing in a manner analogous to the supercritical aerodynamic methods developed for transonic speeds. The application of supercritical aerodynamics to supersonic speeds is carried out using Supercritical Conical Camber (SC3). This report provides an aerodynamic analysis of the effort, with emphasis on wing design using non-linear aerodynamics. The substantial experimental data base is described in three separate wind tunnel reports, while two of the computer programs used in the work are also described in a separate report. Based on the development program it appears that a controlled supercritical crossflow can be obtained reliably on fighter-type wing planforms, with an associated drag due to lift reduction of about 20% projected using this concept.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3763 , NAS 1.26:3763
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  • 181
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    Three-dimensional implicit lambda methods (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: This paper derives the three dimensional lambda-formulation equations for a general orthogonal curvilinear coordinate system and provides various block-explicit and block-implicit methods for solving them, numerically. Three model problems, characterized by subsonic, supersonic and transonic flow conditions, are used to assess the reliability and compare the efficiency of the proposed methods.
    Keywords: AERODYNAMICS
    Type: NASA-CR-172264 , NAS 1.26:172264 , REPT-83-62
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  • 182
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    Multicomponent velocity measurement in a pison-cylinder configuration using laser velocimetry (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Because of its nonintrusive nature, Laser Doppler Velocimetry (LDV) has become a popular tool for velocity measurements in internal combustion engines. This work shows how one can use an on-axis measurement technique, in conjunction with the standard two channel LDV technique, to make simultaneous three-component measurements using a single focusing lens. Simultaneous measurement of two of these three components in a piston-cylinder configuration is demonstrated.
    Keywords: AERODYNAMICS
    Type: NASA-TM-83534 , E-1835 , NAS 1.15:83534
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  • 183
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    Aerodynamic characteristics, including effect of body shape, of a Mach 6 aircraft concept (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Longitudinal aerodynamic characteristics for a hydrogen-fueled hypersonic transport concept at Mach 6 are presented. The model components consist of four bodies with identical longitudinal area distributions but different cross-sectional shapes and widths, a wing, horizontal and vertical tails, and a set of wing-mounted nacelles simulated by slid bodies on the wing upper surface. Lift-drag ratios were found to be only sightly affected by fuselage planform width or cross sectional shape. Relative distribution of fuselage volume above and below the wing was found to have an effect on the lift-drag ratio, with a higher lift drag ratio produced by the higher wing position.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2235 , L-15675 , NAS 1.60:2235
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  • 184
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    Effect of thrust reverser operation on the lateral-directional characteristics of a three-surface F-15 model at transonic speeds (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the Langley 16 Foot Transonic Tunnel to determine the lateral directional aerodynamic characteristics of a fully metric 0.04 scale model of the F-15 three surface configuration (canards, horizontal tails) with twin two dimensional nozzles and twin axisymmetric nozzles installed. The effects of two dimensional nozzle in flight thrust reversing and rudder deflection were also determined. Test data were obtained at static conditions and at Mach numbers from 0.60 to 1.20 over an angle of attack range from -2 deg to 15 deg. Reynolds number varied from 2.6 million to 3.8 million. Angle of sideslip was set at approximately 0 deg and -5 deg for all configurations and at -10 deg for selected configurations.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2234 , L-15648 , NAS 1.60:2234
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  • 185
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    Thrust-induced effects on subsonic longitudinal aerodynamic characteristics of a vectored-engine-over-wing configuration (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the Langley 4 by 7 Meter Tunnel of the thrust induced effects on the longitudinal aerodynamic characteristics of a vectored-engine-over-wing fighter aircraft. The investigation was conducted at Mach numbers from 0.14 to 0.17 over an angle-of-attack range from -2 deg to 26 deg. The major model variables were the spanwise blowing nozzle sweep angle and main nozzle vector angle along with trailing edge, flap deflections. The overall thrust coefficient (main and spanwise nozzles) was varied from 0 (jet off) to 2.0. The results indicate that the thrust-induced effects from the main nozzle alone were small and mainly due to boundary-layer control affecting a small area behind the nozzle. When the spanwise blowing nozzles were included, the induced effects were larger than the main nozzle alone and were due to both boundary layer control and induced circulation lift. No leading edge vortex effects were evident.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2228 , L-15629 , NAS 1.60:2228
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  • 186
