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  • AERODYNAMICS  (629)
  • 1985-1989  (629)
  • 1950-1954
  • 1987  (629)
  • 101
    Publication Date: 2011-08-19
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 0021-8669); 24; 377-385
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  • 102
    Publication Date: 2011-08-19
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 0021-8669); 24; 392-398
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  • 103
    Publication Date: 2011-08-19
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 25; 252-259
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  • 104
    Publication Date: 2011-08-19
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 25; 266-274
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  • 105
    Publication Date: 2011-08-19
    Description: A special array system has been designed to examine noise source distributions over a helicopter rotor model. The particular measurement environment is for a rotor operating in the open jet of an anechoic wind tunnel. An out-of-flow directional microphone element array is used with a directivity pattern whose major directional lobe projects on the rotor disk. If significant contributions from extraneous tunnel noise sources in the direction of the side lobes are excluded, the dominant output from the array would be that noise emitted from the projected area on the rotor disk. The design incorporates an array element signal blending features which serves to control the spatial resolution of the size of the directional lobes. (Without blending, the resolution and side lobe size are very strong functions of frequency, which severely limits the array's usefulness).
    Keywords: AERODYNAMICS
    Type: Journal of Sound and Vibration (ISSN 0022-460X); 112; 192-197
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  • 106
    Publication Date: 2011-08-19
    Description: A one-parameter family of explicit and implicit total variation diminishing (TVD) schemes is developed which permits incorporation of an expanded group of slope and flux limiters. The numerical technique is intended for use in calculations which include a time-differencing scheme and an optional Lax-Wendroff scheme. Methods of extending the TVD models to nonlinear scalar equations and systems of hyperbolic conservation equations are described. Sample results are presented from calculations of shocked flows around NACA 0012 and NACA 0018 airfoils.
    Keywords: AERODYNAMICS
    Type: Journal of Computational Physics (ISSN 0021-9991); 68; 151-179
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  • 107
    Publication Date: 2011-08-19
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 25; 294-299
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  • 108
    Publication Date: 2011-08-19
    Keywords: AERODYNAMICS
    Type: Journal of Propulsion and Power (ISSN 0748-4658); 3; 63-70
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  • 109
    Publication Date: 2013-08-29
    Description: Numerical simulation of compressible viscous flow fields is performed for a transonic transport configuration. A single structured grid system is constructed using analytical transformations such as conformal mapping, shearing/twisting/rotating/clustering/stretching transformations. The Reynolds-averaged, thin-layer Navier-Stokes equations are solved on a supercomputer, FACOM VP-400, using the LU-ADI factorization method.
    Keywords: AERODYNAMICS
    Type: National Aerospace Lab., Proceedings of the 5th NAL Symposium on Aircraft Computational Aerodynamics; p 85-89
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  • 110
    Publication Date: 2013-08-29
    Description: The slightly supersonic viscous flow about the space-plane under development at the National Aerospace Laboratory (NAL) in Japan was simulated numerically using the LU-ADI algorithm. The wind-tunnel testing for the same plane also was conducted with the computations in parallel. The main purpose of the simulation is to capture the phenomena which have a great deal of influence to the aerodynamic force and efficiency but is difficult to capture by experiments. It includes more accurate representation of vortical flows with high angles of attack of an aircraft. The space-plane shape geometry simulated is the simplified model of the real space-plane, which is a combination of a flat and slender body and a double-delta type wing. The comparison between experimental results and numerical ones will be done in the near future. It could be said that numerical results show the qualitatively reliable phenomena.
    Keywords: AERODYNAMICS
    Type: National Aerospace Lab., Proceedings of the 5th NAL Symposium on Aircraft Computational Aerodynamics; p 13-18
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  • 111
    Publication Date: 2013-08-31
    Description: The concept of circulation control was successfully demonstrated in flight using an A-6 aircraft. Circulation control can provide an aircraft with STOL performance of heavy lift capability. For ship based Naval aircraft the lower takeoff and landing velocities result in reduced deck gear and wind over the deck requirements. Circulation control airfoils can be mechanically less complex and lightweight compared to multi-element high lift airfoils.
    Keywords: AERODYNAMICS
    Type: NASA. Ames Research Center Proceedings of the Circulation-Control Workshop, 1986; p 479-489
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  • 112
    Publication Date: 2013-08-31
    Description: The rotorcraft dynamics analysis was used to predict the aeroelastic responses of a representative X-wing model with a 10 ft diameter rotor. The aeroelastic methodology used and the tests and assumptions involved are reviewed. Results are reported on the findings concerning control power and higher harmonic control in hover, transition flight, vibratory loads at forward speed, and responses in conversion. It is concluded that the analysis can give satisfactory predictions of X-wing behavior.
    Keywords: AERODYNAMICS
    Type: NASA. Ames Research Center Proceedings of the Circulation-Control Workshop, 1986; p 383-398
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  • 113
    Publication Date: 2013-08-31
    Description: The thrust vectoring of supersonic Coanda jets was improved by designing a nozzle to skew the initial jet velocity profile. A new nozzle design procedure, based on the method of characteristics, was developed to design a nozzle which produces a specified exit velocity profile. The thrust vectoring of a simple convergent nozzle, a convergent-divergent nozzle, and a nozzle which produces a skewed velocity profile matched to the curvature of the Coanda surface were expermentially compared over a range of pressure ratios from 1.5 to 3.5. Elimination of the expansion shocks with the C-D nozzle is shown to greatly improve the thrust vectoring; elimination of turning shocks with the skewed profile nozzle further improves the vectoring.
    Keywords: AERODYNAMICS
    Type: NASA. Ames Research Center Proceedings of the Circulation-Control Workshop, 1986; p 289-312
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  • 114
    Publication Date: 2013-08-31
    Description: Wind tunnel measurements of boundary layer and wake velocity profiles and surface static pressure distributions are presented for a swept, circulation control wing. The model is an aspect ratio four semispan wing mounted on the tunnel side wall as a sweep angle of 45 deg. A full span, tangetial, rearward blowing, circulation control slot is located ahead of the trailing edge on the upper surface. Flow surveys were obtained at mid-semispan at freestream Mach numbers of 0.425 and 0.70. Boundary layer profiles measured on the forward portions of the wing are approximately streamwise and two dimensional. The flow in the vicinity of the jet exit and in the near wake is highly three dimensional. The jet flow near the slot on the Coanda surface is directed normal to the slot. Near wake surveys show large outboard flows at the center of the wake. At Mach 0.425 and a 5 deg angle of attack, a range of jet blowing rates was found for which an abrupt transition from incipient separation to attached flow occurs in the boundary layer upstream of the slot. The variation in the lower surface separation location with blowing rate was determined from the boundary layer measurements at Mach 0.425.
    Keywords: AERODYNAMICS
    Type: Proceedings of the Circulation-Control Workshop, 1986; p 239-266
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  • 115
    Publication Date: 2013-08-31
    Description: Classical and zero-total pressure-loss sets of Euler equations were applied to sharp- and round-edge delta wings. The origin of the total pressure was explained in the classical set. For sharp-edged delta wings, all sets of Euler equations produce the same separated flow solutions. For round-edged delta wings and for coarse grids, the solution depends on the level of dissipation, the accuracy of the surface boundary condition, and the type of Euler equations set. For round-edged delta wings and for fine grids, attached flow solutions are obtained. Also presented were three dimensional flow solutions and asymmetric flow solutions including unsteady flow for sharp-edged delta wings. Euler equations should be restricted to sharp-edged wings for real flow solutions. For roung-edged wings, Navier-Stokes equations must be used.
    Keywords: AERODYNAMICS
    Type: Unsteady Hybrid Vortex Technique for Transonic Vortex Flows and Flutter Applications; 29 p
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  • 116
    Publication Date: 2013-08-31
    Description: Two transonic computational schemes which are based on the Integral Equation Formulation of the full potential equation were presented. The first scheme is a Shock Capturing-Shock Fitting (SCSF) scheme which uses the full potential equation throughout with the exception of the shock wave where the Rankine-Hugoniot relations are used to cross and fit the shock. The second scheme is an Integral Equation with Embedded Euler (IEEE) scheme which uses the full potential equation with an embedded region where the Euler equations are used. The two schemes are applied to several transonic airfoil flows and the results were compared with numerous computational results and experimental domains with fine grids. The SCSF-scheme is restricted to flows with weak shock, while the IEEE-scheme can handle strong shocks. Currently, the IEEE scheme is applied to other transonic flows with strong shocks as well as to unsteady pitching oscillations.
    Keywords: AERODYNAMICS
    Type: Unsteady Hybrid Vortex Technique for Transonic Vortex Flows and Flutter Applications; 15 p
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  • 117
    Publication Date: 2013-08-31
    Description: The general problem of calculating the flow fields associated with hypersonic airbreathing aircraft is presented. Unique aspects of hypersonic aircraft aerodynamics are introduced and their demands on computational fluid dynamics are outlined. Example calculations associated with inlet/forebody integration and hypersonic nozzle design are presented to illustrate the nature of the problems considered.
    Keywords: AERODYNAMICS
    Type: NASA. Ames Research Center, Supercomputing in Aerocomputing; p 239-255
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  • 118
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2013-08-31
    Description: A glimpse is provided of the research program in stability, transition, and turbulence based on numerical simulations. This program includes both the so-called abrupt and the restrained transition processes. Attention is confined to the prototype problems of channel flow and the parallel boundary layer in the former category and the Taylor-Couette flow in the latter category. It covers both incompressible flows and supersonic flows. Some representative results are presented.
    Keywords: AERODYNAMICS
    Type: NASA. Ames Research Center, Supercomputing in Aerospace; p 211-220
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  • 119
    Publication Date: 2013-08-31
    Description: The effect of near proximity to the ground are investigated on a low aspect ratio propulsive wing/canard concept at STOL conditions. Data were obtained on a wing/body and wing/body/canard configuration at various heights above the ground, ranging from free air to approximately 1/4 of the mean aerodynamic chord (MAC) above the ground. The data presented and discussed include force and moment coefficients, surface pressure distributions, and downwash angles measured one MAC behind the wing. The test technique, model requirements, and special considerations required for testing these configurations are also discussed. Special model requirements included evenly distributed exit nozzle pressures along four separate nozzles of lengths of one and two feet with only one air supply to the model. Test techniques must recognize and deal with the ground boundary layer as well as the air supply pressure measurement and management.
