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  • AERODYNAMICS  (738)
  • 1980-1984  (485)
  • 1970-1974  (252)
  • 1925-1929  (1)
  • 1984  (485)
  • 1972  (252)
  • 1926  (1)
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  • 1980-1984  (485)
  • 1970-1974  (252)
  • 1925-1929  (1)
Year
  • 101
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    Supersonic wing design concepts employing nonlinear flows (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Three nonlinear flow concepts for the design of supersonic wings are reviewed. The specific concepts are: leading-edge thrust, supercritical crossflow, and leading-edge vortex flow. The major results of the experimental-theoretical studies supporting the development of these concepts are presented and discussed. Also, supporting aerodynamic prediction methods are described and example applications are given. Recommendations for further development of each concept are made.
    Keywords: AERODYNAMICS
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  • 102
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    Validation of a transonic analysis code for use in preliminary design of advanced transport configurations (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: The WBPPW code has the capability of analyzing flow-field effects about configurations which include wing pylons and engine nacelles or pods in addition to the basic wing/fuselage combination. Using the concept of grid embedding, the code solves the extended small disturbance transonic flow equation for complex flow interactions of the various configuration components. A general description of the code and solution algorithm is included. Results are presented and compared with experiment for various configurations which encompass the code capabilities. These include wing planform and wing contour modifications and variations in nacelle position beneath a high-aspect-ratio wing. Results are analyzed in the light of preliminary design, where the capability to accurately compute flow-field effects resulting from various configuration perturbations is important. The comparisons show that the computational results are sensitive to subtle design modifications and that the code could be used as an effective guide during the design process for transport configurations.
    Keywords: AERODYNAMICS
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  • 103
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    Prediction of transonic separated flows (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Johnson et al. (1982) have provided a detailed comparison between a thoroughly documented transonic flow with shock-induced separations and solutions of the flow using the Navier-Stokes equations. According to this comparison, there were several deficiencies in the computations. The present investigation takes into account new experimental data which have been obtained in a larger wind tunnel with the same test model for a wider range of freestream Mach numbers. The results of new Navier-Stokes computations using more compatible boundary conditions are shown, and the effects of the turbulence model choice on predicting Mach number trends are assessed.
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 1001-100
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  • 104
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    Improved design of subcritical and supercritical cascades using complex characteristics and boundary-layer correction (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: The method of complex characteristics and hodograph transformation for the design of shockless airfoils was extended to design supercritical cascades with high solidities and large inlet angles. This capability was achieved by introducing a conformal mapping of the hodograph domain onto an ellipse and expanding the solution in terms of Tchebycheff polynomials. A computer code was developed based on this idea. A number of airfoils designed with the code are presented. Various supercritical and subcritical compressor, turbine and propeller sections are shown. The lag-entrainment method for the calculation of a turbulent boundary layer was incorporated to the inviscid design code. The results of this calculation are shown for the airfoils described. The elliptic conformal transformation developed to map the hodograph domain onto an ellipse can be used to generate a conformal grid in the physical domain of a cascade of airfoils with open trailing edges with a single transformation. A grid generated with this transformation is shown for the Korn airfoil. Previously announced in STAR as N83-24474
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 950-956
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  • 105
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    Turbulent boundary-layer relaxation with application to skin-friction drag reduction (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Previously cited in issue 05, p. 584, Accession no. A83-16633
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 871
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  • 106
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    Surface pressures on a flat plate with dual jet configurations (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Previously cited in issue 17, p. 2456, Accession no. A83-38677
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 0021-8669); 21; 484-490
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  • 107
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    Two-dimensional wind-tunnel interference from measurements on two contours (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: This paper describes how wall-induced velocities near a model in a two-dimensional wind tunnel can be estimated from upwash distributions measured along two contours surrounding a model. The method is applicable to flows that can be represented by linear theory. It was derived by applying the Schwarz Integral Formula separately to the two contours and by exploiting the free-air relationship between upwashes along the contours. Advantages of the method are that only one flow quantity need by measured and no representation of the model is required. A weakness of the method is that it assumes streamwise interference velocity vanishes far upstream of the model. This method was applied to a simple theoretical model of flow in a solid-wall wind tunnel. The theoretical interference velocities and the velocities computed using the method were in excellent agreement. The method was then used to analyze experimental data acquired during adaptive-wall experiments at Ames Research Center. This analysis confirmed that the wall adjustments reduced wall-induced velocities near the model.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 0021-8669); 21; 414-419
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  • 108
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    Effect of boundary layers on solid walls in three-dimensional subsonic wind tunnels (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Previously cited in issue 05, p. 580, Accession no. A83-16553
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 365-371
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  • 109
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    Effect of a buried-wire gage on the separation bubble Numerical study (1984)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: A numerically simulated buried-wire separation gage is investigated with emphasis on its effect on the separation bubble. The conjugated problem of a supersonic, time-dependent, two-dimensional flowfield above a conductive solid wall with an embedded heat source is solved using implicit finite difference algorithms. Steady-state and transient cases were computed for different locations of the heat source within the bubble. Results show that by using a steady heat source, the flow direction near the wall can be detected, without distorting the flowfield, only if the source is located in regions where the bubble is thick (i.e., not too close to the separation). The flow direction near separation can be detected by using a temperature pulse at the solid/fluid interface with insignificant distortion of the flowfield.
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 1539-154
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  • 110
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    Rotary-wing aerodynamics. Volume 1 - Basic theories of rotor aerodynamics (With application to helicopters). Volume 2 Performance prediction of helicopters (2nd revised and enlarged edition) (1984)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: Basic theories of rotor aerodynamics are presented and applied to the performance prediction of helicopters. The very simple physicomathematical model of the rotor offered by momentum theory is addressed first, followed by the combined blade-element and momentum theory. Vortex theory is discussed, and a rotor blade is modeled by means of a vortex filament or vorticity surface. Considerations of airfoil sections suitable for rotors are examined. Detailed performance techniques for a single-rotor helicopter in hover, vertical ascent, and forward flight are described, and winged and tandem-rotor helicopter performance calculations are presented as extensions and modifications of single-rotor methodology. Computer data based on the vortex theory are compared with approximate results obtained from the simplified momentum theory and the blade element solution.
    Keywords: AERODYNAMICS
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  • 111
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    Resonance in flows with vortex sheets and edges (1984)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: It is shown that the vortex sheet in a slot between two semi-infinite plates does not admit incompressible resonant perturbations. The semi-infinite vortex sheet entering a duct does admit incompressible resonance. These results indicate that the vortex-sheet approximation is less useful for impinging shear flows than for non-impinging flows. They also suggest an important role of downstream vortical disturbances in resonant flows. The general solution for perturbations to flow with a vortex sheet and edges is written in terms of a Cauchy integral. Requirements on the behavior of this solution at edges and at downstream infinity fix the criteria for resonance.
    Keywords: AERODYNAMICS
    Type: Journal of Fluid Mechanics (ISSN 0022-1120); 145; 275-285
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  • 112
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    Generation of instability waves in flows separating from smooth surfaces (1984)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Description: This paper analyses the coupling between an imposed disturbance and an instability wave that propagates downstream on a shear layer which emanates from a separation point on a smooth surface. Since the wavelengths of the most-amplified instability waves will generally be small compared with the streamwise body dimensions, the analysis is restricted to this 'high-frequency' limit and the solution is obtained by using matched asymptotic expansions. An 'inner' solution, valid near the separation point, is matched onto an outer solution, which represents an instability wave on a slowly diverging mean flow. The analysis relates the amplitude of this instability to that of the imposed disturbance.
    Keywords: AERODYNAMICS
    Type: Journal of Fluid Mechanics (ISSN 0022-1120); 145; 71-94
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  • 113
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    A time-split finite-volume algorithm for three-dimensional flowfield simulation (1984)
    In:  Other Sources
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    Publication Date: 2011-08-19
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 1564-157
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  • 114
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    A fast viscous correction method for full-potential transonic wing analysis (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: An analysis of the transonic flowfield around a three-dimensional wing is carried out using a strip method. Attention is given to the boundary layer growth in the streamwise direction. A viscous correction technique is defined for the TWING code for solving the full potential equations. A viscous ramp at the base of a shock is superimposed on the boundary layer displacement thickness generated by an integral boundary layer method. A relationship is then obtained between the effective displacement thickness and a vertical component of the surface velocity, a transpirational boundary condition. The viscous correction is found to be unnecessary in weak shock conditions but gives a better shock position and pressure distribution in a strong shock condition when compared with data from an ONERA M6 airfoil and the Hinson and Burdges (1980) Wing A.
    Keywords: AERODYNAMICS
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  • 115
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    A new aerodynamic integral equation based on an acoustic formula in the time domain (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: An aerodynamic integral equation for bodies moving at transonic and supersonic speeds is presented. Based on a time-dependent acoustic formula for calculating the noise emanating from the outer portion of a propeller blade travelling at high speed (the Ffowcs Williams-Hawking formulation), the loading terms and a conventional thickness source terms are retained. Two surface and three line integrals are employed to solve an equation for the loading noise. The near-field term is regularized using the collapsing sphere approach to obtain semiconvergence on the blade surface. A singular integral equation is thereby derived for the unknown surface pressure, and is amenable to numerical solutions using Galerkin or collocation methods. The technique is useful for studying the nonuniform inflow to the propeller.