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    Aerodynamic force measurements with a strain-gage balance in a cryogenic wind tunnel (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Aerodynamic force measurements on a generalized 75 deg delta wing model with sharp leading edges were made with a three component internal strain gage balance in a cryogenic wind tunnel at stagnation temperatures of 300 K, 200 K, and 110 K. The feasibility of using a strain gage balance without thermal control in a cryogenic environment as well as the use of electrical resistance heaters, an insulator between the model and the balance, and a convection shield on the balance was investigated. Force and moment data on the delta wing model as measured by the balance are compared at the different temperatures while holding constant either the Reynolds number or the tunnel stagnation pressure. Tests were made at Mach numbers of 0.3 and 0.5 and at angles of attack up to 29 deg. The results indicate that it is feasible to acquire accurate force and moment data while operating at steady state thermal conditions in a cryogenic wind tunnel, either with or without electrical heaters on the balance. Within the limits of the balance accuracy, there were no apparent Reynolds number effects on the aerodynamic results for the delta wind model.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2251 , L-15685 , NAS 1.60:2251
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  • 187
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    Internal pressure distributions for a two-dimensional thrust-reversing nozzle operating at a free-stream Mach number of zero (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the static test facility of the Langley 16-Foot Transonic Tunnel to measure static pressure distributions inside a nonaxisymmetric thrust reversing nozzle. The tests were made at nozzle total pressures ranging from ambient to about eight times ambient pressure at a free stream Mach number of zero. Tabulated pressure data are presented.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85655 , L-15582 , NAS 1.15:85655
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  • 188
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    Vortex flap flow reattachment line and subsonic longitudinal aerodynamic data on 50 deg to 74 deg Delta wings on common fuselage (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Positions of the primary vortex flow reattachment line and longitudinal aerodynamic data were obtained at Mach number 0.3 for a systematic series of vortex flaps on delta wing body configurations with leading edge sweeps of 50, 58, 66, and 74 deg. The investigation was performed to study the parametric effects of wing sweep, vortex flap geometry and deflection, canards, and trailing edge flaps on the location of the primary vortex reattachment line relative to the flap hinge line. The vortex reattachment line was located via surface oil flow photographs taken at selected angles of attack. Force and moment measurements were taken over an angle of attack range of -1 deg to 22 deg at zero sideslip angle for many configurations to further establish the data base and to assess the aforementioned parametric effects on longitudinal aerodynamics. Both the flow reattachment and aerodynamic data are presented.
    Keywords: AERODYNAMICS
    Type: NASA-TM-84618 , L-15702 , NAS 1.15:84618
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  • 189
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    A Computer Program for the Calculation of Three-Dimensional Transonic Nacelle/Inlet Flowfields (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A highly efficient computer analysis was developed for predicting transonic nacelle/inlet flowfields. This algorithm can compute the three dimensional transonic flowfield about axisymmetric (or asymmetric) nacelle/inlet configurations at zero or nonzero incidence. The flowfield is determined by solving the full-potential equation in conservative form on a body-fitted curvilinear computational mesh. The difference equations are solved using the AF2 approximate factorization scheme. This report presents a discussion of the computational methods used to both generate the body-fitted curvilinear mesh and to obtain the inviscid flow solution. Computed results and correlations with existing methods and experiment are presented. Also presented are discussions on the organization of the grid generation (NGRIDA) computer program and the flow solution (NACELLE) computer program, descriptions of the respective subroutines, definitions of the required input parameters for both algorithms, a brief discussion on interpretation of the output, and sample cases to illustrate application of the analysis.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166528 , NAS 1.26:166528 , LG83ER0163
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  • 190
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    A grid-embedding transonic flow analysis computer program for wing/nacelle configurations (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An efficient grid-interfacing zonal algorithm was developed for computing the three-dimensional transonic flow field about wing/nacelle configurations. the algorithm uses the full-potential formulation and the AF2 approximate factorization scheme. The flow field solution is computed using a component-adaptive grid approach in which separate grids are employed for the individual components in the multi-component configuration, where each component grid is optimized for a particular geometry such as the wing or nacelle. The wing and nacelle component grids are allowed to overlap, and flow field information is transmitted from one grid to another through the overlap region using trivariate interpolation. This report represents a discussion of the computational methods used to generate both the wing and nacelle component grids, the technique used to interface the component grids, and the method used to obtain the inviscid flow solution. Computed results and correlations with experiment are presented. also presented are discussions on the organization of the wing grid generation (GRGEN3) and nacelle grid generation (NGRIDA) computer programs, the grid interface (LK) computer program, and the wing/nacelle flow solution (TWN) computer program. Descriptions of the respective subroutines, definitions of the required input parameters, a discussion on interpretation of the output, and the sample cases illustrating application of the analysis are provided for each of the four computer programs.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166529 , NAS 1.26:166529 , LG83ER0164
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  • 191