    Keywords: AERODYNAMICS
    Type: NASA. Ames Research Center Proceedings of the 1985 NASA Ames Research Center's Ground-Effects Workshop; p 415-444
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  • 120
    Publication Date: 2013-08-31
    Description: An experimental facility was developed in the 1.23 (48 inch) wind tunnel of the Applied Research Lab. at the Pennsylvania State Univ. to model the ground vortex. The purpose of the facility was to study the effect of various parameters on the location and characteristics of a ground vortex. An experimental investigation was conducted in the tunnel into the formation, stability and strength of the ground vortex for several flow parameters. The design of the facility, special instrumentation and results are summarized.
    Keywords: AERODYNAMICS
    Type: NASA. Ames Research Center Proceedings of the 1985 NASA Ames Research Center's Ground-Effects Workship; p 207-238
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  • 121
    Publication Date: 2013-08-31
    Description: Mean velocity and turbulence measurements were conducted on the three dimensional fountain flow field generated by the impingement of two axisymmetric jets on a ground plane with application to vertical takeoff and landing (VTOL) aircraft. The basic instantaneous velocity data were obtained using a two component laser Doppler velocimeter in a plane connecting the nozzle centerlines at different heights above the ground emphasizing the jet impingement region and the fountain upwash region formed by the collision of the wall jets. The distribution of mean velocity components and turbulence quantities, including the turbulence intensity and the Reynolds shear stress, were derived from the basic velocity data. Detailed studies of the characteristics of the fountain revealed self-similarity in the mean velocity and turbulence profiles across the fountain. The spread and mean velocity decay characteristics of the fountain were established. Turbulence intensities of the order of 50% were observed in the fountain.
    Keywords: AERODYNAMICS
    Type: NASA. Ames Research Center Proceedings of the 1985 NASA Ames Research Center's Ground-Effects Workship; p 147-159
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  • 122
    Publication Date: 2013-08-31
    Description: Numerical simulations of a planar turbulent wall jet and a planar VTOL upwash fountain were performed. These are three dimensional simulations which resolve large scale unsteady motions in the flows. The wall jet simulation shows good agreement with experimental data and is presented to verify the simulation methodology. Simulation of the upwash fountain predicts elevated shear stress and a half velocity width spreading rate of 33% which agrees well with experiment. Turbulence mechanisms which contribute to the enhanced spreading rate are examined.
    Keywords: AERODYNAMICS
    Type: NASA. Ames Research Center Proceedings of the 1985 NASA Ames Research center's Ground-Effects Workship; p 195-206
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  • 123
    facet.materialart.
    Unknown
    In:  Other Sources
    Publication Date: 2018-12-01
    Description: An overview of the most recent work conducted by NASA and others to study the potential influence of heavy rain on airfoil performance is presented. Previous analytical investigations are discussed, and some promising experimental methods for evaluating rain effects are examined. Special attention is given to the scaling analysis. Results from wind tunnel tests indicated that a conventional NACA 64-210 airfoil and an unflapped NACA 0012 airfoil have different sensitivities to a simulated rain spray. Very little effect was noted on the lift of the NACA 64-210 airfoil, while the NACA 0012 showed a considerable loss in maximum lift capability. With both airfoils in a flapped configuration, significant reductions in maximum lift capability were noted. For the NACA 64-210 airfoil, a reduction in the angle of attack for maximum lift was observed. For both airfoils, the effect of rain on lift occurred near the region of maximum lift; little effect was observed at lower angles of attack.
    Keywords: AERODYNAMICS
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  • 124
    Publication Date: 2019-06-28
    Description: A method using curved vortex elements was developed for helicopter rotor free wake calculations. The Basic Curve Vortex Element (BCVE) is derived from the approximate Biot-Savart integration for a parabolic arc filament. When used in conjunction with a scheme to fit the elements along a vortex filament contour, this method has a significant advantage in overall accuracy and efficiency when compared to the traditional straight-line element approach. A theoretical and numerical analysis shows that free wake flows involving close interactions between filaments should utilize curved vortex elements in order to guarantee a consistent level of accuracy. The curved element method was implemented into a forward flight free wake analysis, featuring an adaptive far wake model that utilizes free wake information to extend the vortex filaments beyond the free wake regions. The curved vortex element free wake, coupled with this far wake model, exhibited rapid convergence, even in regions where the free wake and far wake turns are interlaced. Sample calculations are presented for tip vortex motion at various advance ratios for single and multiple blade rotors. Cross-flow plots reveal that the overall downstream wake flow resembles a trailing vortex pair. A preliminary assessment shows that the rotor downwash field is insensitive to element size, even for relatively large curved elements.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3958 , NAS 1.26:3958 , CDI-84-6
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  • 125
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of a turboprop-nacelle installation on the pressure distributions over a swept, supercritical wing. The tests were conducted at Mach numbers from 0.20 to 0.80, at angles of attack from 0 to 5 degrees, nacelle nozzle pressure ratios from 1.0 to 1.6, and at propeller tip speeds from 700 to 800 ft/sec. The results of this study indicate that the turboprop nacelle interference, with and without power, on a swept wing is greater on the inboard wing panel than on the outboard wing panel. The over-the-wing nacelle installation with the propeller upwash on the inboard panel had flow separation problems at a Mach number of 0.80. No severe flow separation problems appear to exist for either propeller rotation direction for the under-the-wing nacelle installation. The local flow disturbances caused by the under-the-wing nacelle installation were in general less severe than for the over-the-wing nacelle installation.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2729 , L-16043 , NAS 1.60:2729
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  • 126
    Publication Date: 2019-06-28
    Description: A multigrid algorithm has been developed for solving the steady-state Euler equations in two dimensions on unstructured triangular meshes. The method assumes the various coarse and fine grids of the multigrid sequence to be independent of one another, thus decoupling the grid generation procedure from the multigrid algorithm. The transfer of variables between the various meshes employs a tree-search algorithm which rapidly identifies regions of overlap between coarse and fine grid cells. Finer meshes are obtained either by regenerating new globally refined meshes, or by adaptively refining the previous coarser mesh. For both cases, the observed convergence rates are comparable to those obtained with structured multigrid Euler solvers. The adaptively generated meshes are shown to produce solutions of higher accuracy with fewer mesh points.
    Keywords: AERODYNAMICS
    Type: NASA-CR-178346 , ICASE-87-53 , NAS 1.26:178346
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  • 127
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the Langley Transonic Dynamics Tunnel to evaluate differences between an existing utility-class main-rotor blade and an advanced-design main-rotor blade. The two rotor blade designs were compared with regard to rotor performance oscillatory pitch-link loads, and 4-per-rev vertical fixed-system loads. Tests were conducted in hover and over a range of simulated full-scale gross weights and density altitude conditions at advance ratios from 0.15 to 0.40. Results indicate that the advanced blade design offers performance improvements over the baseline blade in both hover and forward flight. Pitch-link oscillatory loads for the baseline rotor were more sensitive to the test conditions than those of the advanced rotor. The 4-per-rev vertical fixed-system load produced by the advanced blade was larger than that produced by the baseline blade at all test conditions.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89129 , L-16260 , NAS 1.15:89129 , AVSCOM-TM-87-B-8 , AD-A184574
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  • 128
    Publication Date: 2019-06-28
    Description: The Navier-Stokes (NS) equations were integrated numerically for investigating the flow characteristics on the forepart of the spherical nose of a space vehicle such as the AOTV or AFE by a modified Accelerated Successive Replacement (ASR) scheme under hypersonic rarefied conditions. Technical feasibility of the mathematical approach was demonstrated by computing the flowfield on a spherical nose under conditions that the AFE encounters at times t = 15 and 20 seconds after its reentry into the atmosphere. Local similar solutions for the merged layer flow along the stagnation line of the sphere were developed. These are correct to the same degree of accuracy as the NS equations. These solutions provided stagnation line boundary conditions for the domain of integration on the spherical noise. Also, a parametric study of the stagnation line solution was made with a view to understand the flow characteristics in tunnels with different ambient fluids. Analytical expressions for surface slip temperature, jump conditions, and concentration level in the presence of the real gas effects at the top of the Knudsen layer were derived and used to calculate the stagnation line flowfield with nonequilibrium dissociation and ionization. A number of graphics were drawn to illustrate the basic physics of the flowfields. The present analysis can be extended to include real gas effects and to bodies of arbitrary shapes. It can further provide boundary conditions for integrating the NS equations in the near wake region.
    Keywords: AERODYNAMICS
    Type: NASA-CR-179153 , NAS 1.26:179153 , RTR-175-01
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  • 129
    Publication Date: 2019-06-28
    Description: A flight research program was undertaken at the NASA Langley Research Center to apply the vapor-screen technique, widely used in wind tunnels, to an aircraft. The purpose was to obtain qualitative and quantitative information about near-field vortex flows above the wings of fighter aircraft and ascertain the effects of Reynolds and Mach numbers over the angle-of-attack range. The hardware for the systems required for flight application of the vapor-screen technique was successfully developed and integrated. Details of each system, its operational performance on the F-106B aircraft, and pertinent aircraft and environmental data collected are presented.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4004 , L-16306 , NAS 1.15:4004
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  • 130
    Publication Date: 2019-06-28
    Description: The aerodynamic characteristics of the 40- by 80-Foot Wind Tunnel at Ames Research Center were measured by using a 1/50th-scale facility. The model was configured to closely simulate the features of the full-scale facility when it became operational in 1986. The items measured include the aerodynamic effects due to changes in the total-pressure-loss characteristics of the intake and exhaust openings of the air-exchange system, total-pressure distributions in the flow field at locations around the wind tunnel circuit, the locations of the maximum total-pressure contours, and the aerodynamic changes caused by the installation of the acoustic barrier in the southwest corner of the wind tunnel. The model tests reveal the changes in the aerodynamic performance of the 1986 version of the 40- by 80-Foot Wind Tunnel compared with the performance of the 1982 configuration.
