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  • Other Sources  (729)
  • AERODYNAMICS  (546)
  • COMPOSITE MATERIALS  (183)
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  • 1985-1989  (729)
  • 1940-1944
  • 1988  (729)
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  • 1985-1989  (729)
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  • 101
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    Detailed flowfield measurements over a 75 deg swept delta wing for code validation (1988)
    In:  CASI
    Details
    Publication Date: 2017-10-02
    Description: Selected results from an experimental investigation documenting the flowfield over 75 deg swept delta wing at an angle-of-attack of 20.5 deg are presented. Results obtained in the investigation include surface flow visualization, off-body flow visualization, and detailed flowfield surveys for various Reynolds numbers. Flowfield surveys at Reynolds numbers of 0.5, 1.0 and 1.5 million were conducted with both a pitot pressure probe and a 5-hole pressure probe; and 3-component laser Doppler velocimeter surveys were conducted at a Reynolds number of 1.0 million. The pitot pressure surveys were obtained at 5 longitudinal stations, the 5-hole probe surveys were obtained at 3 longitudinal stations and the laser Doppler velocimeter surveys were obtained at one station. The accuracy of each instrumentation system is discussed, as well as, discrepancies in the calculation of vorticity using various algorithms.
    Keywords: AERODYNAMICS
    Type: AGARD, Validation of Computational Fluid Dynamics. Volume 2: Poster Papers; 14 p
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  • 102
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    Experiments and code validation for juncture flows (1988)
    In:  CASI
    Details
    Publication Date: 2017-10-02
    Description: The turbulent flow around a juncture formed by an unswept wing and a flat plate has been experimentally studied, and the effectiveness of modifications near the wing leading edge in controlling the juncture flow field has been evaluated. The results are compared with numerical solutions of the incompressible Reynolds-averaged Navier-Stokes equations. The Baldwin-Lomax turbulence model is used in the computations. The numerical code is very time efficient, and it predicts the flow behavior well, including the detection of leading-edge vortex formation. It tends to over-predict the boundary layer thickness and the location of the vortex. Both the experiment and computations indicate that the leading edge flow separation is eliminated by the use of a leading-edge fillet designed in this study, resulting in drag reduction.
    Keywords: AERODYNAMICS
    Type: AGARD, Validation of Computational Fluid Dynamics. Volume 2: Poster Papers; 11 p
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  • 103
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    Experimental program for real gas flow code validation at NASA Ames Research Center (1988)
    In:  CASI
    Details
    Publication Date: 2017-10-02
    Description: The experimental program for validating real gas hypersonic flow codes at NASA Ames is described. Ground based test facilities used include ballistic ranges, shock tubes and shock tunnels, arcjet facilities and heated air hypersonic wind tunnels. Also included are large scale computer systems for kinetic theory simulations and benchmark code solutions. Flight tests consist of the Aeroassist Flight Experiment, the Space Shuttle, Project Fire 2, and planetary probes such as Galileo, Pioneer Venus and PAET.
    Keywords: AERODYNAMICS
    Type: AGARD, Validation of Computational Fluid Dynamics. Volume 1: Symposium Papers and Round Table Discussion; 16 p
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  • 104
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    On the validation of a code and a turbulence model appropriate to circular control airfoils (1988)
    In:  CASI
    Details
    Publication Date: 2017-10-02
    Description: A computer code for calculating flow about a circulation control airfoil within a wind tunnel test section was developed. This code is being validated for eventual use as an aid to design such airfoils. The concept of code validation being used is explained. The initial stages of the process were accomplished. The present code was applied to a low subsonic, 2-D flow about a circulation control airfoil for which extensive data exist. Two basic turbulence models and variants thereof were successfully introduced into the algorithm, the Baldwin-Lomax algebraic and the Jones-Launder two equation models of turbulence. The variants include adding a history of the jet development for the algebraic model and adding streamwise curvature effects for both models. Numerical difficulties and difficulties in the validation process are discussed. Turbulence model and code improvements to proceed with the validation process are also discussed.
    Keywords: AERODYNAMICS
    Type: AGARD, Validation of Computational Fluid Dynamics. Volume 1: Symposium Papers and Round Table Discussion; 22 p
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  • 105
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    Parametric study of grid size, time step and turbulence modeling on Navier-Stokes computations over airfoils (1988)
    In:  CASI
    Details
    Publication Date: 2017-10-02
    Description: An upwind-biased implicit approximate factorization algorithm is applied to several steady and unsteady turbulent flows. The thin layer form of the compressible Navier-Stokes equation is used. Both the flux vector splitting and flux difference splitting methods are used to determine fluxes, and the results are compared. Flux difference splitting predicts results more accurately than flux vector splitting on a given mesh size, but, in its present implementation, is more severely limited by the maximum CFL number for unsteady time accurate flows. Physical aspects of the computations are also examined. An equilibrium turbulent boundary layer model computes generally better steady and unsteady results than a nonequilibrium model when there is little to no boundary layer separation. Conversely, when a significant region of separation exists, the nonequilibrium model performs in better agreement with experiment.
    Keywords: AERODYNAMICS
    Type: AGARD, Validation of Computational Fluid Dynamics. Volume 1: Symposium Papers and Round Table Discussion; 19 p
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  • 106
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    A critical assessment of wind tunnel results for the NACA 0012 airfoil (1988)
    In:  CASI
    Details
    Publication Date: 2017-10-02
    Description: A large body of experimental results, obtained in more than 40 wind tunnels on a single, well known two-dimensional configuration, was critically examined and correlated. An assessment of some of the possible sources of error was made for each facility, and data which are suspect were identified. It was found that no single experiment provided a complete set of reliable data, although an investigation stands out as superior in many respects. However, from the aggregate of data the representative properties of the NACA 0012 airfoil can be identified with reasonable confidence over wide range of Mach numbers, Reynolds number, and angles of attack. This synthesized information can now be used to assess and validate existing and future wind tunnel results and to evaluate advanced Computational Fluid Dynamics codes.
    Keywords: AERODYNAMICS
    Type: AGARD, Aerodynamic Data Accuracy and Quality: Requirements and Capabilities in Wind Tunnel Testing; 21 p
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  • 107
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    Creep behavior of tungsten/niobium and tungsten/niobium-1 percent zirconium composites (1988)
    In:  CASI
    Details
    Publication Date: 2017-10-02
    Description: A study was conducted to determine the feasibility of using tungsten fiber reinforced niobium or niobium-1 percent zirconium matrix composites to meet the anticipated increased temperature and creep resistance requirements imposed by advanced space power systems. The results obtained on the short time tensile properties indicated that W/Nb composites showed significant improvements in high temperature strength and offer significant mass reductions for high temperature space power systems. The prime material requirement for space power systems applications is long time creep resistance. A study was conducted to determine the effect of high temperature exposure on the properties of these composites, with emphasis upon their creep behavior at elevated temperatures.