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 1337-134
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  • 116
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    Viscous-inviscid interactive procedure for rotational flow in cascades of airfoils (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 1281
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  • 117
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    Hypersonic interactions with surface mass transfer. I - Steady flow. II - Unsteady flow (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Previously cited in issue 15, p. 2345, Accession no. A82-31944
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 449-452
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  • 118
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    Impact of computers on aerodynamics research and development (1984)
    In:  Other Sources
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    Publication Date: 2011-08-18
    Description: Factors motivating the development of computational aerodynamics as a discipline are traced back to the limitations of the tools available to the aerodynamicist before the development of digital computers. Governing equations in exact and approximate forms are discussed together with approaches to their numerical solution. Example results obtained from the successively refined forms of the equations are presented and discussed, both in the context of levels of computer power required and the degree of the effect that their solution has on aerodynamic research and development. Factors pacing advances in computational aerodynamics are identified, including the amount of computational power required to take the next major step in the discipline. Finally, the Numerical Aerodynamic Simulation (NAS) Program - with its 1987 target of achieving a sustained computational rate of 1 billion floating-point operations per second operating on a memory of 240 million words - is briefly discussed in terms of its projected effect on the future of computational aerodynamics.
    Keywords: AERODYNAMICS
    Type: IEEE, Proceedings (ISSN 0018-9219); 72; 68-79
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  • 119
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    Design of advanced technology maneuvering aircraft models for the National Transonic Facility (1984)
    In:  CASI
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    Publication Date: 2017-10-02
    Description: The need for a large High-Reynolds-Number Transonic Wind Tunnel which will provide a tool to study phenomena sensitive to Reynolds number is discussed. The National Transonic Facility (NTF), is in the calibration phase and the desired capability. Its usefulness, however, will be influenced by the ability of industry to develop model systems capable of withstanding the severe operating environment of the facility so necessary to achieve full scale Reynolds number, without degradation of accuracy, and at reasonable cost. The feasibility of designing models of advanced aerodynamic technology maneuvering aircraft and to achieve full scale Reynolds number for each configuration in the NTF are determined. It is concluded that the facility does not offer the potential for making tunnel to full scale data correlations for this type of aircraft configuration.
    Keywords: AERODYNAMICS
    Type: AGARD Wind Tunnels and Testing Tech.; 15 p
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  • 120
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    Technology developments for laminar boundary layer control on subsonic transport aircraft (1984)
    In:  CASI
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    Publication Date: 2017-10-02
    Description: The development of laminar flow technology for commercial transport aircraft is discussed and illustrated in a review of studies undertaken in the NASA Aircraft Energy Efficiency (ACEE) program since 1976. The early history of laminar flow control (LFC) techniques and natural laminar flow (NLF) airfoil designs is traced, and the aims of ACEE are outlined. The application of slotted structures, composites, and electron beam perforated metals in supercritical LFC airfoils, wing panels, and leading edge systems is examined; wind tunnel and flight test results are summarized; studies of high altitude ice effects are described; and hybrid (LFC/NLF designs are characterized. Drawings and photographs are provided.
    Keywords: AERODYNAMICS
    Type: AGARD Improvement of Aerodynamic Performance Through Boundary Layer Control and High Lift Systems; 13 p
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  • 121
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    High Reynolds number tests of the cast 10-2/DOA2 transonic airfoil at ambient and cryogenic temperature conditions (1984)
    In:  CASI
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    Publication Date: 2017-10-02
    Description: The transonic airfoil CAST 10-2/DOA2 was investigated in several major transonic wind tunnels at Reynolds numbers ranging from Re=1 million six hundred thousand to forty five million at ambient and cryogenic temperature conditions. The main objective was to study the degree and extent of the effects of Reynolds on both the airfoil aerodynamic characteristics and the interference effects of various model-wind-tunnel systems. the initial analysis of the 10-2 airfoil results revealed appreciable real Reynolds number effects on this airfoil and, moreover, showed that wall interference, can be significantly affected by changes in Reynolds number thus appearing as true Reynolds number effects.
    Keywords: AERODYNAMICS
    Type: Agard Wind Tunnels and Testing Tech.; 13 p
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  • 122
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    Turbulent drag reduction research (1984)
    In:  CASI
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    Publication Date: 2017-10-02
    Description: Research in the area of turbulent drag reduction for attached flows is summarized. The most promising passive techniques utilize non-planar geometry. Of particular interest is the suitability of these devices for retrofit of existing vehicles. Five methods for reducing turbulent skin friction drag on bodies/fuselages are discussed. They are: (1) large-eddy breakup devices; (2) riblets; (3) slot injection optimization; (4) control of Emmons spot generation; and (5) relaminarization through massive suction. Except for the Emmons spot work these methods all indicate the possibility of sizable net reductions in skin friction for laboratory conditions.
    Keywords: AERODYNAMICS
    Type: AGARD Improvement of Aerodynamic Performance Through Boundary Layer Control and High Lift Systems; 13 p
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  • 123
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    Lattice arrangements for rapid convergence (1926)
    In:  CASI
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    Publication Date: 2016-06-07
    Description: A simple, systematic, optimized vortex-lattice approach is developed for application to lifting-surface problems. It affords a significant reduction in computational costs when compared to current methods. Extensive numerical experiments have been carried out on a wide variety of configurations, including wings with camber and single or multiple flaps, as well as high-lift jetflap systems. Rapid convergence as the number of spanwise or chordwise lattices are increased is assured, along with accurate answers. The results from this model should be useful not only in preliminary aircraft design but also, for example, as input for wake vortex roll-up studies and transonic flow calculations.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Res. Center Vortex-Lattice Utilization; p 325-342
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  • 124
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    Computational aerodynamics and supercomputers (1984)
    In:  Other Sources
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    Publication Date: 2018-12-01
    Description: Some of the progress in computational aerodynamics over the last decade is reviewed. The Numerical Aerodynamic Simulation Program objectives, computational goals, and implementation plans are described. Previously announced in STAR as N84-16139
    Keywords: AERODYNAMICS
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  • 125
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    Navier-Stokes solutions for two-dimensional subsonic base flow (1984)
    In:  Other Sources
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    Publication Date: 2018-12-01
    Description: Methods for determining the effects of mass injection from the trailing edge of a bluff body at low speeds and in transonic flow were numerically studied along with an unmodified blunt-based body to gain insight into the effects of vortex shedding on the base drag. The methodology used to obtain finite-difference solutions to the Navier-Stokes equations for subsonic compressible two-dimensional near-wake flows is presented. The effectiveness of an introduced outflow boundary condition which minimizes reflections back into the computational domain was demonstrated with the solution of a model vortex problem. Calculations of the near-wake flow past a circular cylinder were in excellent agreement with experimental data. Laminar-flow solutions for a blunt-based model with and without a base cavity and with mass injection into the wake agreed qualitatively with experimental observations. The drag reduction capability provided by such base modifications was demonstrated.
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  • 126
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    Experimental analyses of trailing edge flows (1984)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: An experimental study of several of the trailing edge and wake turbulence properties for a NACA 64A010 airfoil section was completed. The experiment was conducted at the Ohio State University Aeronautical and Astronautical Research Laboratory in the 6 inch X 22 inch transonic wind tunnel facility. The data were obtained at a free stream Mach number of 0.80 and a flow Reynolds number (based on chord length) of 5 million. The principle diagnostic tool was a dual-component laser Doppler velocimeter. The experimental data included surface static pressures, chordwise and vertical mean velocities, RMS turbulence intensities, local flow angles, and a determination of turbulence kinetic energy in the wake. Two angles of attack (0 and 2 degrees) were investigated. At these incidence angles, four flow field surveys were obtained ranging in position from the surface of the airfoil, between the transonic shock and the trailing edge, to the far-wake. At both angles of attack, the turbulence intensities and turbulence kinetic energy were observed to decay in the streamwise direction. In the far wake, for the non-lifting case, the turbulence intensities were nearly isotropic. For the two degree case, the horizontal component of the turbulence intensity was observed to be substantially higher than the vertical component.
    Keywords: AERODYNAMICS
    Type: NASA-CR-176904 , NAS 1.26:176904
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  • 127
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    A comparison of computer-generated lift and drag polars for a Wortmann airfoil to flight and wind tunnel results (1984)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: Computations of drag polars for a low-speed Wortmann sailplane airfoil are compared to both wind tunnel and flight results. Excellent correlation is shown to exist between computations and flight results except when separated flow regimes were encountered. Wind tunnel transition locations are shown to agree with computed predictions. Smoothness of the input coordinates to the PROFILE airfoil analysis computer program was found to be essential to obtain accurate comparisons of drag polars or transition location to either the flight or wind tunnel results.