    Keywords: AERODYNAMICS
    Type: NASA-TM-88336 , A-86329 , NAS 1.15:88336
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  • 131
    Publication Date: 2019-06-28
    Description: Subsonic lateral-direction and longitudinal characteristics of a forward-swept-wing fighter configuration were examined in wind-tunnel tests at Mach numbers of 0.2 and 0.5 for angles of attack from -7 to 47 deg. and over a sidelslip range of +/- 15 deg. The effects of a canard, strakes, vertical tail, and leading- and trailing-edge flaps are examined. The canard and strakes both reduce asymmetric moments and side forces at zero sideslip for angles of attack up to about 30 deg. The canard has a small influence on lateral-directional stability; however, strakes produce a substantial reduction in lateral stability for angles of attack greater than about 20 deg. The vertical tail improves directional stability for angles of attack up to 30 deg. Deflection of the leading-edge flap to 20 deg. at high angles of attack on the strake and canard configurations degrades lateral and directional stability. Deflection of the trailing-edge flap to 20 deg. on the canard configuration generally increases lateral and directional stability at high angles of attack. Leading- and trailing-edge flaps on the wing-body and canard configurations are effective for increased lift only for angles of attack up to about 40 deg. The leading-edge flap remains effective on the strake configuration over the entire angle-of-attack range tested.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2727 , L-16206 , NAS 1.60:2727
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  • 132
    Publication Date: 2019-06-28
    Description: The influence of Mach and Reynolds numbers as well as airfoil and planform geometry on the phenomenon of constant shock jump pressure coefficient for conditions of shock induced trailing edge separation (SITES) was studied. It was demonstrated that the phenomenon does exist for a wide variety of two and three dimensional flow cases and that the influence of free stream Mach number was not significant. The influence of Reynolds number was found to be important but was not strong. Airfoil and planform geometric characteristics were found to be very important where the pressure coefficient jump was shown to vary with the sum of: (1) airfoil curvature at the upper surface crest, and (2) camber surface slope at the trailing edge. It was also determined that the onset of SITES could be defined as a function of airfoil geometric parameters and Mach number normal to the leading edge. This onset prediction was shown to predict the angle of onset to within + or - 1 deg accuracy or better for about 90% of the cases studied.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4090 , NAS 1.26:4090
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  • 133
    Publication Date: 2019-06-28
    Description: Numerical methods were developed that will form the computational part of the turbulence closure scheme. A wave model was developed for the two-dimensional shear layer. This configuration is being used as a test case for the closure schemes. Various numerical schemes were examined to give efficient solutions of the Rayleigh equation for this geometry. These include both spectral and finite difference methods. Secondly, numerical methods are under development to solve the non-separable Rayleigh equation. This solution is required for the closure scheme in more complex geometries. A model problem was used to assist in the algorithm development. Two-dimensional spectral methods and a hybrid spectral/finite difference technique were developed. An analytic solution of the Rayleigh equation for a basic elliptic flow was obtained. This will be used to verify the stability codes developed for arbitrary geometries. Other numerical methods for solving the Rayleigh equation based on the boundary element technique were also examined. These solutions are forming the basis of a model for the shock structure in jets of arbitrary geometry.
    Keywords: AERODYNAMICS
    Type: NASA-CR-180558 , NAS 1.26:180558
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  • 134
    Publication Date: 2019-06-28
    Description: The initial phases of a study of the large-amplitude unsteady aerodynamics of wings in severe maneuver are reported. The research centers on vortex flows, their initiation at wing surfaces, their subsequent convection, and interaction dynamically with wings and control surfaces. The focus is on 2D and quasi-2D aspects of the problem and features the development of an exact nonlinear unsteady airfoil theory as well as an approach to the crossflow problem for slender wing applications including leading-edge separation. The effective use of interactive on-line computing in quantifying and visualizing the nonsteady effects of severe maneuver is demonstrated. Interactive computational work is now possible, in which a maneuver can be initiated and its effects observed and analyzed immediately.
    Keywords: AERODYNAMICS
    Type: NASA-CR-181008 , NAS 1.26:181008
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  • 135
    Publication Date: 2019-06-28
    Description: The body surface-panel method SOUSSA is applied to calculate steady and unsteady lift and pitching moment coefficients on a thin fighter-type wing model with and without a tip-mounted missile. Comparisons are presented with experimental results and with PANAIR and PANAIR-related calculations for Mach numbers from 0.6 to 0.9. In general the SOUSSA program, the experiments, and the PANAIR (and related) programs give lift and pitching-moment results which agree at least fairly well, except for the unsteady clean-wing experimental moment and the unsteady moment on the wing tip body calculated by a PANAIR-predecessor program at a Mach number of 0.8.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2736 , L-16262 , NAS 1.60:2736
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  • 136
    Publication Date: 2019-06-28
    Description: A wind-tunnel test has been conducted in the Langley Low-Turbulence Pressure Tunnel to evaluate the performance of a symmetrical NASA LS(1)-0013 airfoil which is a 13-percent-thick, low-speed airfoil. The airfoil contour was obtained from the thickness distribution of a 13-percent-thick, high-performance airfoil developed for general aviation airplanes. The tests were conducted at Mach numbers from 0.10 tp 0.37 over a Reynolds number range from about 0.6 to 12.0 X 10 to the 6th power. The angle of attack varied from about -8 to 20 degrees. The results indicate that the aerodynamic characteristics of the present airfoil are similar to, but slightly better than, those of the NACA 0012 airfoil.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4003 , L-16279 , NAS 1.15:4003
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  • 137
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2019-06-28
    Description: Topics addressed include: numerical aerodynamic simulation; computational mechanics; supercomputers; aerospace propulsion systems; computational modeling in ballistics; turbulence modeling; computational chemistry; computational fluid dynamics; and computational astrophysics.
    Keywords: AERODYNAMICS
    Type: NASA-CP-2454 , A-87082 , NAS 1.55:2454
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  • 138
    Publication Date: 2019-06-28
    Description: Two separate tests have been made on the same blended wing-body hydrogen-fueled transport model at a Mach number of about 6 and a range of Reynolds number (based on theoretical body length) of 1.577 to 55.36 X 10 to the 6th power. The results of these tests, made in a conventional hypersonic blowdown tunnel and a hypersonic shock tunnel, are presented through a range of angle of attack from -1 to 8 deg, with an extended study at a constant angle of attack of 3 deg. The model boundary layer flow appeared to be predominately turbulent except for the low Reynolds number shock tunnel tests. Model wall temperatures varied considerably; the blowdown tunnel varied from about 255 F to 340 F, whereas the shock tunnel had a constant 70 F model wall temperature. The experimental normal-force coefficients were essentially independent of Reynolds number. A current theoretical computer program was used to study the effect of Reynolds number. Theoretical predictions of normal-force coefficients were good, particularly at anticipated cruise angles of attack, that is 2 to 5 deg. Axial-force coefficients were generally underestimated for the turbulent skin friction conditions, and pitching-moment coefficients could not be predicted reliably.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2728 , L-16286 , NAS 1.60:2728
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  • 139
    Publication Date: 2019-06-28
    Description: Wind tunnel experiments were conducted on Wortmann FX67-K170, NACA 0012, and NACA 64-210 airfoils at rain rates of 1000 mm/hr and Reynolds numbers of 310,000 to compare the aerodynamic performance degradation of the airfoils and to attempt to identify the various mechanisms which affect performance in heavy rain conditions. Lift and drag were measured in dry and wet conditions, a variety of flow visualization techniques were employed, and a computational code which predicted airfoil boundary layer behavior was used. At low angles of attack, the lift degradation in wet conditions varied significantly between the airfoils. The Wortmann section had the greatest overall lift degradation and the NACA 64-210 airfoil had the smallest. At high angles of attack, the NACA 64-210 and 0012 airfoils had improved aerodynamic performance in rain conditions due to an apparent reduction of the boundry layer separation. Performance degradation in heavy rain for all three airfoils at low angles of attack could be emulated by forced boundary layer transition near the leading edge. The secondary effect occurs at time scales consistent with top surface water runback times. The runback layer is thought to effectively alter the airfoil geometry. The severity of the performance degradation for the airfoils varied. The relative differences appeared to be related to the susceptibility of each airfoil to premature boundary layer transition.
    Keywords: AERODYNAMICS
    Type: NASA-CR-181119 , NAS 1.26:181119 , ASL-87-1
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  • 140
    Publication Date: 2019-06-28
    Description: A numerical simulation of bubble-type vortex breakdown using a unique discrete form of the full 3-D, unsteady incompressible Navier-Stokes equations was performed. The Navier-Stokes equations were written in a vorticity-velocity form and the physical problem was not restricted to axisymmetric flow. The problem was parametized on a Rossby- Reynolds-number basis. Utilization of this parameter duo was shown to dictate the form of the free-field boundary condition specification and allowed control of axial breakdown location within the computational domain. The structure of the breakdown bubble was studied through time evolution plots of planar projected velocity vectors as well as through plots of particle traces and vortex lines. These results compared favorably with previous experimental studies. In addition, profiles of all three velocity components are presented at various axial stations and a Fourier analysis was performed to identify the dominant circumferential modes. The dynamics of the breakdown process were studied through plots of axial variation of rate of change of integrated total energy and rate of change of integrated enstrophy, as well as through contour plots of velocity, vorticity and pressure.