    Keywords: COMPOSITE MATERIALS
    Type: New Mexico Univ., Transactions of the Fifth Symposium on Space Nuclear Power Systems; p 267-272
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  • 108
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    A review of tilt rotor download research (1988)
    In:  Other Sources
    Details
    Publication Date: 2018-12-01
    Description: Experimental and theoretical research on the forces on a wing immersed in the wake of a hovering rotor is reviewed, with emphasis on the tilt rotor download problem. The basic features of the rotor/wing flow field on a tilt rotor aircraft are described. The effect of important geometric and operational parameters on the wing download is assessed. The magnitude of the download for typical tilt rotor configurations is reviewed, and advanced concepts for download reduction are described. Recommendations are presented for the direction of future research efforts.
    Keywords: AERODYNAMICS
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  • 109
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    Toward lower drag with laminar flow technology (1988)
    In:  Other Sources
    Details
    Publication Date: 2018-12-01
    Description: The most promising method of achieving significantly lower skin-friction drag of an aircraft component is through the stabilization and maintenance of a laminar boundary layer over a large fraction of its surface. In a number of applications, however, laminar flow is unfeasible and/or undesirable, and the most efficient turbulent flow design must be provided. The development of both laminar and turbulent airfoil and wing designs over the past decade is reviewed. High, medium, and low-speed airfoil concepts are discussed, and some overall comparisons are made for different Reynolds number ranges.
    Keywords: AERODYNAMICS
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  • 110
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    Viscous shock-layer solutions for the low-density hypersonic flow past long slender bodies (1988)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Results are obtained for the surface pressure, drag, heat-transfer, and skin-friction coefficients for hyperboloids and sphere cones. Body half angles from 5 to 22.5 degrees are considered for various low-density flow conditions. Recently obtained surface-slip and shock-slip equations are employed to account for the low-density effects. The method of solution employed for the viscous shock-layer (VSL) equations is a partially coupled spatial-marching implicit finite-difference technique. The flow cases analyzed include highly cooled long slender bodies in high Mach number flows. The present perfect-gas VSL calculations compare quite well with available experimental data. Results have also been obtained from the steady-state Navier-Stokes (NS) equations by successive approximations. Comparison between the NS and VSL results indicates that VSL equations even with body and shock-slip boundary conditions may not be adequate in the stagnation region at altitudes greater than about 75 km for the cases analyzed here.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0460
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  • 111
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    Turbulent swirling jets with excitation (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An existing cold-jet facility at NASA Lewis Research Center was modified to produce swirling flows with controllable initial tangential velocity distribution. Two extreme swirl profiles, i.e., one with solid-body rotation and the other predominated by a free-vortex distribution, were produced at identical swirl number of 0.48. Mean centerline velocity decay characteristics of the solid-body rotation jet flow exhibited classical decay features of a swirling jet with S - 0.48 reported in the literature. However, the predominantly free-vortex distribution case was on the verge of vortex breakdown, a phenomenon associated with the rotating flows of significantly higher swirl numbers, i.e., S sub crit greater than or equal to 0.06. This remarkable result leads to the conclusion that the integrated swirl effect, reflected in the swirl number, is inadequate in describing the mean swirling jet behavior in the near field. The relative size (i.e., diameter) of the vortex core emerging from the nozzle and the corresponding tangential velocity distribution are also controlling factors. Excitability of swirling jets is also investigated by exciting a flow with a swirl number of 0.35 by plane acoustic waves at a constant sound pressure level and at various frequencies. It is observed that the cold swirling jet is excitable by plane waves, and that the instability waves grow about 50 percent less in peak r.m.s. amplitude and saturate further upstream compared to corresponding waves in a jet without swirl having the same axial mass flux. The preferred Strouhal number based on the mass-averaged axial velocity and nozzle exit diameter for both swirling and nonswirling flows is 0.4.
    Keywords: AERODYNAMICS
    Type: NASA-CR-180895 , NAS 1.26:180895
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  • 112
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    Three-dimensional singular points in aerodynamics (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: When three-dimensional separation occurs on a body immersed in a flow governed by the incompressible Navier-Stokes equations, the geometrical surfaces formed by the three vector fields (velocity, vorticity and the skin-friction) and a scalar field (pressure) become interrelated through topological maps containing their respective singular points and extremal points. A mathematically consistent description of these singular points becomes inevitable when we want to study the geometry of the separation. A separated stream surface requires, for example, the existence of a saddle-type singular point on the skin-friction surface. This singular point is actually, in the proper language of mathematics, a saddle of index two. The index is a measure of the dimension of the outset (set leaving the singular point). Hence, when a saddle of index two is specified, a two dimensional surface that becomes separated from the osculating plane of the saddle is implied. The three-dimensional singular point is interpreted mathematically and the most common aerodynamical singular points are discussed through this perspective.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100045 , A-88029 , NAS 1.15:100045 , AD-A197978 , USAAVSCOM-TR-87-A-14
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  • 113
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    Tip aerodynamics and acoustics test: A report and data survey (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: In a continuing effort to understand helicopter rotor tip aerodynamics and acoustics, a flight test was conducted by NASA Ames Research Center. The test was performed using the NASA White Cobra and a set of highly instrumented blades. All aspects of the flight test instrumentation and test procedures are explained. Additionally, complete data sets for selected test points are presented and analyzed. Because of the high volume of data acquired, only selected data points are presented. However, access to the entire data set is available to the researcher on request.
    Keywords: AERODYNAMICS
    Type: NASA-RP-1179 , A-87128 , NAS 1.61:1179
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  • 114
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    Wake model for helicopter rotors in high speed flight (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Two alternative approaches are developed to calculate blade-vortex interaction airloads on helicopter rotors: second order lifting-line theory, and a lifting surface theory correction. The common approach of using a larger vortex core radius to account for lifting-surface effects is quantified. The second order lifting-line theory also improves the modeling of yawed flow and swept tips. Calculated results are compared with wind tunnel measurements of lateral flapping, and with flight test measurements of blade section lift on SA349/2 and H-34 helicopter rotors. The tip vortex core radius required for good correlation with the flight test data is about 20 percent chord, which is within the range of measured viscous core sizes for helicopter rotors.
    Keywords: AERODYNAMICS
    Type: NASA-CR-177507 , USAVSCOM-TR-88-A-008 , NAS 1.26:177507
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  • 115
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    Theoretical studies on flapped delta wings (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The effects of leading edge flaps on the aerodynamic characteristics of a low aspect-ratio delta wing are studied theoretically. As an extension of the classical crossflow plane analysis and in order to include separated shear layers, an analogy between three dimensional steady conical and two dimensional unsteady self-similar flows is explored. This analogy provides a simple steady-unsteady relationship. The criteria for the validity of the steady-unsteady analogy are also examined. Two different theoretical techniques are used to represent the separated shear layers based on the steady-unsteady analogy, neglecting the trailing edge effect. In the first approach, each vortex system is represented by a pair of concentrated vortices connected to the separation points by straight feeding sheets. In the second approach, the vortex cloud method is adopted for simulating the flow field in the crossflow plane. The separated shear layers are replaced with a cloud of discrete vortices and the boundary element method is employed to represent the wing trace by a vorticity distribution. A simple merging scheme is used to model the core region of the vortical flow as a single vortex by imposing a restriction on the shear layer rotation angle. The results are compared with experiments and with results from 3-D panel calculations.