    Keywords: AERODYNAMICS
    Type: NASA-CR-176963 , NAS 1.26:176963
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  • 128
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    Experimental Results for a Flapped Natural-laminar-flow Airfoil with High Lift/drag Ratio (1984)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: Experimental results have been obtained for a flapped natural-laminar-flow airfoil, NLF(1)-0414F, in the Langley Low-Turbulence Pressure Tunnel. The tests were conducted over a Mach number range from 0.05 to 0.40 and a chord Reynolds number range from about 3.0 x 10(6) to 22.0 x 10(6). The airfoil was designed for 0.70 chord laminar flow on both surfaces at a lift coefficient of 0.40, a Reynolds number of 10.0 x 10(6), and a Mach number of 0.40. A 0.125 chord simple flap was incorporated in the design to increase the low-drag, lift-coefficient range. Results were also obtained for a 0.20 chord split-flap deflected 60 deg.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85788 , NAS 1.15:85788
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  • 129
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    Theory of wing rock (1984)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: A theory is developed for predicting wing rock characteristics. From available data, it can be concluded that wing rock is triggered by flow asymmetries, developed by negative or weakly positive roll damping, and sustained by nonlinear aerodynamic roll damping. A new nonlinear aerodynamic model that includes all essential aerodynamic nonlinearities is developed. The Beecham-Titchener method is applied to obtain approximate analytic solutions for the amplitude and frequency of the limit cycle based on the three degree-of-freedom equations of motion. An iterative scheme is developed to calculate the average aerodynamic derivatives and dynamic characteristics at limit cycle conditions. Good agreement between theoretical and experimental results is obtained.
    Keywords: AERODYNAMICS
    Type: NASA-CR-176640 , NAS 1.26:176640 , CRINC-FRL-516-1
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  • 130
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    Nonlinear potential analysis techniques for supersonic-hypersonic aerodynamic design (1984)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: Approximate nonlinear inviscid theoretical techniques for predicting aerodynamic characteristics and surface pressures for relatively slender vehicles at supersonic and moderate hypersonic speeds were developed. Emphasis was placed on approaches that would be responsive to conceptual configuration design level of effort. Second order small disturbance and full potential theory was utilized to meet this objective. Numerical 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 experimental results for a variety of wing, body, and wing-body shapes.
    Keywords: AERODYNAMICS
    Type: NASA-CR-172299 , NAS 1.26:172299
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  • 131
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    Locally linearized longitudinal and lateral-directional aerodynamic stability and control derivaties for the X-29A aircraft (1984)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: The locally linearized longitudinal and lateral-directional aerodynamic stability and control derivatives for the X-29A aircraft were calculated for altitudes ranging from sea level to 50,000 ft, Mach numbers from 0.2 to 1.5, and angles of attack from -5 deg to 25 deg. Several other parameters were also calculated, including aerodynamic force and moment coefficients, control face position, normal acceleration, static margin, and reference angle of attack.
    Keywords: AERODYNAMICS
    Type: NASA-TM-84919 , H-1203 , NAS 1.26:84919
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    User's guide for NASCRIN: A vectorized code for calculating two-dimensional supersonic internal flow fields (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A computer program NASCRIN has been developed for analyzing two-dimensional flow fields in high-speed inlets. It solves the two-dimensional Euler or Navier-Stokes equations in conservation form by an explicit, two-step finite-difference method. An explicit-implicit method can also be used at the user's discretion for viscous flow calculations. For turbulent flow, an algebraic, two-layer eddy-viscosity model is used. The code is operational on the CDC CYBER 203 computer system and is highly vectorized to take full advantage of the vector-processing capability of the system. It is highly user oriented and is structured in such a way that for most supersonic flow problems, the user has to make only a few changes. Although the code is primarily written for supersonic internal flow, it can be used with suitable changes in the boundary conditions for a variety of other problems.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85708 , L-15678 , NAS 1.15:85708
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  • 133
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    Supersonic second order analysis and optimization program user's manual (1984)
    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 supersonic and moderate hypersonic speeds were developed. Emphasis was placed on approaches that would be responsive to conceptual configuration design level of effort. Second order small disturbance theory was utilized to meet this objective. Numerical codes were developed for analysis and design of relatively general three dimensional geometries. Results from the computations indicate good agreement with experimental results for a variety of wing, body, and wing-body shapes. Case computational time of one minute on a CDC 176 are typical for practical aircraft arrangement.
    Keywords: AERODYNAMICS
    Type: NASA-CR-172342 , NAS 1.26:172342
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  • 134
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    Static internal performance of single-expansion-ramp nozzles with thrust-vectoring capability up to 60 deg (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation has been conducted at static conditions (wind off) in the static-test facility of the Langley 16-Foot Transonic Tunnel. The effects of geometric thrust-vector angle, sidewall containment, ramp curvature, lower-flap lip angle, and ramp length on the internal performance of nonaxisymmetric single-expansion-ramp nozzles were investigated. Geometric thrust-vector angle was varied from -20 deg. to 60 deg., and nozzle pressure ratio was varied from 1.0 (jet off) to approximately 10.0.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2364 , L-15766 , NAS 1.60:2364
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  • 135
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    Effects of tail span and empennage arrangement on drag of a typical single-engine fighter aft end (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the Langley 16 foot Transonic Tunnel to determine the effects of tail span and empennage arrangement on drag of a single engine nozzle/afterbody model. Tests were conducted at Mach numbers from 0.50 to 1.20, nozzle pressures frm 1.0 (jet off) to 8.0, and angles of attack from -3 to 9 deg, depending upon Mach numbers. Three empennage arrangements (aft, staggered, and forward) were investigated with several different tail spans. The results of the investigation indicate that tail span and position have a significant effect on the drag at transonic speeds. Unfavorable tail interference was largely due to the outer portion of the tail surfaces. The inner portion near the nozzle and afterbody did little to increase drag other than surface skin friction. Tail positions forward of the nozzle generally had lower tail interference.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2352 , L-15742 , NAS 1.60:2352
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  • 136
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    Rotary balance data and analysis for the X-29A airplane for an angle-of-attack range of 0 deg to 90 deg (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The rotational aerodynamic characteristics are discussed for a 1/8 scale model of the X-29A airplane. The effects of rotation on the aerodynamics of the basic model were determined, as well as the influence of airplane components, various control deflections, and several forebody modifications. These data were measured using a rotary balance, over an angle of attack range of 0 to 90 deg, for clockwise and counter clockwise rotations covering an omega b/2V range of 0 to 0.4.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3747 , NAS 1.26:3747
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  • 137
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    Low speed rotary aerodynamics of F-18 configuration for 0 deg to 90 deg angle of attack: Test results and analysis (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Aerodynamic characteristics obtained in a rotational flow environment, utilizing a rotary balance located in the Langley Spin Tunnel, are discussed and presented in tabular form for a 1/10 scale F-18 airplane model. The rotational aerodynamic characteristics were established for the basic airplane, as well as the influence of control deflections and the contribution of airplane components, i.e., body, wing, leading edge extension, horizontal and vertical tails, on these characteristics up to 90 deg angle of attack. Spin equilibrium conditions predicted using the measured data are also presented and compared with spin model and full scale flight results.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3608 , NAS 1.26:3608
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  • 138
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    Large perturbation flow field analysis and simulation for supersonic inlets (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An analysis technique for simulation of supersonic mixed compression inlets with large flow field perturbations is presented. The approach is based upon a quasi-one-dimensional inviscid unsteady formulation which includes engineering models of unstart/restart, bleed, bypass, and geometry effects. Numerical solution of the governing time dependent equations of motion is accomplished through a shock capturing finite difference algorithm, of which five separate approaches are evaluated. Comparison with experimental supersonic wind tunnel data is presented to verify the present approach for a wide range of transient inlet flow conditions.
    Keywords: AERODYNAMICS
    Type: NASA-CR-174676 , NAS 1.26:174676
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  • 139
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    Spin-tunnel investigation of a 1/25-scale model of the General Dynamics F-16XL airplane (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A spin-tunnel investigation of the spin and recovery characteristics of a 1/25-scale model to the General Dynamics F-16XL aircraft was conducted in the Langley Spin Tunnel. Tests included erect and inverted spins at various symmetric and asymmetric loading conditions. The required size of an emergency spin-recovery parachute was determined.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85660 , L-15616 , NAS 1.15:85660
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  • 140
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    Supersonic, nonlinear, attached-flow wing design for high lift with experimental validation (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Results of the experimental validation are presented for the three dimensional cambered wing which was designed to achieve attached supercritical cross flow for lifting conditions typical of supersonic maneuver. The design point was a lift coefficient of 0.4 at Mach 1.62 and 12 deg angle of attack. Results from the nonlinear full potential method are presented to show the validity of the design process along with results from linear theory codes. Longitudinal force and moment data and static pressure data were obtained in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.58, 1.62, 1.66, 1.70, and 2.00 over an angle of attack range of 0 to 14 deg at a Reynolds number of 2.0 x 10 to the 6th power per foot. Oil flow photographs of the upper surface were obtained at M = 1.62 for alpha approx. = 8, 10, 12, and 14 deg.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2336 , L-15787 , NAS 1.60:2336
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  • 141
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    Experimental and analytical investigation of a freezing point depressant fluid ice protection system (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A glycol-exuding porous leading edge ice protection system was tested in the NASA Icing Research Tunnel. Stainless steel mesh, laser drilled titanium, and composite panels were tested on two general aviation wing sections. Two different glycol-water solutions were evaluated. Minimum glycol flow rates required for anti-icing were obtained as a function of angle of attack, liquid water content, volume median drop diameter, temperature, and velocity. Ice accretions formed after five minutes of icing were shed in three minutes or less using a glycol fluid flow equal to the anti-ice flow rate. Two methods of predicting anti-ice flow rates are presented and compared with a large experimental data base of anti-ice flow rates over a wide range of icing conditions. The first method presented in the ADS-4 document typically predicts flow rates lower than the experimental flow rates. The second method, originally published in 1983, typically predicts flow rates up to 25 percent higher than the experimental flow rates. This method proved to be more consistent between wing-panel configurations. Significant correlation coefficients between the predicted flow rates and the experimental flow rates ranged from .867 to .947.