    Keywords: AERODYNAMICS
    Type: NASA-CR-181068 , NAS 1.26:181068
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  • 141
    Publication Date: 2019-06-28
    Description: A parametric investigation of the static internal performance of multifunction two-dimensional convergent-divergent nozzles has been made in the static test facility of the Langley 16-Foot Transonic Tunnel. All nozzles had a constant throat area and aspect ratio. The effects of upper and lower flap angles, divergent flap length, throat approach angle, sidewall containment, and throat geometry were determined. All nozzles were tested at a thrust vector angle that varied from 5.60 tp 23.00 deg. The nozzle pressure ratio was varied up to 10 for all configurations.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2721 , L-16240 , NAS 1.60:2721
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  • 142
    Publication Date: 2019-06-28
    Description: A test to determine the performance differences between the 27-percent-scale models of two rotors for the U.S. Army AH-64 helicopter was conducted in the Langley 14- by 22-Foot Subsonic Tunnel. One rotor, referred to as the baseline rotor, simulated the geometry and dynamic characteristics of the production baseline rotor, and the other rotor, referred to as the advanced rotor, was designed to have improved hover performance. During the performance test, the dynamic pitch-link forces and blade bending and torsion moments were also measured. Dynamic data from the forward flight investigation are reduced and presented. The advanced blade set was designed to have dynamic characteristics similar to those of the baseline rotor so that test conditions would not be limited by potential rotor instability and blade resonances, and so that the measured performance increments could be considered to be due purely to aerodynamic causes. Data show consistent trends with advance ratio for both blade sets with generally higher oscillatory loads occurring for the advanced blade set when compared with the baseline blade set.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89053 , L-16245 , NAS 1.15:89053 , AVSCOM-TM-87-B-7 , AD-A182870
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  • 143
    Publication Date: 2019-06-28
    Description: Direct numerical simulation using the full three dimensional, time dependent Navier-Stokes equations is used to investigate V/STOL jet induced interactions. The objective of this numerical simulation is to compute accurately the details of the flow field and to achieve a better understanding of the physics of the flow, including the role of initial turbulence in the jet, the influence of forward motion on hover aerodynamics, the collision zone and fountain characteristics. Preliminary results are presented.
    Keywords: AERODYNAMICS
    Type: NASA. Ames Research Center Proceedings of the 1985 NASA Ames Research Center's Ground-Effects Workship; p 161-193
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  • 144
    Publication Date: 2019-06-28
    Description: An experimental and theoretical investigation of the longitudinal and lateral-directional stability and control of an axisymmetric cruciform-finned missile has been conducted at Mach 6. The angle-of-attack range extended from 20 to 65 deg to encompass maximum lift. Longitudinal stability, performance, and trim could be accurately predicted with the fins at a fin roll angle of 0 deg but not when the fins were at a fin roll angle of 45 deg. At this roll angle, windward fin choking occurred at angles of attack above 50 deg and reduced the effectiveness of the fins and caused pitch-up.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2733 , L-16287 , NAS 1.60:2733
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  • 145
    Publication Date: 2019-06-28
    Description: A method is presented for aerodynamic design for a specified pressure distribution, using analysis codes only. The method requires a very conservative number of analysis runs, and therefore is appropriate when the analysis code is a large code in terms of storage and/or running time. Three model problems illustrate some capabilities and limitations of the method.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2706 , L-16226 , NAS 1.60:2706
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  • 146
    Publication Date: 2019-06-28
    Description: Increased emphasis on sustained supersonic or hypersonic cruise has revived interest in the supersonic throughflow fan as a possible component in advanced propulsion systems. Use of a fan that can operate with a supersonic inlet axial Mach number is attractive from the standpoint of reducing the inlet losses incurred in diffusing the flow from a supersonic flight Mach number to a subsonic one at the fan face. The design of the experiment using advanced computational codes to calculate the components required is described. The rotor was designed using existing turbomachinery design and analysis codes modified to handle fully supersonic axial flow through the rotor. A two-dimensional axisymmetric throughflow design code plus a blade element code were used to generate fan rotor velocity diagrams and blade shapes. A quasi-three-dimensional, thin shear layer Navier-Stokes code was used to assess the performance of the fan rotor blade shapes. The final design was stacked and checked for three-dimensional effects using a three-dimensional Euler code interactively coupled with a two-dimensional boundary layer code. The nozzle design in the expansion region was analyzed with a three-dimensional parabolized viscous code which corroborated the results from the Euler code. A translating supersonic diffuser was designed using these same codes.
    Keywords: AERODYNAMICS
    Type: NASA-TM-88915 , E-3339 , NAS 1.15:88915
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  • 147
    Publication Date: 2019-06-28
    Description: An investigation was conducted to define pressure distributions for rectangular cavities over a range of free-stream Mach numbers and cavity dimensions. These pressure distributions together with schlieren photographs are used to define the critical values of cavity length-to-depth ratio that separate open type cavity flows from closed type cavity flows. For closed type cavity flow, the shear layer expands over the cavity leading edge and impinges on the cavity floor, whereas for open type cavity flow, the shear layer bridges the cavity. The tests were conducted by using a flat-plate model permitting the cavity length to be remotely varied from 0.5 to 12 in. Cavity depths and widths were varied from 0.5 to 2.5 in. The flat-plate boundary layer approaching the cavity was turbulent and had a thickness of approximately 0.2 in. at the cavity front face for the range of test Mach numbers from 1.5 to 2.86. Presented are a discussion of the results and a complete tabulation of the experimental data.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2683 , L-16215 , NAS 1.60:2683
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  • 148
    Publication Date: 2019-06-28
    Description: A comprehensive static aerodynamic simulation model of the 8 by 8 by 20 ft MILVAN cargo container is determined by combining the wind tunnel data from a 1972 NASA Ames Research Center study taken over the restricted domain (0 is less than or equal to phi is less than or equal to 90 degrees; 0 is less than or equal to alpha is less than or equal to 45 degrees) with extrapolation relations derived from the geometric symmetry of rectangular boxes. It is found that the aerodynamics of any attitude can be defined from the aerodynamics at an equivalent attitude in the restricted domain (0 is less than phi is less than 45 degrees; 0 is less than alpha is less than 90 degrees). However, a similar comprehensive equivalence with the domain spanned by the data is not available; in particular, about two-thirds of the domain with the absolute value of alpha is greater than 45 degrees is unrelated to the data. Nevertheless, as estimate can be defined for this region consistent with the measured or theoretical values along its boundaries and the theoretical equivalence of points within the region. These descrepancies are assumed to be due to measurement errors. Data from independent wind tunnel studies are reviewed; these are less comprehensive than the NASA Ames Research Center but show good to fair agreement with both the theory and the estimate given here.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89433 , A-87126 , NAS 1.15:89433
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  • 149
    Publication Date: 2019-06-28
    Description: The effects of empennage arrangement and afterbody boattail design of nonaxisymmetric nozzles on the aeropropulsive characteristics of a twin-engine fighter-type model have been determined in an investigation conducted in the Langley 16-Foot Transonic Tunnel. Three nonaxisymmetric and one twin axisymmetric convergent-divergent nozzle configurations were tested with three different tail arrangements: a two-tail V-shaped arrangement; a staggered, conventional three-tail arrangement; and a four-tail arrangement similar to that on the F-18. Two of the nonaxisymmetric nozzles were also vectorable. Tests were conducted at Mach numbers from 0.60 to 1.20 over an angle-of-attack range from -3 deg to 9 deg. Nozzle pressure ratio was varied from 1 (jet off) to approximately 12, depending on Mach number. Results indicate that at design nozzle pressure ratio, the medium aspect ratio nozzle (with equal boattail angles on the nozzle sidewalls and upper and lower flaps) had the lowest zero angle of attack drag of the nonaxisymmetric nozzles for all tail configurations at subsonic Mach numbers. The drag levels of the twin axisymmetric nozzles were competitive with those of the medium-aspect-ratio nozzle at subsonic Mach number.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2704 , L-16227 , NAS 1.60:2704
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  • 150
    Publication Date: 2019-06-28
    Description: Computational methods for unsteady transonic flows are surveyed with emphasis on prediction. Computational difficulty is discussed with respect to type of unsteady flow; attached, mixed (attached/separated) and separated. Significant early computations of shock motions, aileron buzz and periodic oscillations are discussed. The maturation of computational methods towards the capability of treating complete vehicles with reasonable computational resources is noted and a survey of recent comparisons with experimental results is compiled. The importance of mixed attached and separated flow modeling for aeroelastic analysis is discussed, and recent calculations of periodic aerodynamic oscillations for an 18 percent thick circular arc airfoil are given.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89106 , NAS 1.15:89106 , AIAA PAPER 87-0107
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  • 151
    Publication Date: 2019-06-28
    Description: An investigation has been conducted to evaluate the effects of several geometric parameters on the internal performance of rectangular thrust-reverser ports for nonaxisymmetric nozzles. Internal geometry was varied with a test apparatus which simulated a forward-flight nozzle with a single, fully deployed reverser port. The test apparatus was designed to simulate thrust reversal (conceptually) either in the convergent section of the nozzle or in the constant-area duct just upstream of the nozzle. The main geometric parameters investigated were port angle, port corner radius, port location, and internal flow blocker angle. For all reverser port geometries, the port opening had an aspect ratio (throat width to throat height) of 6.1 and had a constant passage area from the geometric port throat to the exit. Reverser-port internal performance and thrust-vector angles computed from force-balance measurements are presented.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89061 , L-16211 , NAS 1.15:89061
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  • 152
    Publication Date: 2019-06-28
    Description: The NASA F-106 collected data on the rates of change of electromagnetic parameters on the aircraft surface during over 700 direct lightning strikes while penetrating thunderstorms at altitudes from 15,000 t0 40,000 ft (4,570 to 12,190 m). These in situ measurements provided the basis for the first statistical quantification of the lightning electromagnetic threat to aircraft appropriate for determining indirect lightning effects on aircraft. These data are used to update previous lightning criteria and standards developed over the years from ground-based measurements. The proposed standards will be the first which reflect actual aircraft responses measured at flight altitudes. Nonparametric maximum likelihood estimates of the distribution of the peak electromagnetic rates of change for consideration in the new standards are obtained based on peak recorder data for multiple-strike flights. The linear and nonlinear modeling techniques developed provide means to interpret and understand the direct-strike electromagnetic data acquired on the F-106. The reasonable results obtained with the models, compared with measured responses, provide increased confidence that the models may be credibly applied to other aircraft.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2737 , L-16281 , NAS 1.60:2737
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  • 153
    Publication Date: 2019-06-28
    Description: The flow in a two-dimensional curved channel driven by an azimuthal pressure gradient can become linearly unstable due to axisymmetric perturbations and/or nonaxisymmetric perturbations depending on the curvature of the channel and the Reynolds number. For a particular small value of curvature, the critical neighborhood of this curvature value and critical Reynolds number, nonlinear interactions occur between these perturbations. The Stuart-Watson approach is used to derive two coupled Landau equations for the amplitudes of these perturbations. The stability of the various possible states of these perturbations is shown through bifurcation diagrams. Emphasis is given to those cases which have relevance to external flows.