    Keywords: AERODYNAMICS
    Type: NASA-CR-184795 , NAS 1.26:184795 , JIAA-TR-85
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  • 116
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    Flutter analysis of highly swept delta wings by conventional methods (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The flutter boundaries of six thin highly-swept delta-platform wings have been calculated. Comparisons are made between experimental data and results using several aerodynamic methods. The aerodynamic methods used include a subsonic and supersonic kernel function, second order piston theory, and a transonic small disturbance code. The dynamic equations of motion are solved using analytically calculated mode shapes and frequencies.
    Keywords: AERODYNAMICS
    Type: NASA-TM-101530 , NAS 1.15:101530
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  • 117
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    Nonlinear interaction of near-planar TS waves and longitudinal vortices in boundary-layer transition (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The nonlinear interactions that evolve between a planar or nearly planar Tollmien-Schlichting (TS) wave and the associated longitudinal vortices are considered theoretically for a boundary layer at high Reynolds number. The vortex flow is either induced by the TS nonlinear forcing or is input upstream, and similarly for the nonlinear wave development. Three major kinds of nonlinear spatial evolution, Types 1-3, are found. Each can start from secondary instability and then become nonlinear, Type 1 proving to be relatively benign but able to act as a pre-cursor to the Types 2, 3 which turn out to be very powerful nonlinear interactions. Type 2 involves faster stream-wise dependence and leads to a finite-distance blow-up in the amplitudes, which then triggers the full nonlinear 3-D triple-deck response, thus entirely altering the mean-flow profile locally. In contrast, Type 3 involves slower streamwise dependence but a faster spanwise response, with a small TS amplitude thereby causing an enhanced vortex effect which, again, is substantial enough to entirely alter the meanflow profile, on a more global scale. Streak-like formations in which there is localized concentration of streamwise vorticity and/or wave amplitude can appear, and certain of the nonlinear features also suggest by-pass processes for transition and significant changes in the flow structure downstream. The powerful nonlinear 3-D interactions 2, 3 are potentially very relevant to experimental findings in transition.
    Keywords: AERODYNAMICS
    Type: NASA-CR-181751 , ICASE-88-66 , NAS 1.26:181751
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  • 118
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    Shock tube investigation of dynamic response of pressure transducers for validation of rotor performance measurements (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: For some time now, NASA has had a program under way to aid in the validation of rotor performance and acoustics codes associated with the UH-60 rotary-wing aircraft; and to correlate results of such studies with those obtained from investigations of other selected aircraft rotor performance. A central feature of these studies concerns the dynamic measurement of surface pressure at various locations up to frequencies of 25 KHz. For this purpose, fast-response gauges of the Kulite type are employed. The latter need to be buried in the rotor; they record surface pressures which are transmitted by a pipette connected to the gauge. The other end of the pipette is cut flush with the surface. In certain locations, the pipette configuration includes a rather sharp right-angle bend. The natural question has arisen in this connection: In what way are the pipettes modifying the signals received at the rotor surface and subsequently transmitted to the sensitive Kulite transducer element. The basic details and results of the program performed and recently completed in the High Pressure Shock Tube Laboratory of the Department of Aeronautics and Astronautics at Stanford University are given.
    Keywords: AERODYNAMICS
    Type: NASA-CR-182673 , NAS 1.26:182673 , SU-AERO-48-88
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  • 119
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    High Lift, Low Pitching Moment Airfoils (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Two families of airfoil sections which can be used for helicopter/rotorcraft rotor blades or aircraft propellers of a particular shape are prepared. An airfoil of either family is one which could be produced by the combination of a camber line and a thickness distribution or a thickness distribution which is scaled from these. An airfoil of either family has a unique and improved aerodynamic performance. The airfoils of either family are intended for use as inboard sections of a helicopter rotor blade or an aircraft propeller.
    Keywords: AERODYNAMICS
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  • 120
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    Turbofan forced mixer lobe flow modeling. 2: Three-dimensional inviscid mixer analysis (FLOMIX) (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A three-dimensional potential analysis (FLOMIX) was formulated and applied to the inviscid flow over a turbofan foced mixer. The method uses a small disturbance formulation to analytically uncouple the circumferential flow from the radial and axial flow problem, thereby reducing the analysis to the solution of a series of axisymmetric problems. These equations are discretized using a flux volume formulation along a Cartesian grid. The method extends earlier applications of the Cartesian method to complex cambered geometries. The effects of power addition are also included within the potential formulation. Good agreement is obtained with an alternate small disturbance analysis for a high penetration symmetric mixer in a planar duct. In addition, calculations showing pressure distributions and induced secondary vorticity fields are presented for practical trubofan mixer configurations, and where possible, comparison was made with available experimental data. A detailed description of the required data input and coordinate definition is presented along with a sample data set for a practical forced mixer configuration. A brief description of the program structure and subroutines is also provided.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4147-PT-2 , E-4084 , NAS 1.26:4147-PT-2
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  • 121
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    Turbofan forced mixer lobe flow modeling. 1: Experimental and analytical assessment (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A joint analytical and experimental investigation of three-dimensional flowfield development within the lobe region of turbofan forced mixer nozzles is described. The objective was to develop a method for predicting the lobe exit flowfield. In the analytical approach, a linearized inviscid aerodynamical theory was used for representing the axial and secondary flows within the three-dimensional convoluted mixer lobes and three-dimensional boundary layer analysis was applied thereafter to account for viscous effects. The experimental phase of the program employed three planar mixer lobe models having different waveform shapes and lobe heights for which detailed measurements were made of the three-dimensional velocity field and total pressure field at the lobe exit plane. Velocity data was obtained using Laser Doppler Velocimetry (LDV) and total pressure probing and hot wire anemometry were employed to define exit plane total pressure and boundary layer development. Comparison of data and analysis was performed to assess analytical model prediction accuracy. As a result of this study a planar mixed geometry analysis was developed. A principal conclusion is that the global mixer lobe flowfield is inviscid and can be predicted from an inviscid analysis and Kutta condition.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4147-PT-1 , E-4083 , NAS 1.26:4147-PT-1
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  • 122
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    Two-dimensional aerodynamic characteristics of the OLS/TAAT airfoil (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Two flight tests have been conducted that obtained extension pressure data on a modified AH-1G rotor system. These two tests, the Operational Loads Survey (OLS) and the Tip Aerodynamics and Acoustics Test (TAAT) used the same rotor set. In the analysis of these data bases, accurate 2-D airfoil data is invaluable, for not only does it allow comparison studies between 2- and 3-D flow, but also provides accurate tables of the airfoil characteristics for use in comprehensive rotorcraft analysis codes. To provide this 2-D data base, a model of the OLS/TAAT airfoil was tested over a Reynolds number range from 3 x 10 to the 6th to 7 x 10 to the 7th and between Mach numbers of 0.34 to 0.88 in the NASA Langley Research Center's 6- by 28-Inch Transonic Tunnel. The 2-D airfoil data is presented as chordwise pressure coefficient plots, as well as lift, drag, and pitching moment coefficient plots and tables.