    Keywords: AERODYNAMICS
    Type: NASA-CR-174758 , NAS 1.26:174758
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  • 142
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    Assessment of aerodynamic performance of V/STOL and STOVL fighter aircraft (1984)
    In:  CASI
    Details
    Publication Date: 2015-08-25
    Description: The aerodynamic performance of V/STOL and STOVL fighter/attack aircraft was assessed. Aerodynamic and propulsion/airframe integration activities are described and small-and large-scale research programs are considered. Uncertainties affecting aerodynamic performance that are associated with special configuration features resulting from the V/STOL requirement are addressed. Example uncertainties related to minimum drag, wave drag, high angle of attack characteristics, and power-induced effects. Engine design configurations from several aircraft manufacturers are reviewed.
    Keywords: AERODYNAMICS
    Type: AGARD Spec. Course on V(STOL Aerodyn.; 35 p
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  • 143
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    An experimental investigation of the parallel blade-vortex interaction (1984)
    In:  Other Sources
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    Publication Date: 2018-12-01
    Description: A scheme for investigating the parallel blade vortex interaction (BVI) has been designed and tested. The scheme involves setting a vortex generator upstream of a nonlifting rotor so that the vortex interacts with the blade at the forward azimuth. The method has revealed two propagation mechanisms: a type C shock propagation from the leading edge induced by the vortex at high tip speeds, and a rapid but continuous pressure pulse associated with the proximity of the vortex to the leading edge. The latter is thought to be the more important source. The effects of Mach number and vortex proximity are discussed.
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  • 144
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    A numerical study of the two- and three-dimensional unsteady Navier-Stokes equations in velocity-vorticity variables using compact difference schemes (1984)
    In:  Other Sources
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    Publication Date: 2018-12-01
    Description: A compact finite-difference approximation to the unsteady Navier-Stokes equations in velocity-vorticity variables is used to numerically simulate a number of flows. These include two-dimensional laminar flow of a vortex evolving over a flat plate with an embedded cavity, the unsteady flow over an elliptic cylinder, and aspects of the transient dynamics of the flow over a rearward facing step. The methodology required to extend the two-dimensional formulation to three-dimensions is presented.
    Keywords: AERODYNAMICS
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  • 145
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    Effects of nozzle design parameters on the extent of quiet test flow at Mach 3.5 (1984)
    In:  Other Sources
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    Publication Date: 2018-12-01
    Description: Tests results at the NASA Langley Research Center, involving a Mach 3.5 pilot quiet tunnel, have shown that laminar-layered nozzle walls improve boundary layer stability and reduce stream disturbance levels caused by eddy Mach wave radiation. This type of wall design is required to obtain transition Reynolds numbers on tests models as high as those previously observed in supersonic flight vehicles. The Mach 3.5 pilot nozzle wall boundary layers were tested for Tollmein-Schlichting and Goertler linear amplification, and, in an analysis of Goertler vortices in two axisymmetric Mach 5 nozzles, transition values were found to vary. These values were applied to several nozzles with similar throat heights but different expansion rates. Among the nozzles included in the study, a flat-wall radial flow nozzle and a proposed rod-wall nozzle were tested. For the highest test unit Reynolds number, it was determined that the nozzle wall surface finish should not exceed 0.3 micron. Oil flow studies have indicated that Goertler vortex disturbances were the dominant mechanism causing transition on the walls of the pilot nozzle.
    Keywords: AERODYNAMICS
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  • 146
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    B-70 Aircraft Study (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: This volume contains cost, schedule, and technical information on the following B-70 aircraft subsystems: air induction system, flight control, personnel accommodation and escape, alighting and arresting, mission and traffic control, flight indication, test instrumentation, and installation, checkout, and pre-flight.
    Keywords: AERODYNAMICS
    Type: NASA-CR-115705 , SD-72-SH-0003-VOL-4 , NAS 1.26:115705 , JSC-CN-29834
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  • 147
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    B-70 Aircraft Study (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Volume 2 of the final report on the B-70 aircraft study is presented here. The B-70 Program, at the onset, was a full weapon system capable of sustained Mach 3 flight for the major portion of its design missions. The weapon system was to enter the SAC inventory as an RS-70 with the first intercontinental resonnaissance/bomber wing scheduled to go operational in July, 1964. After several redirections, a two XB-70 air vehicle program emerged with its prime objective being to demonstrate the technical feasibility of sustained Mach 3 flight. This section describes the original Weapon System 110A concepts, the evolution of the RS-70 design, and the XB-70 air vehicles which demonstrated the design, fabrication, and technical feasibility of long range Mach 3 flights at high altitude. The data presented shows that a very large step forward in the state-of-the-art of manned aircraft design was achieved during the B-70 development program and that advances were made and incorporated in every area, including design, materials application, and manufacturing techniques.
    Keywords: AERODYNAMICS
    Type: NASA-CR-115703 , SD-72-SH-0003-VOL-2 , NAS 1.26:115703 , JSC-CN-29832
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  • 148
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    B-70 Aircraft Study (1972)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: This Phase 2 final report for the B-70 aircraft study contains the data location matrix, which provides a summary of the major cost, schedule, and technical items provided in the report; work breakdown structure; cost definitions; and B-70 program level summary data. The Phase 2 objective was to provide the B-70 aircraft data in accordance with the approved study plan. Several minor modifications to the original plan have been made as the result of the Phase 2 effort.
    Keywords: AERODYNAMICS
    Type: NASA-CR-115702 , SD-72-SH-0003-VOL-1 , NAS 1.26:115702 , JSC-CN-29818
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  • 149
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    Aerodynamic heating rate distributions induced by trailing edge controls on hypersonic aircraft configurations at Mach 8 (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Aerodynamic surface heating rate distributions in three dimensional shock wave boundary layer interaction flow regions are presented for a generic set of model configurations representative of the aft portion of hypersonic aircraft. Heat transfer data were obtained using the phase change coating technique (paint) and, at particular spanwise and streamwise stations for sample cases, by the thin wall transient temperature technique (thermocouples). Surface oil flow patterns are also shown. The good accuracy of the detailed heat transfer data, as attested in part by their repeatability, is attributable partially to the comparatively high temperature potential of the NASA-Langley Mach 8 Variable Density Tunnel. The data are well suited to help guide heating analyses of Mach 8 aircraft, and should be considered in formulating improvements to empiric analytic methods for calculating heat transfer rate coefficient distributions.