    Keywords: AERODYNAMICS
    Type: NASA-CR-178286 , ICASE-87-26 , NAS 1.26:178286
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  • 154
    Publication Date: 2019-06-28
    Description: The current scope, recent progress, and plans for research and development of computational methods for unsteady aerodynamics at the NASA Langley Research Center are reviewed. Both integral-equations and finite-difference method for inviscid and viscous flows are discussed. Although the great bulk of the effort has focused on finite-difference solution of the transonic small-perturbation equation, the integral-equation program is given primary emphasis here because it is less well known.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89133 , NAS 1.15:89133
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  • 155
    Publication Date: 2019-06-28
    Description: An investigation to determine the effects of vortex flaps on the flight dynamic characteristics of the F-106B in the area of low-speed, high-angle-of-attack flight was undertaken on a 0.15-scale model of the airplane in the Langley 30- by 60-Foot Tunnel. Static force tests, dynamic forced-oscillation tests, as well as free-flight tests were conducted to obtain a data base on the flight characteristics of the F-106B airplane with vortex flaps. Vortex flap configurations tested included a full-span gothic flap, a full-span constant-chord flap, and a part-span gothic flap.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2700 , L-16202 , NAS 1.60:2700
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  • 156
    facet.materialart.
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    In:  CASI
    Publication Date: 2019-06-28
    Description: Singularities which arise in the solution to elliptic systems are often of great technological importance. This is certainly the case in models of fracture of structures. A survey of the ways singularities are modeled is presented with special emphasis on the effects due to nonlinearities.
    Keywords: AERODYNAMICS
    Type: NASA-CR-178278 , ICASE-87-24 , NAS 1.26:178278
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  • 157
    Publication Date: 2019-06-28
    Description: The analysis and interpretation of wind tunnel pressure data from the Space Shuttle wind tunnel test IA300 are presented. The primary objective of the test was to determine the effects of the Space Shuttle Main Engine (SSME) and the Solid Rocket Booster (SRB) plumes on the integrated vehicle forebody pressure distributions, the elevon hinge moments, and wing loads. The results of this test will be combined with flight test results to form a new data base to be employed in the IVBC-3 airloads analysis. A secondary objective was to obtain solid plume data for correlation with the results of gaseous plume tests. Data from the power level portion was used in conjunction with flight base pressures to evaluate nominal power levels to be used during the investigation of changes in model attitude, eleveon deflection, and nozzle gimbal angle. The plume induced aerodynamic loads were developed for the Space Shuttle bases and forebody areas. A computer code was developed to integrate the pressure data. Using simplified geometrical models of the Space Shuttle elements and components, the pressure data were integrated to develop plume induced force and moments coefficients that can be combined with a power-off data base to develop a power-on data base.
    Keywords: AERODYNAMICS
    Type: NASA-CR-179077 , NAS 1.26:179077 , LMSC-HEC-TR-D951415-1
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  • 158
    Publication Date: 2019-06-28
    Description: The Navier-Stokes Computer is a multi-purpose parallel-processing supercomputer which is currently under development at Princeton University. It consists of multiple local memory parallel processors, called Nodes, which are interconnected in a hypercube network. Details of the procedures involved in implementing an algorithm on the Navier-Stokes computer are presented. The particular finite difference algorithm considered in this analysis was developed for simulation of laminar-turbulent transition in wall bounded shear flows. Projected timing results for implementing this algorithm indicate that operation rates in excess of 42 GFLOPS are feasible on a 128 Node machine.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89119 , NAS 1.15:89119
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  • 159
    Publication Date: 2019-06-28
    Description: Ballistic range tests have been conducted to determine the aerodynamically stable trim attitudes of two proposed general-purpose heat-source module configurations. Tests were conducted at speeds of 4.6 km/sec for the concept module, and at both 4.6 km/sec and 1.4 km/sec for the MOD II. Test results indicated that both configurations were stable when launched in the face-on attitude. When launched in the side-on attitude, the MOD II configuration was found to be stable, while tests of the concept module did not give definitive results.
    Keywords: AERODYNAMICS
    Type: NASA-TM-88345 , A-86368 , NAS 1.15:88345
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  • 160
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2019-06-28
    Description: Nonequilibrium phenomena due to real gas effects are very important features of low density hypersonic flows. The shock shape and emitted nonequilibrium radiation are identified as the bulk flow behavior parameters which are very sensitive to the nonequilibrium phenomena. These parameters can be measured in shock tubes, shock tunnels, and ballistic ranges and used to test the accuracy of computational fluid dynamic (CFD) codes. Since the CDF codes, by necessity, are based on multi-temperature models, it is also desirable to measure various temperatures, most importantly, the vibrational temperature. The CFD codes would require high temperature rate constants, which are not available at present. Experiments conducted at the NASA Electric Arc-driven Shock Tube (EAST) facility reveal that radiation from steel contaminants overwhelm the radiation from the test gas. For the measurement of radiation and the chemical parameters, further investigation and then appropriate modifications of the EAST facility are required.
    Keywords: AERODYNAMICS
    Type: NASA-CR-180441 , NAS 1.26:180441
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  • 161
    Publication Date: 2019-06-28
    Description: A new Navier-Stokes solver is developed by combining the efficiency of the LU-SSOR scheme and the accuracy of the flux-limited dissipation scheme. Application to laminar and turbulent flows and hypersonic flows proves the reliability of the new algorithm.
    Keywords: AERODYNAMICS
    Type: NASA-CR-179608 , E-3499 , NAS 1.26:179608
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  • 162
    Publication Date: 2019-06-28
    Description: The thin-layer Navier-Stokes equations are coupled with a zonal scheme (or domain-decomposition method) to develop the Transonic Navier-Stokes (TNS) wing-alone code. The TNS has a total of 4 zones and is extended to a total of 16 zones for the wing-fuselage version of the code. Results are compared on the Cray X-MP-48 and compared with experimental data.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89421 , A-87066 , NAS 1.15:89421
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  • 163
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the Langley Spin Tunnel of the spin and spin-recovery characteristics of a 1/15-scale model of an Australian trainer airplane. The invesigation included erect and inverted spins; configuration variables such as a long tail, fuselage strakes, 20 deg. elevator cutouts, and rudder modifications; and determination of the parachute size for emergency spin recovery. Also included in the investigation were wing leading-edge modifications to evaluate Reynolds number effects. Results indicate that the basic configuration will spin erect at an angle of attack of about 63 deg. at about 2 to 2.3 seconds per turn. Recovery from this spin was unsatisfactory by rudder reversal or by rudder reversal and ailerons deflected to full with the spin. The elevators had a pronounced effect on the recovery characteristics. The elevators-down position was very adverse to recoveries, whereas the elevators-up position provided favorable recovery effects. Moving the vertical tail aft (producing a long tail configuration) improved the spin characteristics, but the recoveries were still considered marginal. An extension to the basic rudder chord and length made a significant improvement in the spin and recovery characteristics. Satisfactory recoveries were obtained by deflecting the rudder to full against the spin and the elevators and ailerons to neutral.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89049 , L-16191 , NAS 1.15:89049
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  • 164
    Publication Date: 2019-06-28
    Description: Unsteady aerodynamic data were measured on an aspect ratio 10.3 elastic supercritical wing while undergoing high dynamic response above a Mach number of 0.90. These tests were conducted in the NASA Langley Transonic Dynamics Tunnel. A previous test of this wing predicted an unusual instability boundary based on subcritical response data. During the present test no instability was found, but an angle of attack dependent narrow Mach number region of high dynamic wing response was observed over a wide range of dynamic pressures. The effect on dynamic wing response of wing angle of attack, static outbound control surface deflection and a lower surface spanwise fence located near the 60 percent local chordline was investigated. The driving mechanism of the dynamic wing response appears to be related to chordwise shock movement in conjunction with flow separation and reattachment on both the upper and lower surfaces.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89121 , NAS 1.15:89121 , AIAA PAPER 87-0735-CP
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  • 165
    Publication Date: 2019-06-28
    Description: A composite grid was generated in an attempt to improve grid quality for a typical turbine blade with large camber in terms of mesh control, smoothness, and orthogonality. This composite grid consists of the C grid (or O grid) in the immediate vicinity of the blade and the H grid in the upstream region and in the middle of the blade passage between the C grids. It provides a good boundary layer resolution around the leading edge region for viscous calculation, has orthogonality at the blade surface and slope continuity at the C-H (or O-H) interface, and has flexibility in controlling the mesh distribution in the upstream region without using excessive grid points. This composite grid eliminates the undesirable qualities of a single grid when generated for a typical turbine geometry. A finite-volume lower-upper (LU) implicit scheme can be used in solving for the turbine flows on the composite grid. This grid has a special grid node that is connected to more than four neighboring nodes in two dimensions and to more than six nodes in three dimensions. But the finite-volume approach poses no problem at the special point because each interior cell has only four neighboring cells in two dimensions and only six cells in three dimensions. The finite-volume LU implicit scheme was demonstrated to be robust and efficient for both external and internal flows in a broad flow regime.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89828 , E-3477 , NAS 1.15:89828 , USAAVSCOM-TR-87-C-5 , AD-A180145