    Keywords: AERODYNAMICS
    Type: NASA-TM-89435 , A-87132 , NAS 1.15:89435
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  • 123
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    Transonic aeroelastic analysis of launch vehicle configurations (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A numerical study of the aeroelastic stability of typical launch vehicle configurations in transonic flight is performed. Recent computational fluid dynamics techniques are used to simulate the transonic aerodynamic flow fields, as opposed to relying on experimental data for the unsteady aerodynamic pressures. The flow solver is coupled to an appropriate structural representation of the vehicle. The aerodynamic formulation is based on the thin layer approximation to the Reynolds-Averaged Navier-Stokes equations, where the account for turbulent mixing is done by the two-layer Baldwin and Lomax algebraic eddy viscosity model. The structural-dynamic equations are developed considering free-free flexural vibration of an elongated beam with variable properties and are cast in modal form. Aeroelastic analyses are performed by integrating simultaneously in the two sets of equations. By tracing the growth or decay of a perturbed oscillation, the aeroelastic stability of a given constant configuration can be ascertained. The method is described in detail, and results that indicate its application are presented. Applications include some validation cases for the algorithm developed, as well as the study of configurations known to have presented flutter programs in the past.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4186 , NAS 1.26:4186
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  • 124
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    Thrust-reverser flow investigation on a twin-engine transport (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the NASA Langley 14 x 22 foot Subsonic Tunnel to study the effects of engine thrust reversing on an aft-mounted twin-engine transport and to develop effective testing techniques. Testing was done over a fixed and a moving-belt ground plane and over a pressure instrumented ground board. Free-stream dynamic pressure was set at values up to 12.2 psf, which corresponded to a maximum Reynolds number based on the mean aerodynamic chord of 765,000. The thrust reversers examined included cascade, target and four-door configurations. The investigation focused on the range of free-stream velocities and engine thrust-reverser flow rates that would be typical for landing ground-roll conditions. Flow visualization techniques were investigated, and the use of water or smoke injected into the reverser flow proved effective to determine the forward progression of the reversed flow and reingestion limits. When testing over a moving-belt ground plane, as opposed to a fixed ground plane, forward penetration of the reversed flow was reduced. The use of a pressure-instrumented ground board enabled reversed flow ground velocities to be obtained, and it provided a means by which to identify the reversed flow impingement point on the ground.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2856 , L-16426 , NAS 1.60:2856
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  • 125
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    Design of high-Reynolds-number flat-plate experiments in the NTF (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The design of an experiment to measure skin friction and turbulent boundary layer characteristics at Reynolds numbers exceeding 1 x 10 to the 9th is described. The experiment will be conducted in a zero-pressure-gradient flow on a flat plate in the National Transonic Facility (NTF). The development of computational codes to analyze the aerodynamic loads and the blockage is documented. Novel instrumentation techniques and models, designed to operate in cryogenic environments, are presented. Special problems associated with aerodynamic loads, surface finish, and hot-wire anemometers are discussed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-184627 , NAS 1.26:184627
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    Aerodynamic design via control theory (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The question of how to modify aerodynamic design in order to improve performance is addressed. Representative examples are given to demonstrate the computational feasibility of using control theory for such a purpose. An introduction and historical survey of the subject is included.
    Keywords: AERODYNAMICS
    Type: NASA-CR-181749 , ICASE-88-64 , NAS 1.26:181749 , AIAA PAPER 89-0434-REV
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    Development of three-dimensional code for the analysis of jet mixing problem. Part 1: Laminar solution (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Future aircraft will eventually feature nonaxisymmetric or rectangular nozzles. Developing a three-dimensional code to stimulate the characteristics of the jet exhaust plume, issuing from nonaxisymmetric nozzles, in general, at different flight conditions, is very important. Two three-dimensional codes were developed to simulate the shock-cell structure of circular nozzles. These codes were developed to solve the parabolized and simplified Navier-Stokes equations respectively. Both codes are based on a method previously developed by Newsome et al. These codes are fully vectorized on the VPS 32 at NASA Langley Research Center. The axisymmetric underexpanded supersonic jet flow problem, exhausting into still air, was used as a test case for developing an efficient three-dimensional problems and preserving crossplane symmetry of the flow downstream of the jet exit.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4200 , NAS 1.26:4200
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  • 128
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    User's manual for an aerodynamic optimization scheeme that updates flow variables and design parameters simultaneously (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: This user's manual is presented for an aerodynamic optimization program that updates flow variables and design parameters simultaneously. The program was developed for solving constrained optimization problems in which the objective function and the constraint function are dependent on the solution of the nonlinear flow equations. The program was tested by applying it to the problem of optimizing propeller designs. Some reference to this particular application is therefore made in the manual. However, the optimization scheme is suitable for application to general aerodynamic design problems. A description of the approach used in the optimization scheme is first presented, followed by a description of the use of the program.
    Keywords: AERODYNAMICS
    Type: NASA-CR-182180 , NAS 1.26:182180
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    Wall adjustment strategy software for use with the NASA Langley 0.3-meter transonic cryogenic tunnel adaptive wall test section (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The Wall Adjustment Strategy (WAS) software provides successful on-line control of the 2-D flexible walled test section of the Langley 0.3-m Transonic Cryogenic Tunnel. This software package allows the level of operator intervention to be regulated as necessary for research and production type 2-D testing using and Adaptive Wall Test Section (AWTS). The software is designed to accept modification for future requirements, such as 3-D testing, with a minimum of complexity. The WAS software described is an attempt to provide a user friendly package which could be used to control any flexible walled AWTS. Control system constraints influence the details of data transfer, not the data type. Then this entire software package could be used in different control systems, if suitable interface software is available. A complete overview of the software highlights the data flow paths, the modular architecture of the software and the various operating and analysis modes available. A detailed description of the software modules includes listings of the code. A user's manual is provided to explain task generation, operating environment, user options and what to expect at execution.