    Keywords: AERODYNAMICS
    Type: NASA-TM-87453 , RM-799 , NAS 1.15:87453
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  • 150
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    Real-gas simulation for the Shuttle Orbiter and planetary entry configurations including flight results (1984)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: By testing configurations in a gas (like CF4) which can produce high normal-shock density ratios, such as those encountered during hypersonic entry, certain aspects of real-gas effects can be simulated. Results from force-moment, shock-shape and oil flow visualization tests are presented for both the Shuttle Orbiter and a 45 deg sphere-cone in CF4 and air at M = 6, and comparisons are made with flight results. Pitching-moment coefficients measured on a Shuttle Orbiter model in CF4 showed a nose-up increment, compared with air results, that was almost identical to the difference between preflight predictions and flight in the high hypersonic regime. The drag coefficient measured in CF4 on the 45 deg sphere-cone, which is the same configuration used on the forebody of the Pioneer Venus entry vehicles, showed excellent agreement with flight data at M = 6.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0489
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  • 151
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    Boundary layer transition effects on flow separation around V/STOL engine inlets at high incidence (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Numerical methods for calculating laminar and turbulent boundary layer development around vertical-short take off and landing engine inlets at high incidence angles are investigated. Various transition models were compared and evaluated in calculations off flow separation bound inside the inlet. Results of the transition effects on the boundary layer characteristics at onset of separation for two types of engine inlet geometries are presented. Some of the numerical results are compared with existing wind-tunnel test data for scaled inlet models to demonstrate the effects of transition models in the numerical scheme. The effects of transition modeling on the boundary layer development are illustrated for typical engine operating conditions.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0432
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  • 152
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    Euler and Navier-Stokes solutions for supersonic shear flow past a circular cylinder (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Euler and Navier-Stokes solutions of the supersonic shear flow past a circular cylinder are obtained. These solutions are used to study the basic flow structure around the cylinder. Both the inviscid and viscous calculations show the formation of a large recirculating flow region around the front stagnation point. The calculations further show that the overall size of the recirculating region is approximately the same for the Euler and Navier-Stokes solutions but the inside structure is quite different. The inviscid flow shows only one vortex whereas the viscous flow shows two vortices inside the recirculating flow region. The inner vortex in the Navier-Stokes solution is formed primarily due to the viscous effects near the body surface and its size depends upon the Reynolds number. It is found that with increasing Reynolds number, the inner vortex diminishes in size and the Navier-Stokes solution asymptotically approaches the Euler solution. These results indicate that the Euler equations may correctly predict certain high Reynolds number separation phenomenon in flows with natural inviscid vorticity source.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0339
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  • 153
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    A large-scale investigation of V/STOL ground effects (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A large-scale ground-effects test of a single jet in hover was conducted as a first-case study for future tests to provide the V/STOL community with an improved data base. The objectives for this single-jet hover test were to (1) document the jet characteristics and then (2) gather the associated force data. These data are then compared with results obtained from existing prediction methods. A conically convergent nozzle was mounted to a turbojet engine, and an 8-ft-diam suckdown-plate model measured the lift-loss forces in ground effect. Jet-exit characteristics (pressure profile, temperature, turbulence, etc.) are documented for several nozzle pressure ratios. Characteristics that may give rise to scale effects are discussed. Results from this first study indicate that small-scale tests, and current prediction methods, will lead to significant errors in the lift-loss estimation of a single-jet configured aircraft, hovering in ground effect.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0336
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  • 154
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    Data from tests of a R4 airfoil in the Langley 0.3-meter transonic cryogenic tunnel (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Aerodynamic data for the DFVLR R4 airfoil are presented in both graphic and tabular form. The R4 was tested in the Langley 0.3-Meter Transonic Cryogenic Tunnel (TCT) at Mach number from 0.60 to 0.78 at angles of attack from -2.0 to 8.0 degrees. The airfoil was tested at Reynolds numbers of 4, 6, 10, 15, 30, and 40 million based on the 152.32 mm chord.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85739 , NAS 1.15:85739
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  • 155
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    A conservative treatment of zonal boundaries for Euler equation calculations (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Finite-difference calculations require the generation of a grid for the region of interest. A zonal approach, wherein the given region is subdivided into zones and the grid for each zone is generated independently, makes the grid-generation process for complicated topologies and for regions requiring selective grid refinement a fairly simple task. This approach results in new boundaries within the given region, that is, zonal boundaries at the interfaces of the various zones. The zonal-boundary scheme (the integration scheme used to update the points on the zonal boundary) for the Euler equations must be conservative, accurate, stable, and applicable to general curvilinear coordinate systems. A zonal-boundary scheme with these desirable properties is developed in this study. The scheme is designed for explicit, first-order-accurate integration schemes but can be modified to accommodate second-order-accurate explicit and implicit integration schemes. Results for inviscid flow, including supersonic flow over a cylinder, blast-wave diffraction by a ramp, and one-dimensional shock-tube flow are obtained on zonal grids. The conservative nature of the zonal-boundary scheme permits the smooth transition of the discontinuities associated with these flows from one zone to another. The calculations also demonstrate the continuity of contour lines across zonal boundaries that can be achieved with the present zonal scheme.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0164
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  • 156
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    Prediction of vortex lift on interacting delta wings in incompressible flow (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Vortex flow modeling is used to calculate the steady inviscid incompressible flow past one, two and three delta wing configurations. The wings are modeled with vortex lattices and the leading and trailing-edge sheets are modeled by segmented straight vortex filaments. Aerodynamic characteristics are obtained for a range of geometry and angle of attack.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0136
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  • 157
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    Application of a full potential method for computation of three-dimensional supersonic flows (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A nonlinear aerodynamic analysis technique based on the full potential equation in conservative form has been modified to permit treatment of supersonic flows with embedded subsonic regions (typically near the fuselage-canopy juncture and the wing leading edge). Solution procedures for the equations do not require any specific form of geometry or physical grid system. This results in the capability to analyze easily very complex geometries provided the posed problem lies within the isentropic restrictions of the full potential theory. Characteristic signal propagation theory is used to monitor the type dependent flow and a conservative switching scheme is employed to transition from the supersonic marching algorithm to a subsonic relaxation procedure and vice versa. An implicit approximate factorization scheme is used to solve the finite-difference equations. These modifications now permit analysis of fully three-dimensional flowfields including the interference effects due to lifting surface wakes. Improved grid generation capability allows analysis of complete complex aircraft geometries (fuselage, wing, tail, wing wake, and tail wake). Results are presented showing very good correlations with experimental surface pressure data and aerodynamic force data at both design and off-design operating points. Configurations examined include several waverider concepts, an arrow wing-body with wake, an advanced tactical fighter concept, and a fighter forebody-canard configuration.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0139
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  • 158
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    The influence of leading-edge load alleviation on supersonic wing design (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A theoretical and experimental program to assess the effect of leading-edge load constraints on wing design and performance was conducted. For a planform characterized by a highly swept leading edge on the inboard region, linear theory was used to design camber surfaces which produced minimum drag-due-to-lift at the design lift coefficient of 0.08 and a design Mach number of 2.4. In an effort to delay the formation of leading edge vortices which often occur on highly swept wings, two approaches were used in the design criteria to limit the loadings on the leading edge. One wing was constrained to have the normal Mach number less than one everywhere along the leading edge and the second wing was constrained to have a pressure coefficient of zero on the leading edge. Force tests were run on the two constrained wings, on a flat reference wing and on an optimized wing with no leading edge constraints. All wings had identical planforms and thicknesses and were tested over a range of Mach numbers from 1.8 to 2.8 and a range in angles of attack from -5 deg to 8 deg. A comparison of the experimental performance of these four models is shown. Correlations of these results with theoretical predictions and flow visualization photographs are also included.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0138
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  • 159
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    Numerical Investigation of unsteady inlet flow fields (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The flow field within an unsteady, two-dimensional inlet is studied numerically, using a two dimensional Navier Stokes and a one-dimensional inviscid model. Unsteadiness is introduced by varying the outflow pressure boundary condition. The cases considered include outflow pressure variations which were a single pressure pulse, a rapid increase and a sine function. The amplitude of the imposed exit plane pressure disturbance varied between 1 percent and 20 percent of the mean exit pressure. At the higher levels of pressure fluctuation, the viscous flow field results bore little resemblance to the inviscid ones. The viscous solution included such phenomena as shock trains and bifurcating separation pockets. The induced velocity at the outflow plane predicted by the viscous model differs significantly from accoustical theory or small perturbation results.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0031
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    Three-dimensional turbulent boundary-layer development on a fan rotor blade (1984)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Previously cited in issue 15, p. 2347, Accession no. A82-31965
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 83-89
    Format: text
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  • 161
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    A review of some Reynolds number effects related to bodies at high angles of attack (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A review of some effects of Reynolds number on selected aerodynamic characteristics of two- and three-dimensional bodies of various cross sections in relation to fuselages at high angles of attack at subsonic and transonic speeds is presented. Emphasis is placed on the Reynolds number ranges above the subcritical and angles of attack where lee side vortex flow or unsteady wake type flows predominate. Lists of references, arranged in subject categories, are presented with emphasis on those which include data over a reasonable Reynolds number range. Selected Reynolds number data representative of various aerodynamic flows around bodies are presented and analyzed and some effects of these flows on fuselage aerodynamic parameters are discussed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3809 , NAS 1.26:3809
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  • 162
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    Calculation of the 3-D viscous flow at the endwall leading edge region of an axial annular turbine cascade (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A three-dimensional viscous computer code (VANS/MD) was employed to calculate the turbulent flow field at the end wall leading edge region of a 20 inch axial annular turbine cascade. The initial boundary layer roll-up and formation of the end wall vortices were computed at the vane leading edge. The calculated flow field was found to be periodic with a frequency of approximately 1600 Hz. The calculated size of the separation region for the hub endwall vortex compared favorably with measured endwall oil traces. In an effort to determine the effects of the turbulence model on the calculated unsteadiness, a laminar calculation was made. The periodic nature of the calculated flow field persisted with the frequency essentially unchanged.