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  • 166
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2019-06-28
    Description: The method of complex characteristics and hodograph transformation for the design of shockless airfoils was introduced by Bauer, Garabedian, and Korn and has been extended by the author to design subcritical and supercritical cascades with high solidities and large inlet angles. This new capability was achieved by introducing a new conformal mapping of the hodograph domain onto an ellipse and expanding the solution in terms of Chebyshev polynomials. A new computer code, the NASA Lewis inverse design code, was developed based on this idea. This new design code is an efficient method for the design of airfoils in cascade. In particular, the design of subcritical cascades of airfoils is a very fast, robust, and versatile process. The inverse design code can be made to interact with a turbulent boundary layer calculation to obtain airfoils with no separated flows at the design condition. This report is intended to serve as a users manual for this design code. Material previously reported by the author is included here for completeness and quick access to the user. The manual contains a description of the method followed by a discussion of the design procedure and examples. The input parameters necessary to run the code are then described and their default values given. Output listings corresponding to six different blade shapes designed with the code are given, as well as the necessary input data to reproduce the computer runs. The examples have been chosen to show that a wide range of applications can be covered with the code, ranging from supercritical propeller sections to wind tunnel turning vanes that can operate with a large inlet flow angle range.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2676 , E-3221 , NAS 1.60:2676
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  • 167
    Publication Date: 2019-06-28
    Description: A transonic unsteady aerodynamic and aeroelasticity code has been developed for application to realistic aircraft configurations. The new code is called CAP-TSD which is an acronym for Computational Aeroelasticity Program - Transonic Small Disturbance. The CAP-TSD code uses a time-accurate approximate factorization (AF) algorithm for solution of the unsteady transonic small-disturbance equation. The AF algorithm is very efficient for solution of steady and unsteady transonic flow problems. It can provide accurate solutions in only several hundred time steps yielding a significant computational cost savings when compared to alternative methods. The new code can treat complete aircraft geometries with multiple lifting surfaces and bodies including canard, wing, tail, control surfaces, launchers, pylons, fuselage, stores, and nacelles. Applications are presented for a series of five configurations of increasing complexity to demonstrate the wide range of geometrical applicability of CAP-TSD. These results are in good agreement with available experimental steady and unsteady pressure data. Calculations for the General Dynamics one-ninth scale F-16C aircraft model are presented to demonstrate application to a realistic configuration. Unsteady results for the entire F-16C aircraft undergoing a rigid pitching motion illustrated the capability required to perform transonic unsteady aerodynamic and aeroelastic analyses for such configurations.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89120 , NAS 1.15:89120 , AIAA PAPER 87-0850
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  • 168
    Publication Date: 2019-06-28
    Description: No serious studies of the relationship of heavy rain to aircraft safety were made until 1981 when it was suggested that the torrential rain which often occurs at the time of severe wind shear might substantially increase the danger to aircraft operating at slow speeds and high lift in the vicinity of airports. While these data were not published until early 1983, appropriate measures were taken by NASA to study the effect of heavy rain on the lift of wings typical of commercial aircraft. One of the aspects of these tests that seemed confirmed by the data was the existence of a velocity effect on the lift data. The data seemed to indicate that when all the normal non-dimensional aerodynamic parameters were used to sort out the data, the effect of velocity was not accounted for, as it usually is, by the effect of dynamic pressure. Indeed, the measured lift coefficients at high lift indicated a dropoff in lift coefficient for the same free-stream water content as velocity was increased. indicated a drop-off in lift coefficient for the same free-stream water content as velocity was increased.
    Keywords: AERODYNAMICS
    Type: NASA-CR-178248 , NAS 1.26:178248 , ARAP-597
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  • 169
    Publication Date: 2019-06-28
    Description: An investigation of the aerodynamic performance of leading-edge flaps on three clipped delta and three clipped double-delta wing planforms with aspect ratios of 1.75, 2.11, and 2.50 was conducted in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.90, and 2.16. A primary set of fullspan leading-edge flaps with similar root and tip chords were investigated on each wing, and several alternate flap planforms were investigated on the aspect-ratio-1.75 wings. All leading-edge flap geometries were effective in reducing the drag at lifting conditions over the range of wing aspect ratios and Mach numbers tested. Application of a primary flap resulted in better flap performance with the double-delta planform than with the delta planform. The primary flap geometry generally yielded better performance than the alternate flap geometries tested. Trim drag due to flap-induced pitching moments was found to reduce the leading-edge flap performance more for the delta planform than for the double-delta planform. Flow-visualization techniques showed that leading-edge flap deflection reduces crossflow shock-induced separation effects. Finally, it was found that modified linear theory consistently predicts only the effects of leading-edge flap deflection as related to pitching moment and lift trends.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2656 , L-16143 , NAS 1.60:2656
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  • 170
    Publication Date: 2019-06-28
    Description: Measurements have been made in the wake of a semi-span NACA 0015 airfoil with emphasis on the region of the wing tip vortex. The spanwise and streamwise velocity components were measured using a two-component laser Doppler velocimeter. The purpose of the study was to initiate the operation of a laser velocimeter system and to perform preliminary wake measurements in preparation for a more extensive study of the structure and near field development of a tip vortex.
    Keywords: AERODYNAMICS
    Type: NASA-TM-88343 , A-86207 , NAS 1.15:88343 , AVSCOM-TM-86-A-2
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  • 171
    Publication Date: 2019-06-28
    Description: The influence of cold and heated secondary flow on the instability of a two-stream, coplanar jet having a 0.7 Mach number heated primary jet for a nominal fan to primary velocity ratio of 0.68 was investigated by means of inviscid linearized stability theory. The instability properties of spatially growing axisymmetric and first order azimuthal disturbances were studied. The instability characteristics of the two-stream jet with a velocity ratio of 0.68 are very different from those of a single stream jet, and a two-stream, coplanar jet having a 0.9 Mach number heated primary jet and a cold secondary jet for a fan to primary velocity ratio of 0.30. For X/D = 1 and in comparison to the case where the velocity ratio was 0.3, the presence of the fan stream with a velocity ratio of 0.68 enhanced the instability of the jet and increased the unstable frequency range. However, the axisymmetric mode (m = 0) and the first order azimuthal mode (m = 1) have similar spatial growth rates where the velocity ratio is 0.68 while for a velocity ratio of 0.3 the growth rate of the first order azimuthal mode (m = 1) is greater. Comparing the cold and hot secondary flow results showed that for a velocity ratio of 0.68 the growth rate is greater for cold.
    Keywords: AERODYNAMICS
    Type: NASA-TM-88922 , E-3255 , NAS 1.15:88922 , AIAA PAPER 87-0056
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  • 172
    Publication Date: 2019-06-28
    Description: The effect of initial turbulence level on the development of a jet and on the susceptability of the jet to discrete tone excitation was experimentally investigated. Turbulence intensity was varied, over the range 0.15 to 5 percent, by using screens and grids placed upstream of an 8.8 cm diameter nozzle. Top-hat mean velocity profiles with approximately identical initial boundary layer states were ensured in all cases; the turbulence spectra were broadband. It was found, contrary to earlier reports, that the natural jet decay remained essentially unchanged for varying initial turbulence. For a fixed amplitude of the tonal excitation, increasing the initial turbulence damped out the growth of the instability wave; as a result, the excitability, assessed from the mean velocity decay on the axis, was found to diminish. However, the degree of damping in the amplification of the instability wave was only slight compared to the large increase in the initial turbulence. The jet with 5 percent turbulence could be measurably altered by excitation with a velocity perturbation amplitude as little as 0.25 percent of the jet velocity. The amplitude effect data indicate an upper bound of the extent to which a jet could be excited, and thus its plume shortened, by the plane wave, single frequency excitation. An additional data set with no grid or trip, yielding a nominally laminar boundary layer, re-emphasizes the profound effect of initial boundary layer state on jet evolution as well as on its excitability. This jet decayed the fastest naturally, and consequently, it was the least excitable in spite of its turbulence being the least.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100178 , E-3702 , NAS 1.15:100178 , AIAA PAPER 87-2725
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  • 173
    Publication Date: 2019-06-28
    Description: In a variety of aeronautical applications, the flow around conical bodies at incidence is of interest. Such applications include, but are not limited to, highly maneuverable aircraft with delta wings, the aerospace plane and nose portions of spike inlets. The theoretical model used has three parts. First, the single line vortex model is used within the framework of slender body theory, to compute the outer inviscid field for specified separation lines. Next, the three dimensional boundary layer is represented by a momentum equation for the cross flow, analogous to that for a plane boundary layer; a von Karman Pohlhausen approximation is applied to solve this equation. The cross flow separation for both laminar and turbulent layers is determined by matching the pressure at the upper and lower separation points. This iterative procedure yields a unique solution for the separation lines and consequently for the position of the vortices and the vortex lift on the body. Lastly, control of separation is achieved by blowing tangentially from a slot located along a cone generator. It is found that for very small blowing coefficients, the separation can be postponed or suppressedy completely.