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    Analytical model of rotor wake aerodynamics in ground effect (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The model and the computer program developed provides the velocity, location, and circulation of the tip vortices of a two-blade helicopter in and out of the ground effect. Comparison of the theoretical results with some experimental measurements for the location of the wake indicate that there is excellent accuracy in the vicinity of the rotor and fair amount of accuracy far from it. Having the location of the wake at all times enables us to compute the history of the velocity and the location of any point in the flow. The main goal of out study, induced velocity at the rotor, can also be calculated in addition to stream lines and streak lines. Since the wake location close to the rotor is known more accurately than at other places, the calculated induced velocity over the disc should be a good estimate of the real induced velocity, with the exception of the blade location, because each blade was replaced only by a vortex line. Because no experimental measurements of the wake close to the ground were available to us, quantitative evaluation of the theoretical wake was not possible. But qualitatively we have been able to show excellent agreement. Comparison of flow visualization with out results has indicated the location of the ground vortex is estimated excellently. Also the flow field in hover is well represented.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166533 , NAS 1.26:166533
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  • 192
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    Investigation of installation effects on twin-engine convergent-divergent nozzles (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine installation effects on convergent-divergent nozzles applicable to twin-engine reduced-power supersonic cruise aircraft. Tests were conducted at Mach numbers from 0.50 to 1.20, angles of attack from -5 deg to 9 deg, and at nozzle pressure ratios from jet off (1.0) to 8.0. The effects of empennage arrangement, nozzle length, and afterbody closure on total and component drag coefficients were investigated.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2205 , L-15609 , NAS 1.60:2205
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  • 193
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    Specification for a Program for an Interative Aeroelastic Solution (PIAS) (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An engineering and software specification which was written for a computer program to calculate aeroelastic structural loads including the effects of nonlinear aerodynamics is presented. The procedure used in the program for an iterative aeroelastic solution (PIAS) is to alternately execute two computer codes: one to calculate aerodynamic loads for a specific wing shape, and another to calculate the deflected shape caused by this loading. A significant advantage to the design of PIAS is that the initial aerodynamic module can be replaced with others. The leading edge vortex (LEV) program is used as the aerodynamic module in PIAS. This provides the capability to calculate aeroelastic loads, including the effects of a separation induced leading edge vortex. The finite element method available in ATLAS Integrated structural analysis and design system is used to determine the deflected wing shape for the applied aerodynamics and inertia loads. The data management capabilities in ATLAS are used by the execution control monitor (ECM) of PIAS to control the solution process.
    Keywords: AERODYNAMICS
    Type: NASA-CR-172200 , NAS 1.26:172200 , D6-52134
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  • 194
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    Tabulation of data from tests of an NPL 9510 airfoil in the Langley 0.3-meter transonic cryogenic tunnel (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The tabulated data from tests of a six inch chord NPL 9510 airfoil in the Langley 0.3-Meter Transonic Cryogenic Tunnel. The tests were performed over the following range of conditions: Mach numbers of 0.35 to 0.82, total temperature of 94 K to 300 K, total pressure of 1.20 to 5.81 atm, Reynolds number based on chord of 1.34 x 10 to the 6th to 48.23 x 10 to the 6th, and angle of attack of 0 deg to 6 deg. The NPL 9510 airfoil was observed to have decreasing drag coefficient up to the highest test Reynolds number.
    Keywords: AERODYNAMICS
    Type: NASA-TM-84579 , NAS 1.15:84579
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  • 195
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    XTRAN2L: A program for solving the general-frequency unsteady transonic small disturbance equation (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A program, XTRAN2L, for solving the general-frequency unsteady transonic small disturbance potential equation was developed. It is a modification of the LTRAN2-NLR code. The alternating-direction-implicit (ADI) method of Rizzetta and Chin is used to advance solutions of the potential equation in time Engquist-Osher monotone spatial differencing is used in the ADI solution algorithm. As a result, the XTRAN2L code is more robust and more efficient than similar codes that use Murman-Cole type-dependent spatial differencing. Nonreflecting boundary conditions that are consistent with the general-frequency equation have been developed and implemented at the far-field boundaries. Use of those conditions allow the computational boundaries to be moved closer to the airfoil with no loss of accuracy. This makes the XTRAN2L code more economical to use.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85723 , NAS 1.15:85723
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  • 196
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    Predicting aerodynamic characteristics of vortical flows on three-dimensional configurations using a surface-singularity panel method (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A general low-order surface-singularity panel method is used to predict the aerodynamic characteristics of a problem where a wing-tip vortex from one wing closely interacts with an aft mounted wing in a low Reynolds Number flow; i.e., 125,000. Nonlinear effects due to wake roll-up and the influence of the wings on the vortex path are included in the calculation by using a coupled iterative wake relaxation scheme. The interaction also affects the wing pressures and boundary layer characteristics: these effects are also considered using coupled integral boundary layer codes and preliminary calculations using free vortex sheet separation modelling are included. Calculated results are compared with water tunnel experimental data with generally remarkably good agreement.