    Keywords: AERODYNAMICS
    Type: NASA-CR-168275 , NAS 1.26:168275 , USAAVSCOM-TR-84-C-3 , AD-A145641
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  • 163
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    Comparison of the full potential and Euler formulations for computing transonic airfoil flows (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A quantitative comparison between the Euler and full potential formulations with respect to speed and accuracy is presented. The robustness of the codes used is tested by a number of transonic airfoil cases. The computed results are from four transonic airfoil computer codes. The full potential codes use fully implicit iteration algorithms. The first Euler code uses a fully implicit ADI iteration scheme. The second Euler code uses an explicit Runge Kutta time stepping algorithm which is enhanced by a multigrid convergence acceleration scheme. Quantitative comparisons are made using various plots of lift coefficient versus the average mesh spacing along the airfoil. Besides yielding an asymptotic limit to the lift coefficient, these results also demonstrate the truncation error behavior of the various codes. Quantitative conclusions regarding the full potential and Euler formulations with respect to accuracy, speed, and robustness can be presented.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85983 , A-9816 , NAS 1.15:85983
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  • 164
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    Potential application of artificial concepts to aerodynamic simulation (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The concept of artificial intelligence as it applies to computational fluid dynamics simulation is investigated. How expert systems can be adapted to speed the numerical aerodynamic simulation process is also examined. A proposed expert grid generation system is briefly described which, given flow parameters, configuration geometry, and simulation constraints, uses knowledge about the discretization process to determine grid point coordinates, computational surface information, and zonal interface parameters.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85976 , A-9798 , NAS 1.15:85976
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  • 165
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    Unsteady transonic aerodynamic and aeroelastic calculations about airfoils and wings (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The development and application of transonic small disturbance codes for computing two dimensional flows, using the code ATRAN2, and for computing three dimensional flows, using the code ATRAN3S, are described. Calculated and experimental results are compared for unsteady flows about airfoils and wings, including several of the cases from the AGARD Standard Aeroelastic Configurations. In two dimensions, the results include AGARD priority cases for the NACA 64A006, NACA 64A010, NACA 0012, and MBB-A3 airfoils. In three dimensions, the results include flows about the F-5 wing, a typical wing, and the AGARD rectangular wings. Viscous corrections are included in some calculations, including those for the AGARD rectangular wing. For several cases, the aerodynamic and aeroelastic calculations are compared with experimental results.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85986 , A-9822 , NAS 1.15:85986
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  • 166
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    Development and application of an analysis of axisymmetric body effects on helicopter rotor aerodynamics using modified slender body theory (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A computationally efficient body analysis designed to couple with a comprehensive helicopter analysis is developed in order to calculate the body-induced aerodynamic effects on rotor performance and loads. A modified slender body theory is used as the body model. With the objective of demonstrating the accuracy, efficiency, and application of the method, the analysis at this stage is restricted to axisymmetric bodies at zero angle of attack. By comparing with results from an exact analysis for simple body shapes, it is found that the modified slender body theory provides an accurate potential flow solution for moderately thick bodies, with only a 10%-20% increase in computational effort over that of an isolated rotor analysis. The computational ease of this method provides a means for routine assessment of body-induced effects on a rotor. Results are given for several configurations that typify those being used in the Ames 40- by 80-Foot Wind Tunnel and in the rotor-body aerodynamic interference tests being conducted at Ames. A rotor-hybrid airship configuration is also analyzed.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85934 , A-9704 , NAS 1.15:85934
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  • 167
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    Computational aerodynamics and artificial intelligence (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The general principles of artificial intelligence are reviewed and speculations are made concerning how knowledge based systems can accelerate the process of acquiring new knowledge in aerodynamics, how computational fluid dynamics may use expert systems, and how expert systems may speed the design and development process. In addition, the anatomy of an idealized expert system called AERODYNAMICIST is discussed. Resource requirements for using artificial intelligence in computational fluid dynamics and aerodynamics are examined. Three main conclusions are presented. First, there are two related aspects of computational aerodynamics: reasoning and calculating. Second, a substantial portion of reasoning can be achieved with artificial intelligence. It offers the opportunity of using computers as reasoning machines to set the stage for efficient calculating. Third, expert systems are likely to be new assets of institutions involved in aeronautics for various tasks of computational aerodynamics.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85994 , A-9807 , NAS 1.15:85994
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  • 168
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    Acceleration to a steady state for the Euler equations (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A multi-stage Runge-Kutta method is analyzed for solving the Euler equations exterior to an airfoil. Highly subsonic, transonic and supersonic flows are evaluated. Various techniques for accelerating the convergence to a steady state are introduced and analyzed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-172398 , NAS 1.26:172398 , ICASE-84-32
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  • 169
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    Interference drag in a simulated wing-fuselage juncture (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The interference drag in a wing fuselage juncture as simulated by a flat plate and a body of constant thickness having a 1.5:1 elliptical leading edge is evaluated experimentally. The experimental measurements consist of mean velocity data taken with a hot wire at a streamwise location corresponding to 16 body widths downstream of the body leading edge. From these data, the interference drag is determined by calculating the total momentum deficit (momentum area) in the juncture and also in the two dimensional turbulent boundary layers on the flat plate and body at locations sufficiently far from the juncture flow effect. The interference drag caused by the juncture drag as measured at this particular streamwise station is -3% of the total drag due to the flat plate and body boundary layers in isolation. If the body is considered to be a wing having a chord and span equal to 16 body widths, the interference drag due to the juncture is only -1% of the frictional drag of one surface of such a wing.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3811 , NAS 1.26:3811
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  • 170
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    A numerical simulation of three-dimensional flow in an adaptive wall wind tunnel (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Numerical simulations of three dimensional flows in a prototype adaptive wall wind tunnel are conducted at the Mach number of 0.6 to investigate: (1) wind tunnel wall interference, (2) active streamline control by varying air removal or injection along the walls, and (3) to develop a method for establishing wall boundary conditions for interference free flows. Wind tunnel wall interference could be controlled by using only the vertical velocity components. For the configuration tested, interference free flow with solid sidewalls can be approximated by using only floor and ceiling blowing/suction.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2351 , A-9622 , NAS 1.60:2351
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  • 171
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    Three-dimensional transonic potential flow about complex 3-dimensional configurations (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An analysis has been developed and a computer code written to predict three-dimensional subsonic or transonic potential flow fields about lifting or nonlifting configurations. Possible condfigurations include inlets, nacelles, nacelles with ground planes, S-ducts, turboprop nacelles, wings, and wing-pylon-nacelle combinations. The solution of the full partial differential equation for compressible potential flow written in terms of a velocity potential is obtained using finite differences, line relaxation, and multigrid. The analysis uses either a cylindrical or Cartesian coordinate system. The computational mesh is not body fitted. The analysis has been programmed in FORTRAN for both the CDC CYBER 203 and the CRAY-1 computers. Comparisons of computed results with experimental measurement are presented. Descriptions of the program input and output formats are included.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3814 , NAS 1.26:3814 , D6-52329
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  • 172
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    A preliminary design study on an acoustic muffler for the laminar flow transition research apparatus (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An acoustic muffler design of a research tool for studying laminar flow and the mechanisms of transition, the Laminar Flow and Transition Research Apparatus (LFTRA) is investigated. Since the presence of acoustic pressure fluctuations is known to affect transition, low background noise levels in the test section of the LFTRA are mandatory. The difficulties and tradeoffs of various muffler design concepts are discussed and the most promising candidates are emphasized.
    Keywords: AERODYNAMICS
    Type: NASA-CR-172374 , NAS 1.26:172374
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  • 173
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    Calculation of unsteady aerodynamics for four AGARD standard aeroelastic configurations (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Calculated unsteady aerodynamic characteristics for four Advisory Group for Aeronautical Research Development (AGARD) standard aeroelastic two-dimensional airfoils and for one of the AGARD three-dimensional wings are reported. Calculations were made using the finite-difference codes XTRAN2L (two-dimensional flow) and XTRAN3S (three-dimensional flow) which solve the transonic small disturbance potential equations. Results are given for the 36 AGARD cases for the NACA 64A006, NACA 64A010, and NLR 7301 airfoils with experimental comparisons for most of these cases. Additionally, six of the MBB-A3 airfoil cases are included. Finally, results are given for three of the cases for the rectangular wing.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85817 , NAS 1.15:85817
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  • 174
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    Numerical simulation of boundary-layer transition (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The transition to turbulence in boundary layers was investigated by direct numerical solution of the nonlinear, three-dimensional, incompressible Navier-Stokes equations in the half-infinite domain over a flat plate. Periodicity was imposed in the streamwise and spanwise directions. A body force was applied to approximate the effect of a nonparallel mean flow. The numerical method was spectra, based on Fourier series and Jacobi polynomials, and used divergence-free basis functions. Extremely rapid convergence was obtained when solving the linear Orr-Sommerfeld equation. The early nonlinear and three-dimensional stages of transition, in a boundary layer disturbed by a vibrating ribbon, were successfully simulated. Excellent qualitative agreement was observed with either experiments or weakly nonlinear theories. In particular, the breakdown pattern was staggered or nonstaggered depending on the disturbance amplitude.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85984 , A-9796 , NAS 1.15:85984
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  • 175
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    A practical adaptive-grid method for complex fluid-flow problems (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A practical solution, adaptive-grid method utilizing a tension and torsion spring analogy is proposed for multidimensional fluid flow problems. The tension spring, which connects adjacent grid points to each other, controls grid spacings. The torsion spring, which is attached to each grid node, controls inclinations of coordinate lines and grid skewness. A marching procedure was used that results in a simple tridiagonal system of equations at each coordinate line to determine grid-point distribution. Multidirectional adaptation is achieved by successive applications of one-dimensional adaptation. Examples of applications for axisymmetric afterbody flow fields and two dimensional transonic airfoil flow fields are shown.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85989 , A-9803 , NAS 1.15:85989
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  • 176
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    Shapes for rotating airfoils (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An airfoil which has particular application to the blade or blades of rotor aircraft and aircraft propellers is presented. The airfoil thickness distribution, camber and leading edge radius are shaped to locate the airfoil crest at a more aft position along the chord, and to increase the freestream Mach number at which sonic flow is attained at the airfoil crest. The reduced slope of the airfoil causes a reduction in velocity at the airfoil crest at lift coefficients from zero to the maximum lift coefficient. The leading edge radius is adjusted so that the maximum local Mach number at 1.25 percent chord and at the designed maximum lift coefficient is limited to about 0.48 when the Mach number normal to the leading edge is approximately 0.20. The lower surface leading edge radius is shaped so that the maximum local Mach number at the leading edge is limited to about 0.29 when the Mach number normal to the leading edge is approximately 0.20. The drag divergence Mach number associated with the airfoil is moved to a higher Mach number over a range of lift coefficients resulting in superior aircraft performance.