    Keywords: AERODYNAMICS
    Type: NASA-CR-181206 , NAS 1.26:181206 , JIAA-TR-78
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  • 174
    Publication Date: 2019-06-28
    Description: A compressible, unsteady, full Navier-Stokes, finite difference code was developed for modeling transonic flow through two-dimensional, oscillating cascades. The procedure introduces a deforming grid technique to capture the motion of the airfoils. Results using a deforming grid are presented for both isolated and cascaded airfoils. The load histories and unsteady pressure distributions are predicted for the NASA 64A010 isolated airfoil and compared with existing experimental data. Results show that the deforming grid technique can be used to successfully predict the unsteady flow properties around an oscillating airfoil. The deforming grid technique was extended for modeling unsteady flow in a cascade. The use of a deforming grid simplifies the specification of boundary conditions. Unsteady flow solutions similar to the isolated airfoil predictions are found for a NACA 0012 cascade with zero interblade phase angle and zero stagger. Experimental data for these cases are not available for code validation, but computational results are presented to show sample predictions from the code. Applications of the code to typical turbomachinery flow conditions will be presented in future work.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89890 , E-3532 , NAS 1.15:89890 , AIAA PAPER 87-1316
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  • 175
    Publication Date: 2019-06-28
    Description: Three-dimensional, nonlinear numerical simulations are presented for the K-type and H-type transitions for channel flow. There are two objectives. The first is to establish firmly the resolution requirements for the various stages in the transition process. Comparisons between calculations on various grids suggest a set of guidelines for maintaining a physically meaningful calculation. The second objective is to map out the structure of the hairpin vortices which arise in K-type and H-type transitions in channel flow, to the latest stage currently feasible. Flow field details are presented for both a subcritical Reynolds number of 1500 and a supercritical Reynolds number of 8000. The diagnostics include illustrations of the vertical shear, streamwise and spanwise vorticity, helicity, vortex stretching, and vortex diffusion fields.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2667 , L-16204 , NAS 1.60:2667
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  • 176
    Publication Date: 2019-06-28
    Description: The numerical simulation of three-dimensional transonic flow about propeller blades is discussed. The equations for the unsteady potential flow about propellers is given for an arbitrary coordinate system. From this the small disturbance form of the equation is derived for a new helical coordinate system. The new coordinate system is suited to propeller flow and allows cascade boundary conditions to be applied straightforward. A numerical scheme is employed which solves the steady flow as an asymptotic limit of unsteady flow. Solutions are presented for subsonic and transonic flow about a 5 percent thick bicircular arc blade of an eight bladed cascade. Both high and low advance ratio cases are given which include a lifting case as well as nonlifting cases. The nonlifting cases are compared to solutions from a Euler code.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89826 , E-3475 , NAS 1.15:89826
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  • 177
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the Langley 16-Foot Transonic Tunnel at static conditions to measure the pressure distributions inside a nonaxisymmetric nozzle with simultaneous partial thrust reversing (50-percent deployment) and thrust vectoring of the primary (forward-thrust) nozzle flow. Geometric forward-thrust-vector angles of 0 and 15 deg. were tested. Test data were obtained at static conditions while nozzle pressure ratio was varied from 2.0 to 4.0. Results indicate that, unlike the 0 deg. vector angle nozzle, a complicated, asymmetric exhaust flow pattern exists in the primary-flow exhaust duct of the 15 deg. vectored nozzle.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89044 , L-16210 , NAS 1.15:89044
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  • 178
    Publication Date: 2019-06-28
    Description: A small-scale powered rotor model was designed for use as a research tool in the exploratory testing of rotors and helicopter models. The model, which consists of a 29 hp rotor drive system, a four-blade fully articulated rotor, and a fuselage, was designed to be simple to operate and maintain in wind tunnels of moderate size and complexity. Two six-component strain-gauge balances are used to provide independent measurement of the rotor and fuselage aerodynamic loads. Commercially available standardized hardware and equipment were used to the maximum extent possible, and specialized parts were designed so that they could be fabricated by normal methods without using highly specialized tooling. The model was used in a hover test of three rotors having different planforms and in a forward flight investigation of a 21-percent-scale model of a U.S. Army scout helicopter equipped with a mast-mounted sight.
    Keywords: AERODYNAMICS
    Type: NASA-TM-87762 , L-16165 , NAS 1.15:87762 , AVSCOM-TM-86-B-4 , AD-A178047
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  • 179
    Publication Date: 2019-06-28
    Description: Research in the area of computational, unsteady transonic flows about airfoils and wings, including aeroelastic effects is reviewed. In the last decade, there have been extensive developments in computational methods in response to the need for computer codes with which to study fundamental aerodynamic and aeroelastic problems in the critical transonic regime. For example, large commercial aircraft cruise most effectively in the transonic flight regime and computational fluid dynamics (CDF) provides a new tool, which can be used in combination with test facilities to reduce the costs, time, and risks of aircraft development.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89414 , A-87042 , NAS 1.15:89414
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  • 180
    Publication Date: 2019-06-28
    Description: The results of an investigation of the effects of far field boundary conditions on the solution of the three dimensional Euler equations governing the flow field of a high speed single rotation propeller are presented. The results show that the solutions obtained with the nonreflecting boundary conditions are in good agreement with experimental data. The specification of nonreflecting boundary conditions is effective in reducing the dependence of the solution on the location of the far field boundary. Details of the flow field within the blade passage and the tip vortex are presented. The dependence of the computed power coefficient on the blade passage and the tip vortex are presented. The dependence of the computed power coefficient on the blade setting angle is examined.
    Keywords: AERODYNAMICS
    Type: NASA-TM-88955 , E-3399 , NAS 1.15:88955
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  • 181
    Publication Date: 2019-06-28
    Description: Unsteady velocity field measurements made within the stator row of a transonic axial-flow fan are presented. Measurements were obtained at midspan for two different stator blade rows using a laser anemometer. The first stator row consists of double circular-arc airfoils with a solidity of 1.68. The second features controlled-diffusion airfoils with a solidity of 0.85. Both were tested at design-speed peak efficiency conditions. In addition, the controlled-diffusion stator was also tested at near stall conditions. The procedures developed here are used to identify the rotor wake generated and unresolved unsteadiness from the velocity measurements (rotor wake generated unsteadiness refers to the unsteadiness generated by the rotor wake velocity deficit and unresolved unsteadiness refers to all remaining unsteadiness which contributes to the spread in the distribution of velocities such as vortex shedding, turbulence, etc.). Auto and cross correlations of these unsteady velocity fluctuations are presented to show their relative magnitude and spatial distributions. Amplification and attenuation of both rotor wake generated and unresolved unsteadiness are shown to occur within the stator blade passage.
    Keywords: AERODYNAMICS
    Type: NASA-TM-88946 , E-3394 , NAS 1.15:88946 , USAAVSCOM-TR-86-C-31
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  • 182
    Publication Date: 2019-06-28
    Description: An interactive inviscid core flow-boundary layer method is presented for the calculation of turbomachine channel flows. For this method, a one-dimensional inviscid core flow is assumed. The end-wall and blade surface boundary layers are calculated using an integral entrainment method. The boundary layers are assumed to be collateral and thus are two-dimensional. The boundary layer equations are written in a streamline coordinate system. The streamwise velocity profiles are approximated by power law profiles. Compressibility is accounted for in the streamwise direction but not in the normal direction. Equations are derived for the special cases of conical and two-dimensional rectangular diffusers. For these cases, the assumptions of a one-dimensional core flow and collateral boundary layers are valid. Results using the method are compared with experiment and good quantitative agreement is obtained.
    Keywords: AERODYNAMICS
    Type: NASA-TM-88928 , E-3264 , NAS 1.15:88928 , USAAVSCOM-TR-86-C-36 , AD-A178164
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  • 183
    Publication Date: 2019-06-28
    Description: A simplified fourwall interference assessment method has been described, and a computer program developed to facilitate correction of the airfoil data obtained in the Langley 0.3-m Transonic Cryogenic Tunnel (TCT). The procedure adopted is to first apply a blockage correction due to sidewall boundary-layer effects by various methods. The sidewall boundary-layer corrected data are then used to calculate the top and bottom wall interference effects by the method of Capallier, Chevallier and Bouinol, using the measured wall pressure distribution and the model force coefficients. The interference corrections obtained by the present method have been compared with other methods and found to give good agreement for the experimental data obtained in the TCT with slotted top and bottom walls.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4042 , NAS 1.26:4042
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  • 184
    Publication Date: 2019-06-28
    Description: The use of a transonic airfoil code for analysis, inverse design, and direct optimization of an airfoil immersed in propfan slipstream is described. A summary of the theoretical method, program capabilities, input format, output variables, and program execution are described. Input data of sample test cases and the corresponding output are given.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4044 , NAS 1.26:4044 , KU-FRL-602-1
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  • 185
    Publication Date: 2019-06-28
    Description: A pressure experiment at high subsonic speeds was conducted by a cambered and twisted thick delta wing at the design condition (Mach number 0.80), as well as at nearby Mach numbers (0.75 and 0.83) and over an angle-of-attack range. Effects of twin vertical tails on the wing pressure measurements were also assessed. Comparisons of detailed theoretical and experimental surface pressures and sectional characteristics for the wing alone are presented. The theoretical codes employed are FLO-57, FLO-28, PAN AIR, and the Vortex Lattice Method-Suction Analogy.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2713 , L-16224 , NAS 1.60:2713
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  • 186
    Publication Date: 2019-06-28
    Description: The applicability of active control of transition by periodic suction-blowing is investigated via direct numerical simulations of the Navier-Stokes equations. The time-evolution of finite-amplitude disturbances in plane channel flow is compared in detail with and without control. The analysis indicates that, for relatively small three dimensional amplitudes, a two dimensional control effectively reduces disturbance growth rates even for linearly unstable Reynolds numbers. After the flow goes through secondary instability, three dimensional control seems necessary to stabilize the flow. An investigation of the temperature field suggests that passive temperature contamination is operative to reflect the flow dynamics during transition.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4095 , NAS 1.26:4095
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  • 187
    Publication Date: 2019-06-28
    Description: The aeropropulsive characteristics of an advanced twin-engine fighter aircraft designed for supersonic cruise have been studied in the Langley 16-Foot Tansonic Tunnel and the Lewis 10- by 10-Foot Supersonic Tunnel. The objective was to determine multiaxis control-power characteristics from thrust vectoring. A two-dimensional convergent-divergent nozzle was designed to provide yaw vector angles of 0, -10, and -20 deg combined with geometric pitch vector angles of 0 and 15 deg. Yaw thrust vectoring was provided by yaw flaps located in the nozzle sidewalls. Roll control was obtained from differential pitch vectoring. This investigation was conducted at Mach numbers from 0.20 to 2.47. Angle of attack was varied from 0 to about 19 deg, and nozzle pressure ratio was varied from about 1 (jet off) to 28, depending on Mach number. Increments in force or moment coefficient that result from pitch or yaw thrust vectoring remain essentially constant over the entire angle-of-attack range of all Mach numbers tested. There was no effect of pitch vectoring on the lateral aerodynamic forces and moments and only very small effects of yaw vectoring on the longitudinal aerodynamic forces and moments. This result indicates little cross-coupling of control forces and moments for combined pitch-yaw vectoring.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2712 , L-16213 , NAS 1.60:2712
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  • 188
    Publication Date: 2019-06-28
    Description: The purpose of the workshop was to discuss the current technology base for aerodynamic ground effects and to establish directions for further research of advanced, high performance aircraft designs, particularly those concepts utilizing powered lift systems; e.g., V/STOL, ASTOVL, and STOL aircraft. Fourteen papers were presented in the following areas: suckdown and fountain effects in hover; STOL ground vortex and hot gas ingestion; and vortex lift and jet flaps in ground effect. These subject areas were chosen with regard to current activities in the field of aircraft ground effects research.