    Keywords: AERODYNAMICS
    Type: NASA-CR-181694 , NAS 1.26:181694
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  • 130
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    A generalized chemistry version of SPARK (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An extension of the reacting H2-air computer code SPARK is presented, which enables the code to be used on any reacting flow problem. Routines are developed calculating in a general fashion, the reaction rates, and chemical Jacobians of any reacting system. In addition, an equilibrium routine is added so that the code will have frozen, finite rate, and equilibrium capabilities. The reaction rate for the species is determined from the law of mass action using Arrhenius expressions for the rate constants. The Jacobian routines are determined by numerically or analytically differentiating the law of mass action for each species. The equilibrium routine is based on a Gibbs free energy minimization routine. The routines are written in FORTRAN 77, with special consideration given to vectorization. Run times for the generalized routines are generally 20 percent slower than reaction specific routines. The numerical efficiency of the generalized analytical Jacobian, however, is nearly 300 percent better than the reaction specific numerical Jacobian used in SPARK.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4196 , NAS 1.26:4196
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    Theoretical derivation and calibration technique of a hemispherical-tipped, five-hole probe (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A technique is presented for the calibration of a hemispherical tipped 0.125 inch diameter 5-hole probe. The derivation of equations from the potential flow over a sphere relating the flow angle and velocity to pressure differentials measured by the probe is presented. The technique for acquiring the calibration data and the technique used to calculate the calibration coefficients are presented. The accuracy of the probe in both the uniform calibration flow field and the nonuniform flow field over a 75 degree swept delta wing is discussed.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4047 , L-16454 , NAS 1.15:4047
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    A water tunnel study of Gurney flaps (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Several Gurney flap configurations were tested in the NASA Langley 16 x 24 inch Water Tunnel. These devices provided an increased region of attached flow on a wing upper surface relative to the wing without the flaps. The recirculation region behind the flap was visualized and shown to be consistent with hypotheses stated in previous research. Although the test Reynolds number for this study was several orders of magnitude below those in previous investigations, the effect of the Gurney flaps is in qualitative agreement with them. This is as would be expected from first order effects for high lift devices.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4071 , L-16467 , NAS 1.15:4071
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    Analysis of supersonic plug nozzle flowfield and heat transfer (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A number of problems pertaining to the flowfield in a plug nozzle, designed as a supersonic thruster nozzle, with provision for cooling the plug with a coolant stream admitted parallel to the plug wall surface, were studied. First, an analysis was performed of the inviscid, nonturbulent, gas dynamic interaction between the primary hot stream and the secondary coolant stream. A numerical prediction code for establishing the resulting flowfield with a dividing surface between the two streams, for various combinations of stagnation and static properties of the two streams, was utilized for illustrating the nature of interactions. Secondly, skin friction coefficient, heat transfer coefficient and heat flux to the plug wall were analyzed under smooth flow conditions (without shocks or separation) for various coolant flow conditions. A numerical code was suitably modified and utilized for the determination of heat transfer parameters in a number of cases for which data are available. Thirdly, an analysis was initiated for modeling turbulence processes in transonic shock-boundary layer interaction without the appearance of flow separation.
    Keywords: AERODYNAMICS
    Type: NASA-CR-179554 , NAS 1.26:179554
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    Development and validation of an advanced low-order panel method (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A low-order potential-flow panel code, PMARC, for modeling complex three-dimensional geometries, is currently being developed at NASA Ames Research Center. The PMARC code was derived from a code named VSAERO that was developed for Ames Research Center by Analytical Methods, Inc. In addition to modeling potential flow over three-dimensional geometries, the present version of PMARC includes several advanced features such as an internal flow model, a simple jet wake model, and a time-stepping wake model. Data management within the code was optimized by the use of adjustable size arrays for rapidly changing the size capability of the code, reorganization of the output file and adopting a new plot file format. Preliminary versions of a geometry preprocessor and a geometry/aerodynamic data postprocessor are also available for use with PMARC. Several test cases are discussed to highlight the capabilities of the internal flow model, the jet wake model, and the time-stepping wake model.
    Keywords: AERODYNAMICS
    Type: NASA-TM-101024 , A-88275 , NAS 1.15:101024
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  • 135
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    Sidewall boundary-layer measurements with upstream suction in the Langley 0.3-meter transonic cryogenic tunnel (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The Langley 0.3 Meter Transonic Cryogenic Tunnel has provision for boundary removal from the sidewalls to reduce sidewall interference effects on the test data. The tests carried out to determine the change in the empty test section sidewall boundary layer thickness at the model station with upstream boundary layer mass removal are described. The boundary layer measurements showed that the upstream removal region is effective in reducing the boundary layer thickness at the model station. The boundary layer displacement thickness reduced from about 1.2 percent to about .4 percent of the test section width. The boundary layer velocity profiles followed a power law variation in the outer region and showed good correlation when plotted in terms of boundary layer momentum thickness.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4192 , NAS 1.26:4192
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    Structural efficiency study of composite wing rib structures (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A series of short stiffened panel designs which may be applied to a preliminary design assessment of an aircraft wing rib is presented. The computer program PASCO is used as the primary design and analysis tool to assess the structural efficiency and geometry of a tailored corrugated panel, a corrugated panel with a continuous laminate, a hat stiffened panel, a blade stiffened panel, and an unstiffened flat plate. To correct some of the shortcomings in the PASCO analysis when shear is present, a two step iterative process using the computer program VICON is used. The loadings considered include combinations of axial compression, shear, and lateral pressure. The loading ranges considered are broad enough such that the designs presented may be applied to other stiffened panel applications. An assessment is made of laminate variations, increased spacing, and nonoptimum geometric variations, including a beaded panel, on the design of the panels.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-CR-183004 , NAS 1.26:183004 , CCMS-88-18 , VPI-E-88-29
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  • 137
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    Experimental aerodynamic performance of advanced 40 deg-swept 10-blade propeller model at Mach 0.6 to 0.85 (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A propeller designated as SR-6, designed with 40 deg of sweep and 10 blades to cruise at Mach 0.8 at an altitude of 10.7 km (35,000 ft), was tested in the NASA Lewis Research Center's 8- by 6-Foot Wind Tunnel. This propeller was one of a series of advanced single rotation propeller models designed and tested as part of the NASA Advanced Turboprop Project. Design-point net efficiency was almost constant to Mach 0.75 but fell above this speed more rapidly than that of any previously tested advanced propeller. Alternative spinners that further reduced the near-hub interblade Mach numbers and relieved the observed hub choking improved performance above Mach 0.75. One spinner attained estimated SR-6 Design-point net deficiencies of 80.6 percent at Mach 0.75 and 79.2 percent at Mach 0.8, higher than the measured performance of any previously tested advanced single-rotation propeller at these speeds.