    Keywords: AERODYNAMICS
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  • 177
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    Recent modifications and calibration of the Langley low-turbulence pressure tunnel (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Modifications to the Langley Low-Turbulence Pressure Tunnel are presented and a calibration of the mean flow parameters in the test section is provided. Also included are the operational capability of the tunnel and typical test results for both single-element and multi-element airfoils. Modifications to the facility consisted of the following: replacement of the original cooling coils and antiturbulence screens and addition of a tunnel-shell heating system, a two dimensional model-support and force-balance system, a sidewall boundary layer control system, a remote-controlled survey apparatus, and a new data acquisition system. A calibration of the mean flow parameters in the test section was conducted over the complete operational range of the tunnel. The calibration included dynamic-pressure measurements, Mach number distributions, flow-angularity measurements, boundary-layer characteristics, and total-pressure profiles. In addition, test-section turbulence measurements made after the tunnel modifications have been included with these calibration data to show a comparison of existing turbulence levels with data obtained for the facility in 1941 with the original screen installation.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2328 , L-15728 , NAS 1.60:2328
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  • 178
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    V/STOL concepts in the United States: Past, present, and future (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Nonhelicopter types of V/STOL aircraft developed in the United States are reviewed, and some lessons learned from a selected number of concepts are highlighted. The AV-8B, which was developed by modifications to the British Harrier is the only current concept examined. Configurations proposed for the future subsonic, multimissing aircraft and the future supersonic fighter/attack aircraft are described. Emphasis is on these supersonic concepts.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85938 , A-9695 , NAS 1.15:85938
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  • 179
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    ALESEP: A computer program for the analysis of airfoil leading edge separation bubbles (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The ALESEP program for the analysis of the inviscid/viscous interaction which occurs due to the presence of a closed laminar transitional separation bubble on an airflow is presented. The ALESEP code provides a iterative solution of the boundary layer equations expressed in an inverse formulation coupled to a Cauchy integral representation of the inviscid flow. This interaction analysis is treated as a local perturbation to a known solution obtained from a global airfoil analysis. Part of the required input to the ALESEP code are the reference displacement thickness and tangential velocity distributions. Special windward differencing may be used in the reversed flow regions of the separation bubble to accurately account for the flow direction in the discretization of the streamwise convection of momentum. The ALESEP code contains a forced transition model based on a streamwise intermittency function and a natural transition model based on a solution of the integral form of the turbulent kinetic energy equation. Instructions for the input/output, and program usage are presented.
    Keywords: AERODYNAMICS
    Type: NASA-CR-172310 , NAS 1.26:172310
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  • 180
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    The design and operational development of self-streamlining 2-dimensional flexible walled test sections (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Self streamlining two dimensional flexible walled test sections eliminate the uncertainties found in data from conventional test sections particularly at transonic speeds. The test section sidewalls are rigid, while the floor and ceiling are flexible and are positioned to streamline shapes by a system of jacks, without reference to the model. The walls are therefore self streamlining. Data are taken from the model when the walls are good streamlines such that the inevitable residual wall induced interference is acceptably small and correctable. Successful two dimensional validation testing at low speeds has led to the development of a new transonic flexible walled test section. Tunnel setting times are minimized by the development of a rapid wall setting strategy coupled with on line computer control of wall shapes using motorized jacks. Two dimensional validation testing using symmetric and cambered aerofoils in the Mach number range up to about 0.85 where the walls are just supercritical, shows good agreement with reference data using small height-chord ratios between 1.5 and unity.
    Keywords: AERODYNAMICS
    Type: NASA-CR-172328 , NAS 1.26:172328
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  • 181
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    Inlet flowfield investigation. Part 2: Computation of the flow about a supercruise forebody at supersonic speeds (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A numerical procedure which solves the parabolized Navier-Stokes (PNS) equations on a body fitted mesh was used to compute the flow about the forebody of an advanced tactical supercruise fighter configuration in an effort to explore the use of a PNS method for design of supersonic cruise forebody geometries. Forebody flow fields were computed at Mach numbers of 1.5, 2.0, and 2.5, and at angles-of-attack of 0 deg, 4 deg, and 8 deg. at each Mach number. Computed results are presented at several body stations and include contour plots of Mach number, total pressure, upwash angle, sidewash angle and cross-plane velocity. The computational analysis procedure was found reliable for evaluating forebody flow fields of advanced aircraft configurations for flight conditions where the vortex shed from the wing leading edge is not a dominant flow phenomenon. Static pressure distributions and boundary layer profiles on the forebody and wing were surveyed in a wind tunnel test, and the analytical results are compared to the data. The current status of the parabolized flow flow field code is described along with desirable improvements in the code.
    Keywords: AERODYNAMICS
    Type: NASA-CR-172315 , NAS 1.26:172315 , D180-27939-2
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  • 182
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    Documentation of ice shapes on the main rotor of a UH-1H helicopter in hover (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A helicopter icing flight test program in the hover mode was conducted with a UH-1H aircraft. The ice formations were documented after landing by means of silicone rubber molds, stereo photography and outline tracings for later use in aerodynamic analyses. The documentation techniques are described and the results presented for a typical flight.
    Keywords: AERODYNAMICS
    Type: NASA-CR-168332 , NAS 1.26:168332
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  • 183
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    Tests on a CAST 7 two-dimensional airfoil in a streamlining test section (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A unique opportunity has arisen to test one and the same airfoil model of CAST-7 section in two wind tunnels having adaptive walled test sections. The tunnels are very similar in terms of size and the available range of test conditions, but differ principally in their wall setting algorithms. Detailed data from the tests of the model in the Southampton tunnel, are included with comparisons between various sources of data indicating that both adaptive walled test sections provide low interference test conditions.
    Keywords: AERODYNAMICS
    Type: NASA-CR-172291 , NAS 1.26:172291
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  • 184
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    Computational aerodynamics and supercomputers (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Some of the progress in computational aerodynamics over the last decade is reviewed. The Numerical Aerodynamic Simulation Program objectives, computational goals, and implementation plans are described.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85887 , A-9583 , NAS 1.15:85887
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  • 185
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    Comparison of frequency-domain and time-domain rotorcraft vibration control methods (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Active control of rotor-induced vibration in rotorcraft has received significant attention recently. Two classes of techniques have been proposed. The more developed approach works with harmonic analysis of measured time histories and is called the frequency-domain approach. The more recent approach computes the control input directly using the measured time history data and is called the time-domain approach. The report summarizes the results of a theoretical investigation to compare the two approaches. Five specific areas were addressed: (1) techniques to derive models needed for control design (system identification methods), (2) robustness with respect to errors, (3) transient response, (4) susceptibility to noise, and (5) implementation difficulties. The system identification methods are more difficult for the time-domain models. The time-domain approach is more robust (e.g., has higher gain and phase margins) than the frequency-domain approach. It might thus be possible to avoid doing real-time system identification in the time-domain approach by storing models at a number of flight conditions. The most significant error source is the variation in open-loop vibrations caused by pilot inputs, maneuvers or gusts. The implementation requirements are similar except that the time-domain approach can be much simpler to implement if real-time system identification were not necessary.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166570 , NAS 1.26:166570
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  • 186
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    Flow-field measurements on an airfoil with an oscillating trailing-edge using holographic interferometry (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Holographic interferometry data were acquired on an NACA 64A010 airfoil with an oscillating flap. The airfoil was installed in the Ames 11-Foot Transonic Wind Tunnel between splitter plates. Recordings were made at discrete phase angles of the oscillation. The interferometry results provided detailed flow visualization of the shock boundary-layer interaction and the separated flow. Quantitative results were extracted from the interferograms to produce pressure data. These results were compared to the surface pressures obtained with the surface pressure taps. Excellent agreement was found for low angles of incidence. At larger angles of incidence, the flow had greater three-dimensionality, and the results were not in good agreement in some regions of the flow field. Mach contours were traced for representative flow conditions. Wake profiles were also obtained using the assumption of constant pressure across the wake and the Crocco relationship.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166604 , NAS 1.26:166604
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  • 187
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    Application of the Green's function method for 2- and 3-dimensional steady transonic flows (1984)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A Time-Domain Green's function method for the nonlinear time-dependent three-dimensional aerodynamic potential equation is presented. The Green's theorem is being used to transform the partial differential equation into an integro-differential-delay equation. Finite-element and finite-difference methods are employed for the spatial and time discretizations to approximate the integral equation by a system of differential-delay equations. Solution may be obtained by solving for this nonlinear simultaneous system of equations in time. This paper discusses the application of the method to the Transonic Small Disturbance Equation and numerical results for lifting and nonlifting airfoils and wings in steady flows are presented.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0425
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  • 188
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    A method for modeling finite-core vortices in wake-flow calculations (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A numerical method for computing nonplanar vortex wakes represented by finite-core vortices is presented. The approach solves for the velocity on an Eulerian grid, using standard finite-difference techniques; the vortex wake is tracked by Lagrangian methods. In this method, the distribution of continuous vorticity in the wake is replaced by a group of discrete vortices. An axially symmetric distribution of vorticity about the center of each discrete vortex is used to represent the finite-core model. Two distributions of vorticity, or core models, are investigated: a finite distribution of vorticity represented by a third-order polynomial, and a continuous distribution of vorticity throughout the wake. The method provides for a vortex-core model that is insensitive to the mesh spacing. Results for a simplified case are presented. Computed results for the roll-up of a vortex wake generated by wings with different spanwise load distributions are presented; contour plots of the flow-field velocities are included; and comparisons are made of the computed flow-field velocities with experimentally measured velocities.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0417