    Keywords: AERODYNAMICS
    Type: AGARD Aerodyn. of Vortical Type Flows in Three Dimensions; 12 p
    Format: text
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  • 197
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    Helicopter rotor wake geometry and its influence in forward flight. Volume 2: Wake geometry charts (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Isometric and projection view plots, inflow ratio nomographs, undistorted axial displacement nomographs, undistorted longitudinal and lateral coordinates, generalized axial distortion nomographs, blade/vortex passage charts, blade/vortex intersection angle nomographs, and fore and aft wake boundary charts are discussed. Example condition, in flow ratio, undistorted axial location, longitudinal and lateral coordinates, axial coordinates distortions, blade/tip vortex intersections, angle of intersection, and fore and aft wake boundaries are also discussed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3727 , NAS 1.26:3727 , R83-912666-58
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  • 198
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    Experimental investigation of forebody and wing leading-edge vortex interactions at high angles of attack (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental investigation was conducted to assess the vortex flow-field interactions on an advanced, twin-jet fighter aircraft configuration at high angles of attack. Flow-field surveys were conducted on a small-scale model in the Northrop 0.41 - by 0.60-meter water tunnel and, where appropriate, the qualitative observations were correlated with low-speed wind tunnel data trends obtained on a large-scale model of the advanced fighter in the NASA Langley Research Center 30- by 60-foot (9.1- by 18.3-meter) facility. Emphasis was placed on understanding the interactions of the forebody and LEX-wing vortical flows, defining the effects on rolling moment variation with sideslip, and identifying modifications to control or regulate the vortex interactions at high angles of attack. The water tunnel flow visualization results and wind tunnel data trend analysis revealed the potential for strong interactions between the forebody and LEX vortices at high angles of attack. In particular, the forebody flow development near the nose could be controlled by means of carefully-positioned radome strakes. The resultant strake-induced flow-field changes were amplified downstream by the more powerful LEX vortical motions with subsequent large effects on wing flow separation characteristics.
    Keywords: AERODYNAMICS
    Type: AGARD Aerodyn. of Vortical Type Flows in Three Dimensions; 20 p
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  • 199
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    A computer program for estimating the aerodynamic characteristics of NACA 16-series airfoils (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A computer program written in a table ""look-up'' format, is presented which provides a comprehensive data base on NACA 16-series airfoils. The geometry covered is limited to cambers for a design-lift coefficient from 0.0 to 0.7 and thickness ratios from 4 to 21%. The data include Mach numbers from 0.3 to 1.6, angles of attack from -4 to 8 degrees, and lift coefficients from 0.0 to 0.8. Extrapolation is used to obtain data from Mach numbers, angles of attack, and lift coefficients beyond those for which data are available. A routine to adjust the lift and drag coefficients beyond stall is included. The uses and limitations of the program are also discussed.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85696 , NAS 1.15:85696
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  • 200
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    Effects of Nacelle configuration/position on performance of subsonic transport (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental study was conducted to explore possible reductions in installed propulsion system drag due to underwing aft nacelle locations. Both circular (C) and D inlet cross section nacelles were tested. The primary objectives were: to determine the relative installed drag of the C and D nacelle installations; and, to compare the drag of each aft nacelle installation with that of a conventional underwing forward, drag of each aft nacelle installation with that of a conventional underwing forward, pylon mounted (UTW) nacelle installation. The tests were performed in the NASA-Langley Research Center 16-Foot Transonic Wind Tunnel at Mach numbers from 0.70 to 0.85, airplane angles of attack from -2.5 to 4.1 degrees, and Reynolds numbers per foot from 3.4 to 4.0 million. The nacelles were installed on the NASA USB full span transonic transport model with horizontal tail on. The D nacelle installation had the smallest drag of those tested. The UTW nacelle installation had the largest drag, at 6.8 percent larger than the D at Mach number 0.80 and lift coefficient (C sub L) 0.45. Each tested configuration still had some interference drag, however. The effect of the aft nacelles on airplane lift was to increase C sub L at a fixed angle of attack relative to the wing body. There was higher lift on the inboard wing sections because of higher pressures on the wing lower surface. The effects of the UTW installation on lift were opposite to those of the aft nacelles.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3743 , NAS 1.26:3743 , LR-30436
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