    Keywords: AERODYNAMICS
    Type: NASA-TM-88969 , E-3437 , NAS 1.15:88969
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    A mixed-mode bending apparatus for delamination testing (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A mixed-mode delamination test procedure was developed combining double cantilever beam mode I loading and end notch flexure mode II loading on a split unidirectional laminate. By loading the specimen with a lever, a single applied load simultaneously produces mode I and II bending loads on the specimen. This mixed mode bending (MMB) test was analyzed using both finite element procedures and beam theory to calculate the mode I and II components of strain energy release rate, G sub I and G sub II, respectively. The analyses showed that a wide range of G sub I/G sub II ratios could be produced by varying the applied load position on the loading lever. As the delamination extended, the G sub I/G sub II ratios varied by less than 5 percent. The simple beam theory equations were modified to account for the elastic interaction between the two arms of the specimen and to account for shear deformations. The resulting equations agreed closely with the finite element results and provide a basis for selection of G sub I/G sub II test ratios and a basis for computing the mode I and II components of measured delamination toughness. The MMB specimen analysis and test procedures were demonstrated using unidirectional laminates.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-TM-100662 , NAS 1.15:100662
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    Three component laser anemometer measurements in an annular cascade of core turbine vanes with contoured end wall (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The three mean velocity components were measured in a full-scale annular turbine stator cascade with contoured hub end wall using a newly developed laser anemometer system. The anemometer consists of a standard fringe configuration using fluorescent seed particles to measure the axial and tangential components. The radial component is measured with a scanning confocal Fabry-Perot interferometer. These two configurations are combined in a single optical system that can operate simultaneously in a backscatter mode through a single optical access port. Experimental measurements were obtained both within and downstream of the stator vane row and compared with calculations from a three-dimensional inviscid computer program. In addition, detailed calibration procedures are described that were used, prior to the experiment, to accurately determine the laser beam probe volume location relative to the cascade hardware.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2846 , E-4183 , NAS 1.60:2846
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  • 140
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    Validation of a pair of computer codes for estimation and optimization of subsonic aerodynamic performance of simple hinged-flap systems for thin swept wings (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Extensive correlations of computer code results with experimental data are employed to illustrate the use of linearized theory attached flow methods for the estimation and optimization of the aerodynamic performance of simple hinged flap systems. Use of attached flow methods is based on the premise that high levels of aerodynamic efficiency require a flow that is as nearly attached as circumstances permit. A variety of swept wing configurations are considered ranging from fighters to supersonic transports, all with leading- and trailing-edge flaps for enhancement of subsonic aerodynamic efficiency. The results indicate that linearized theory attached flow computer code methods provide a rational basis for the estimation and optimization of flap system aerodynamic performance at subsonic speeds. The analysis also indicates that vortex flap design is not an opposing approach but is closely related to attached flow design concepts. The successful vortex flap design actually suppresses the formation of detached vortices to produce a small vortex which is restricted almost entirely to the leading edge flap itself.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2828 , L-16428 , NAS 1.60:2828
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    Experimental investigation of the performance of a supersonic compressor cascade (1988)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Supersonic cascade wind tunnel results are presented for a linear, supersonic compressor cascade derived from the near-tip section of a high-throughflow axial flow compressor rotor over the inlet Mach number range of 1.30-1.71. Laser anemometry was used to obtain flow-velocity measurements showing the wave pattern in the entrance region. Attention is given to the unique-incidence relationship for this cascade, which relates the supersonic inlet Mach number to the inlet flow direction. An empirical correlation is obtained for the influence of the independent parameters of back pressure, axial velocity density ratio, and blade element performance.
    Keywords: AERODYNAMICS
    Type: ASME PAPER 88-GT-306
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    3-D plume flow computations with an upwind solver (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: In the present study multi-nozzle plume flows are computed with a three-dimensional, Navier-Stokes solver. Numerical simulations are performed with the flux-split, two-factor, time symptotic, viscous flow solver. The two factor splitting provides a stable three-dimensional solution procedure under ideal-gas assumptions. Viscous, ideal-gas solutions for a thin lip symmetrical nozzle are compared with experimental and numerical solutions. Computed solutions to axisymmetric and three-dimensional, multi-nozzle problems at various altitudes and flight conditions demonstrate flow field complexity and three-dimensional effects.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-3158
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    Configuration trade and code validation study on a conical hypersonic vehicle (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A study has been conducted on a generic wing-cone transatmospheric vehicle at Mach numbers form 2.5 to 4.5. The objectives of the study were to experimentally define the aerodynamic characteristics of the vehicle and evaluate several computational aerodynamic prediction methods through comparison with the experimental results. The baseline wing-cone configuration fuselage consisted of a 5 deg half-angle cone forebody, cylindrical midbody, and 9 deg truncated cone afterbody. The 4-percent-thick diamond airfoil wing had an aspect ratio of 1. Several configuration variables were investigated to provide trade information on canard, wing-position and incidence, vertical tail, and nose bluntness effects. Results of the study show that wing-position and wing-incidence effects on the longitudinal aerodynamic characteristics can be significantly influenced by wing-body interference. The use of positive wing incidence to provide favorable forebody orientation for possible inlet performance improvement is accompanied by trim drag and lift-drag ratio penalties. The lateral-directional stability characteristics were strongly influenced by the location of the vertical tails. The higher-order full-potential method provided better estimates of the aerodynamic characteristics than either the linearized supersonic potential method or the tangent-cone/tangent-wedge/shock-expansion on method.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-4505
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    Boundary layer crossflow stabilization of high subsonic speed LFC transport airplanes (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: With a view to the elaboration of the design of Mach 0.83 and 0.97 cruise-speed long-range aircraft employing LFC, a study is conducted of the laminar flow characteristics of supercritical airfoils of blunt leading-edge X88 type, for the case of lightly loaded wings that dispense with leading-edge flaps for low-speed operations. The boundary layer crossflow in the front acceleration zone of these airfoils' upper surface is optimally stabilized by suction in the upstream portion of the zone, yielding a crossflow that is neutrally stable.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-0275
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    Transonic flow modes of an axisymmetric blunt body (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An experimental investigation of transonic flow past a plane-nosed circular cylinder with a plane-nosed circular probe extended coaxially ahead is reported. The possibilities of significant transonic drag reduction and the fluid mechanic phenomena which occur are examined. The probe length and diameter and the approaching flow Mach number are the independent variables. The relations which exist among the probe/cylinder geometry, Mach number, and flow field as revealed by measurements of the drag forces acting on the body are explored.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-3536
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    Summary of low-speed wind tunnel results of several high-speed counterrotation propeller configurations (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The low speed aerodynamic performance characteristics of several advanced counterrotation pusher propeller configurations with cruise design Mach numbers of 0.72 and 0.80 were investigated in the NASA Low Speed Wind Tunnel. The tests were conducted at Mach numbers representative of the takeoff and landing flight regime. The investigation included: (1) the propeller performance characteristics over a range of blade angle settings and rotational speeds at a Mach number of 0.20; (2) the effect on the propeller performance of varying the axial rotor spacing and mismatching the power and rotational speeds on the propeller rotors; and (3) determining the reverse thrust performance characteristics at Mach numbers of 0.0, 0.10, 0.15 and 0.20. The results of the investigation indicated that the overall low speed performance of the counterrotation propeller configurations was reasonable.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-3149