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  • 189
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    An experimental investigation of surface pressure measurements on an advanced winged entry vehicle at Mach 10 (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Surface pressure measurements have been made at Mach 10 in air on an instrumented 0.006-scale model of an advanced (control configured) winged entry vehicle. The tests were conducted in the Langley Continuous Flow Hypersonic Tunnel. Data were obtained at 83 surface pressure stations, which include locations on the lower and upper surface centerlines, spanwise positions along the lower and upper surfaces of the wing, the lower surface of the body flap, and radial locations on the fuselage. Data were obtained for angles of attack ranging from zero to 40 deg, sideslip angles of -2 deg to +5 deg, Reynolds numbers of 0.5, 1.0, and 2.0 million per foot, and body-flap deflections of zero, 10, and 20 deg. Test conditions and orifice locations were chosen to correspond directly with those for the heat transfer measurements previously reported on the same configuration. Comparison of windward symmetry plane data with predictions based upon an approximate engineering method was found to yield reasonable agreement for angles of attack from 20 to 40 deg. The leeward surface pressure data were observed to be roughly an order of magnitude lower than the corresponding windward data. At low angles of attack, regions of high pressure were noted on the windward wing surface. The result is attributed to vortical action or shock impingement. High pressures were also measured on the deflected body flap, a critical region for this type of vehicle. Reynolds number effects were found to be insignificant.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0308
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  • 190
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    Engineering flowfield method with angle-of-attack applications (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An approximate inviscid flowfield method has been extended to include heat-transfer predictions using a technique to account for variable-entropy edge conditions. The engineering code computes the flowfield over hyperboloids, ellipsoids, paraboloids, and sphere cones at 0 deg angle of attack (AOA). For angle-of-attack applications, an approximation to sphere-cone streamline-spreading effects on the heat transfer along the windward and leeward rays and an empirical circumferential heating technique have been incorporated also in the method. The present engineering calculations yield good comparisons with existing pressure and heating data over sphere cones even at high incidence values with the restriction that the sonic-line location remain on the spherical cap.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0303
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  • 191
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    Transonic analysis of canted winglets (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A computational method developed to provide a transonic analysis for upper/lower surface wing-tip mounted winglets is described. Winglets with arbitrary planform, cant and toe angle, and airfoil section can be modeled. The embedded grid approach provides high flow field resolution and the required geometric flexibility. In particular, coupled Cartesian/cylindrical grid systems are used to model the complex geometry presented by canted upper/lower surface winglets. A new rotated difference scheme is introduced in order to maintain the stability of the small-disturbance formulation in the presence of large spanwise velocities. Wing and winglet viscous effects are modeled using a two-dimensional 'strip' boundary layer analysis. Correlations with wind tunnel and flight test data for three transport configurations are included.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0302
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  • 192
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    Comparison of measured and calculated airloads on an energy efficient transport wing model equipped with oscillating control surfaces (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Wind-tunnel measurements of steady and unsteady pressures for a high-aspect-ratio supercritical wing model are compared with calculations by the linear unsteady aerodynamic lifting-surface theory, known as the Doublet Lattice method, at Mach numbers of 0.650 (subsonic) and 0.78 (transonic). The steady-pressure data comparisons are made for incremental changes in angle of attack and control-surface deflection. The unsteady-pressure data comparisons are made for oscillating control-surface deflections. Some differences between the measured and calculated aerodynamics are attributed to viscous and transonic effects that are not accounted for in the Doublet Lattice analysis. Comparisons of the transonic unsteady-pressure data for the oscillating control surfaces are improved by applying empirical corrections based on the steady-pressure measurements to the unsteady Doublet Lattice calculations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0301
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  • 193
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    Transonic solutions for a multielement airfoil using the full-potential equation (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Transonic flow solutions are obtained over a multielement airfoil (augmentor-wing) using the full-potential equation. Solutions obtained for a subcritical case and a strong shock case show good quantitative agreement with experiment in regions not dominated by viscous effects. In those regions where viscous effects are dominant, the results are still in good qualitative agreement. For the strong shock case, Mach number and angle-of-attack corrections were necessary to match experimental coefficient of lift. Typical results from the transonic augmentor-wing Potential Code on the Cray-1S computer require about 10 sec of CPU time for a three-order-of-magnitude drop in the maximum residual. The speed with which solutions can be generated, and the associated low cost, will make this code a practical tool for the design aerodynamicist.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0300
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  • 194
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    An incompressible Navier-Stokes flow solver in three-dimensional curvilinear coordinate systems using primitive variables (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An implicit, finite-difference computer code has been developed to solve the incompressible Navier-Stokes equations in a three-dimensional, curvilinear coordinate system. The pressure-field solution is based on the pseudo compressibility approach in which the time derivative pressure term is introduced into the mass conservation equation to form a set of hyperbolic equations. The solution procedure employs an implicit, approximate factorization scheme. The Reynolds stresses, that are uncoupled from the implicit scheme, are lagged by one time-step to facilitate implementing various levels of the turbulence model. Test problems for external and internal flows are computed, and the results are compared with existing experimental data. The application of this technique for general three-dimensional problems is then demonstrated.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0253
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  • 195
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    Status of orifice induced pressure error studies (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The construction of high Reynolds number facilities, such as the National Transonic Facility (NTF) and the 0.3-m Transonic Cryogenic Tunnel (TCT), has stimulated interest again in the study of orifice induced static pressure. In a high Reynolds number facility, the orifice will have a much larger effect on the boundary layer than in a conventional wind tunnel. The present investigation was performed in the 0.3-m TCT at Mach numbers in the range from 0.60 to 0.80 and Reynolds numbers in the range from 6,000,000 to 40,000,000 with the objective to compare the porous plug orifices to conventional 0.025 cm orifices in a high Reynolds number environment. It was found that there was an error at high Reynolds numbers which could not be neglected and that the use of a porous metal disk in a conventional orifice could virtually eliminate the orifice induced pressure error.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0245
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  • 196
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    Effect of sidewall suction on flow in two-dimensional wind tunnels (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A closed-form analysis of flow in a two-dimensional subsonic wind tunnel which uses sidewall suction around the model to reduce sidewall boundary-layer effects is presented. The model problem which is treated involves a flat plate airfoil in a tunnel with a suction window shaped to permit an analytic solution. This solution shows that the lift coefficient depends explicitly on the porosity parameter of the suction window and implicitly on the suction pressure differential. For a given sidewall displacement thickness, the lift coefficient increases as the suction-window porosity decreases.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0242
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  • 197
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    Computer graphic visualization of orbiter lower surface boundary-layer transition (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Computer graphic techniques are applied to the processing of Shuttle Orbiter flight data in order to create a visual presentation of the extent and movement of the boundary-layer transition front over the orbiter lower surface during entry. Flight-measured surface temperature-time histories define the onset and completion of the boundary-layer transition process at any measurement location. The locus of points which define the spatial position of the boundary-layer transition front on the orbiter planform is plotted at each discrete time for which flight data are available. Displaying these images sequentially in real-time results in an animated simulation of the in-flight boundary-layer transition process.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0228
    Format: text
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  • 198
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    Vectorized schemes for conical potential flow using the artificial density method (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A method is developed to determine solutions to the full-potential equation for steady supersonic conical flow using the artificial density method. Various update schemes used generally for transonic potential solutions are investigated. The schemes are compared for speed and robustness. All versions of the computer code have been vectorized and are currently running on the CYBER-203 computer. The update schemes are vectorized, where possible, either fully (explicit schemes) or partially (implicit schemes). Since each version of the code differs only by the update scheme and elements other than the update scheme are completely vectorizable, comparisons of computational effort and convergence rate among schemes are a measure of the specific scheme's performance. Results are presented for circular and elliptical cones at angle of attack for subcritical and supercritical crossflows.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0162
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  • 199
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    Simulation of transonic separated airfoil flow by finite difference viscous-inviscid interaction (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A finite difference viscous inviscid interaction program has been developed for simulating the separated transonic flow about lifting airfoils, including the wake. In contrast to most interaction programs, this code combines a finite difference boundary layer algorithm with the inviscid program. The recently developed finite difference boundary layer code efficiently simulates attached and reversed compressible boundary layer and wake flows. New viscous inviscid interaction algorithms were also developed to couple the boundary layer code with the inviscid transonic full potential program. Transonic cases with shock induced and trailing edge separation are computed and compared with experimental and Navier-Stokes results.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85980 , A-9812 , NAS 1.15:85980
    Format: application/pdf
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  • 200
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    Fuselage and nozzle pressure distributions of a 1/12-scale F-15 propulsion model at transonic speeds. Effect of fuselage modifications and nozzle variables (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Static pressure coefficient distributions on the forebody, afterbody, and nozzles of a 1/12 scale F-15 propulsion model was determined in the 16 foot transonic tunnel for Mach numbers from 0.60 to 1.20, angles of attack from -2 deg to 7 deg and ratio of jet total pressure to free stream static pressure from 1 up to about 7, depending on Mach number. The effects of nozzle geometry and horizontal tail deflection on the pressure distributions were investigated. Boundary layer total pressure profiles were determined at two locations ahead of the nozzles on the top nacelle surface. Reynolds number varied from about 1.0 x 10 to the 7th power per meter, depending on Mach number.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2333 , L-15755 , NAS 1.60:2333
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