    Keywords: AERODYNAMICS
    Type: NASA-CP-2462 , A-86391 , NAS 1.55:2462
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  • 189
    Publication Date: 2019-06-28
    Description: The hover performance of a 27 percent scale model baseline rotor and advanced rotor with a 3:1 tapered tip (TR3) for the AH-64 attack helicopter was investigated and compared. Hover results from a previously tested advanced rotor with a 5:1 tapered tip (TR5) were also compared. Rotor thrust was varied over a range for two tip Mach numbers. The results indicated that the TR3 blades had improved performance compared with the TR5 blades, and both the TR3 and TR5 blades were superior to the baseline rotor. The additional margin in performance for the TR3 blades was likely due to an increase in blade area and Reynolds number in the tip region of the blades.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89145 , L-16267 , NAS 1.15:89145 , AVSCOM-TM-87-B-10 , AD-A184287
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  • 190
    Publication Date: 2019-06-28
    Description: An exploratory investigation was conducted of the nonlinear aerodynamic and stability characteristics of a tailless generic fighter configuration featuring a chine-shaped forebody coupled to a slender cropped delta wing in the NASA Langley Research Center's 12-Foot Low-Speed Wind Tunnel. Forebody and wing vortex flow mechanisms were identified through off-body flow visualizations to explain the trends in the longitudinal and lateral-directional characteristics at extreme attitudes (angles of attack and sideslip). The interactions of the vortical motions with centerline and wing-mounted vertical tail surfaces were studied and the flow phenomena were correlated with the configuration forces and moments. Single degree of freedom, free-to-roll tests were used to study the wing rock susceptibility of the generic fighter model. Modifications to the nose region of the chine forebody were examined and fluid mechanisms were established to account for their ineffectiveness in modulating the highly interactive forebody and wing vortex systems.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89447 , A-87174 , NAS 1.15:89447
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  • 191
    Publication Date: 2019-06-28
    Description: A six degree-of-freedom (DOF) isolation system using six LIVE units has been installed under an Army/NASA contract on a Bell 206LM helicopter. This system has been named the Total Rotor Isolation System, or TRIS. To determine the effectiveness of TRIS in reducing helicopter vibration, a flight verification study was conducted at Bell's Flight Research Center in Arlington, Texas. The flight test data indicate that the 4/rev vibration level at the pilot's seat were suppressed below the 0.04g level throughout the transition envelope. Flight tests indicate over 95% suppression of vibration level from the rotor hub to the pilot's seat. The TRIS installation was designed with a decoupled control system and has shown a significant improvement in aircraft flying qualities, such that it permitted the trimmed aircraft to be flown hands-off for a significant period of time, over 90 seconds. The TRIS flight test program has demonstrated a system that greatly reduces vibration levels of a current-generation helicopter, while significantly improving the flying qualities to a point where stability augmentation is no longer a requirement.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4082 , NAS 1.26:4082 , REPT-699-099-055
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  • 192
    Publication Date: 2019-06-28
    Description: An experimental investigation was performed in which surface pressure data, flow visualization data, and force and moment data were obtained on four conical delta wing models which differed in leading edge camber only. Wing leading edge camber was achieved through a deflection of the outboard 30% of the local wing semispan of a reference 75 deg swept flat delta wing. The four wing models have leading edge deflection angles delta sub F of 0, 5, 10, and 15 deg measured streamwise. Data for the wings with delta sub F = 10 and 15 deg showed that hinge line separation dominated the lee-side wing loading and prohibited the development of leading edge separation on the deflected portion of wing leading edge. However, data for the wing with delta sub F = 5 deg showed that at an angle of attack of 5 deg, a vortex was positioned on the deflected leading edge with reattachment at the hinge line. Flow visualization results were presented which detail the influence of Mach number, angle of attack, and camber on the lee-side flow characteristics of conically cambered delta wings. Analysis of photographic data identified the existence of 12 distinctive lee-side flow types.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2660-PT-2 , L-16192 , NAS 1.60:2660-PT-2
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  • 193
    Publication Date: 2019-06-28
    Description: A wind tunnel study at Mach 0.4 was conducted for a slender wing-body configuration with a leading edge vortex flap of curved planform that is deflectable about a 74 degree swept hinge line. The basic data consist of a unique combination of longitudinal aerodynamic, surface pressure, and vortex flap hinge-moment measurements on a common model. The longitudinal aerodynamic, pressure and hinge-moment data are presented without analysis in tabular format. Plots of the tabulated pressure data are also given.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89101 , L-16265 , NAS 1.15:89101
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  • 194
    Publication Date: 2019-06-28
    Description: A hover test of a 0.658-scale model of a V-22 rotor and wing was conducted at the Outdoor Aerodynamic Research Facility at Ames Research Center. The primary objectives of the test were to obtain accurate measurements of the hover performance of the rotor system, and to measure the aerodynamic interactions between the rotor and wing. Data were acquired for rotor tip Mach numbers ranging from 0.1 to 0.73. This report presents data on rotor performance, rotor-wake downwash velocities, rotor system loads, wing forces and moments, and wing surface pressures.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89419 , A-87058 , NAS 1.15:89419
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  • 195
    Publication Date: 2019-06-28
    Description: High speed wind tunnel aerodynamic performance tests of the SR-7A advanced prop-fan have been completed in support of the Prop-Fan Test Assessment (PTA) flight test program. The test showed that the SR-7A model performed aerodynamically very well. At the cruise design condition, the SR-7A prop fan had a high measured net efficiency of 79.3 percent.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89917 , E-3610 , NAS 1.15:89917 , AIAA PAPER 87-1893
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  • 196
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the effect of a boattail angle and wedge-size trade on the performance of nonaxisymmetric wedge nozzles installed on a generic twin-engine fighter aircraft model. Test data were obtained at static conditions and at Mach numbers from 0.60 to 1.25. Angle of attack was held constant at 0 deg. High-pressure air was used to simulate jet exhaust, and the nozzle pressure ratio was varied from 1.0 (jet off) to slightly over 15.0. For the configurations studied, the results indicate that wedge size can be reduced without affecting aeropropulsive performance.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2717 , L-16248 , NAS 1.60:2717
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  • 197
    Publication Date: 2019-06-28
    Description: Described are the results of an experiment which determined whether flow conditioning screens and honeycombs would ice up in a closed-loop icing wind tunnel when placed downstream of the heat exchanger and upstream of the spray bars. The experiment was performed in the Icing Research Tunnel (IRT) at NASA Lewis Research Center. The investigation involved two separate tests: one to find the icing characteristics of flow conditioners in the IRT, and the second to find the icing characteristics of flow conditioners in the proposed rehabilitation of the Altitude Wind Tunnel (AWT). Both experiments showed that the heat exchanger removed nearly all of the icing cloud so that icing of the flow conditioners would cause no serious tunnel performance degradation during the course of a day's run. Only extremely cold conditions caused frost formation on the flow conditioners. The significance of this frost formation was minimized because frost buildup on the heat exchanger caused a much more severe pressure drop than did icing of the flow conditioners.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89824 , E-3474 , NAS 1.15:89824
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  • 198
    Publication Date: 2019-06-28
    Description: Standard techniques used to model chemically-reacting flows require an artificial viscosity for stability in the presence of strong shocks. The resulting shock is smeared over at least three computational cells, so that the thickness of the shock is dictated by the structure of the overall mesh and not the shock physics. A gas passing through a strong shock is thrown into a nonequilibrium state and subsequently relaxes down over some finite distance to an equilibrium end state. The artificial smearing of the shock envelops this relaxation zone which causes the chemical kinetics of the flow to be altered. A method is presented which can investigate these issues by following the chemical kinetics and flow kinetics of a gas passing through a fully resolved shock wave at hypersonic Mach numbers. A nonequilibrium chemistry model for air is incorporated into a spectral multidomain Navier-Stokes solution method. Since no artificial viscosity is needed for stability of the multidomain technique, the precise effect of this artifice on the chemical kinetics and relevant flow features can be determined.
    Keywords: AERODYNAMICS
    Type: NASA-CR-178308 , ICASE-87-35 , NAS 1.26:178308
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  • 199
    Publication Date: 2019-06-28
    Description: The hovering performance predictions of the TFAR1 and OPLIN codes and the experimental data are discussed. The TFAR1 program solves the full-potential equation in a rotor-fixed coordinate system by use of the line relaxation method. The OPLIN program calculates the positions of wake vortices and rotor performance using the influence-coefficient-and-lifing-line method. The two programs are combined by adding the induced velocities from the OPLIN code to the near flow-field of the rotor from the TFAR1 code. Results show that the TFAR1 program converges better with the downwash-coupling method than the twist correction method to include the wake downwash.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89494 , A-87193 , NAS 1.15:89494
    Format: application/pdf
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  • 200
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    In:  CASI
    Publication Date: 2019-06-28
    Description: The Boussinesq approximation is extended so as to explicitly account for the transfer of fluid energy through viscous action into thermal energy. Ideal and dissipative integral invariants are discussed, in addition to the general equations for thermal-fluid motion.
    Keywords: AERODYNAMICS
    Type: NASA-CR-178242 , NAS 1.26:178242
    Format: application/pdf
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