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  • 147
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    A block space-marching method for the Navier-Stokes equations (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A solution procedure is presented for predominantly supersonic viscous flows. The procedure approximately solves the Navier-Stokes equations by marching blocks of grid points in the streamwise direction and solving the fully elliptic equations within each block. In this manner elliptic effects of limited streamwise extent may be accurately calculated. Results are presented for calculations of a Mach 2 laminar flat-plate shock/boundary-layer interaction, and for a Mach 10 hypervelocity interceptor.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-3199
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  • 148
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    A comparison of experimental and computational results for 5 and 10 degree cones at high Mach numbers (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Experiments were performed on 5 and 10 deg slender cones at a velocity of approximately 5 km/sec in the NASA-Ames ballistic ranges. The flowfields for the cones were computed using ideal-gas and chemical nonequilibrium-air parabolized Navier-Stokes codes. Experimentally determined drag coefficients and shock shapes are compared with the results of the computer codes. Both the flight-data analysis methods and the computational codes are examined to achieve the most meaningful comparison. Under the conditions of the experiments, skin-friction drag makes up approximately 50 percent of the total drag for the 5 deg cone and 30 percent of the total drag for the 10 deg cone. Computed drag coefficients of the 10 deg cone agree well with the experimental values; predictions fall below the experimental values for the 5 deg cone.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2705
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    A computational study of heat transfer for laminar hypervelocity flows (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Computational Fluid Dynamics (CFD) codes are routinely used to predict the flowfield and the heating environment around complex reentry configurations. At hypervelocities, where the velocity is greater than 3 km/sec, the AFWAL version of the blunt body code predicts the correct surface pressure distributions but underpredicts laminar wall heat fluxes. This study was performed to determine the reasons for the underprediction. The computer code chosen solves thin-layer Navier-Stokes equations in a time-asymptotic manner and assumes a constant isentropic exponent. Flowfields around a spherical configuration at various entry velocities are computed. The computed pressure distributions agree well with the tabulated, inviscid results of Lyubimov and Rusanov for entry velocities ranging from 0.6 to 5.92 km/sec. At hypervelocities, the calculated stagnation point heat transfer rates were lower by roughly fifty percent when compared to engineering correlations available in the literature. Good comparisons between heat transfer rates are obtained at hypervelocity entry conditions provided the CFD code is modified to include equilibrium air properties.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2666
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    Stagnation flow field analysis for an aeroassist flight experiment vehicle (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Continuum methods are used to analyze the stagnation flow field of the aeroassist flight experiment (AFE) vehicle. For the lower altitude portion of an AFE trajectory, the viscous shock-layer equations are employed. At higher altitudes, the full Navier-Stokes equations with chemical nonequilibrium and surface slip are used. Particular attention is given to the effect of surface catalyticity on surface heating, electron number density, and flow field structure.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2613
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    An implicit Navier-Stokes analysis of turbine rotor-stator interaction (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An implicit Navier-Stokes analysis using a single deforming mesh has been developed for the unsteady rotor-stator interaction problem. The technique has been used to simulate the flow through a turbine stator-rotor stage. Periodic two-dimensional solutions have been obtained using 1000 time steps per cycle without iteration at each time step. Computed surface pressure distributions compare favorably with experimental data available for this configuration.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-3090
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    Artificial dissipation models for hypersonic external flow (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Four artificial dissipation models which augment central difference schemes were examined for hypersonic external flows. The models were a first and third order dissipation model, a directionally scaled first and third order dissipation model, a flux limited dissipation model, and a flux difference split dissipation model. Each model was implemented in the lower-upper symmetric-Gauss-Seidel (LU-SGS) algorithm to solve the full Navier-Stokes equations. The latter two models can be regarded as total variation diminishing (TVD) schemes. Test results for model problems showed that the flux limited dissipation model was robust enough to predict a high speed blunt body flow with strong shock and expansion waves. The flux difference split dissipation model was capable of shock capturing with higher resolution, but was less robust. First and third order dissipation models turned out to be neither accurate nor robust enough for high Mach number flow computations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-3277
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    High temperature metal matrix composites for future aerospace systems (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Research was conducted on metal matrix composites and intermetallic matrix composites to understand their behavior under anticipated future operating conditions envisioned for aerospace power and propulsion systems of the 21st century. Extremes in environmental conditions, high temperature, long operating lives, and cyclic conditions dictate that the test evaluations not only include laboratory testing, but simulated flight conditions. The various processing techniques employed to fabricate composites are discussed along with the basic research underway to understand the behavior of high temperature composites, and the relationship of this research to future aerospace systems.
    Keywords: COMPOSITE MATERIALS
    Type: AIAA PAPER 88-3059
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  • 154
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    Turbulence modeling in hypersonic inlets (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A study is conducted to analyze the performance of different turbulence models when applied to flow through a Mach 7.4 hypersonic inlet. The analysis, which is two-dimensional, is done by comparing computational results from a Parabolized Navier-Stokes code and a full Navier-Stokes code, with experimental data. The McDonald-Camarata (MC) and Baldwin-Lomax (BL) models were the two zero-equation models used in the study. The Turbulent Kinetic Energy (TKE) model was chosen as a representative higher order model. The MC model, when run with transition of flow, provides a solution which compares excellently with the data. Transition has a first order effect on the overall solution provided by the code. The BL model predicts separation of flow in the inlet, which contradicts experimental findings. The TKE model does not perform any better than the MC and BL models, despite the fact that it is a higher order turbulence model. The BL and TKE models predict transition in the inlet at a location which is much earlier than observed in the experiment. This may be attributed to the empirical constants used to determine the point of transition.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2957
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  • 155
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    Hypersonic turbulent wall boundary layer computations (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The Baldwin-Lomax (1978) algebraic turbulence model was modified for hypersonic flow conditions. Two coefficients in the outer-layer eddy-viscosity model were determined as functions of Mach number and temperature ratio. By matching the solutions from the Baldwin-Lomax model to those from the Cebeci-Smith (1974) model for a flat plate at hypersonic speed, the new values of the coefficients were obtained. The results show that the values of C(cp) and C(kleb) are functions of both Mach number and wall temperature ratio. The C(cp) and C(kleb) variations with Mach number and wall temperature were used for the calculations of both a 4-deg wedge flow at Mach 18 and an axisymmetric Mach 20 nozzle flow. The Navier-Stokes equations with thin-layer approximation were solved for the above hypersonic flow conditions and the results were compared with existing experimental data. The agreement between the numerical solutions and the existing experimental data were good. The modified Baldwin-Lomax model thus is useful in the computations of hypersonic flows.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2829
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  • 156
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    Navier-Stokes structure of merged layer flow on the spherical nose of a space vehicle (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Hypersonic merged layer flow on the forepart of a spherical surface of a space vehicle has been investigated on the basis of the full steady-state Navier-Stokes equations using slip and temperature jump boundary conditions at the surface and free-stream conditions far from the surface. The shockwave-like structure was determined as part of the computations. Using an equivalent body concept, computations were carried out under conditions that the Aeroassist Flight Experiment (AFE) Vehicle would encounter at 15 and 20 seconds in its flight path. Emphasis was placed on understanding the basic nature of the flow structure under low density conditions. Particular attention was paid to the understanding of the structure of the outer shockwave-like region as the fluid expands around the sphere. Plots were drawn for flow profiles and surface characteristics to understand the role of dissipation processes in the merged layer of the spherical nose of the vehicle.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2692
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  • 157
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    Numerical simulation of hypersonic flow over an aeroassist flight experiment vehicle (1988)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Solutions of wind-tunnel and entry-flight flow around the vehicle are obtained from the Navier-Stokes equations coupled with the chemical species continuity equations if needed. The time-iterative method employs several techniques: shock fitting, chemistry-split ADI and an algebraic grid in conformal spherical-polar space. Sensitivities of the results to numerical parameters and to frozen, equilibrium and finite rate reactions are investigated in the forebody computation. Quantitative results are obtained for the shock layer and the near wake for the entire vehicle corresponding to both ground test and flight conditions. Complex flow characteristics are analyzed on the basis of the complete flowfield over the aerobrake and simplified afterbodies. The method is stable and cost effective, and has yielded shock locations and wall pressure distributions which are in good agreement with wind-tunnel data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2675
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  • 158
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    Porous wind tunnel corrections for counterrotation propeller testing (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Wind tunnel interference corrections have direct impact on measured propeller efficiency. A systematic series of wind tunnel tests was done in the porous-wall NASA Lewis 8- by 6-Foot Wind Tunnel to determine the wind tunnel interference corrections to the NASA Lewis counterrotation propeller test data. The test results were compared with calculations from a potential flow code to determine the interference corrections. At a Mach number of 0.8, the interference corrections resulted in a -0.008 Mach number correction which reduced the counterrotation propeller net efficiency data by 0.46 percent at the reduced Mach number. Additional wind tunnel tests were done to measure the effect of propeller thrust on wind tunnel wall interference. No wall interference corrections due to propeller thrust were found necessary for the high speed counterrotation propeller data obtained in the porous wall NASA Lewis 8- by 6-Foot Wind Tunnel.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2055
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  • 159
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    The computation of radiation from nonequilibrium hypersonic flows (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The results of the solution of the equations that describe a hypersonic ionized flow about an elliptically blunted cone are presented. The flow conditions correspond to those of the proposed Aeroassist Flight Experiment (AFE) vehicle at altitudes between the perigee at 78 km and the approximate limit of the continuum regime at 90 km. For the free-stream velocities of interest, about 9 km/sec, the flowfield is out of thermo-chemical equilibrium, electronically excited, ionized and radiating. The gas consists of eight-chemical species including free electrons. The thermal state of the gas is modeled with a translational-rotational temperature, four vibrational temperatures for the diatomic species and an electron-electronic temperature. The electronic excitation of molecules is included. The nonequilibrium air radiation from each fluid element is computed and the radiative heat flux at the body surface is determined. The stagnation point radiative heating result agrees with previous calculations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2678
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  • 160
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    Vectorization of a particle simulation method for hypersonic rarefied flow (1988)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: An efficient particle simulation technique for hypersonic rarefied flows is presented at an algorithmic and implementation level. The implementation is for a vector computer architecture, specifically the Cray-2. The method models an ideal diatomic Maxwell molecule with three translational and two rotational degrees of freedom. Algorithms are designed specifically for compatibility with fine grain parallelism by reducing the number of data dependencies in the computation. By insisting on this compatibility, the method is capable of performing simulation on a much larger scale than previously possible. A two-dimensional simulation of supersonic flow over a wedge is carried out for the near-continuum limit where the gas is in equilibrium and the ideal solution can be used as a check on the accuracy of the gas model employed in the method. Also, a three-dimensional, Mach 8, rarefied flow about a finite-span flat plate at a 45 degree angle of attack was simulated. It utilized over 10 to the 7th particles carried through 400 discrete time steps in less than one hour of Cray-2 CPU time. This problem was chosen to exhibit the capability of the method in handling a large number of particles and a true three-dimensional geometry.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2735
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  • 161
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    Viscous shock-layer solutions with nonequilibrium chemistry for hypersonic flows past slender bodies (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Laminar nonequilibrium heat transfer to slender vehicles is discussed, with heating-rate results presented as a ratio of the noncatalytic to the corresponding fully catalytic value to illustrate the maximum potential for a heating reduction in dissociated nonequilibrium flow at a given flight condition. Larger blunted cone half-angles are shown to produce the most significant nonequilibrium effects at distances beyond 100 nose radii, except in the fore-cone region. Increasing nose bluntness is found to produce large reductions in the ratio for the smaller cone angles at relatively large downstream surface lengths. It is noted that the nose radius and freestream density are not independent scaling parameters in nonequilibrium flow.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2709
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  • 162
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    Effects of surface pressures and streamline metrics on the calculation of laminar heating rates (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The effect of streamline geometry and pressure distributions on surface heating rates is examined for slender, spherically blunted cones. The modifications to the approximate aeroheating code include a curve fit of pressures computed by an Euler solution over a range of Mach numbers and cone angles. The streamline geometry is then found using the surface pressures and inviscid surface properties. Previously, streamlines were determined using the inviscid properties at the edge of the boundary layer when accounting for the effects of entropy-layer swallowing. Streamline calculations are now based on inviscid surface conditions rather than boundary-layer edge properties. However, the heating rates are calculated using inviscid properties at the edge of the boundary layer. Resulting heating rates compare favorably with solutions from the viscous-shock-layer equations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2708
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  • 163
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    Radiation heat transfer in the wake of a hypersonic vehicle (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Radiation heat transfer (RHT) from the wake of a hypersonic vehicle to its afterbody is evaluated from Gnoffo and Greene's (1987) calculated wake flowfield and the radiative properties of ionized high-temperature air with the calculated nonequilibrium composition. The 4.2-m aeroassisted flight experiment at an altitude of 75 km and velocity of 8900 m/s causes a 0.1-m-thick layer initially at T = 10,000 K and P = 1 kN/sq m to separate from the shoulder of the forebody heat shield and spread aft to form a wake at approximately T = 5000 K and P = 20 N/sq m. Gas in the separated flow region at approximately T = 3000 K and P = 10 N/sq m, recirculates about the afterbody. It is shown that the radiating layer, recirculating gases, and wake are optically thin for purposes of making engineering RHT calculations. Directional, spectral, and spatial variations of the radiation incident upon the afterbody are presented.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2634
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  • 164
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    An upwind parabolized Navier-Stokes code for chemically reacting flows (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A new upwind, parabolized Navier-Stokes (PNS) code has been developed to compute the hypersonic, viscous, chemically reacting flow around two-dimensional or axisymmetric bodies. The new code is an extension of the upwind (perfect gas) PNS code of Lawrence et al. (1986). The upwind algorithm is based on Roe's flux-difference splitting scheme which has been modified to account for real gas effects. The algorithm solves the gas dynamic and species continuity equations in a 'loosely' coupled manner. The new code has been validated by computing the laminar flow (at free stream Mach number 25) of chemically reacting air over a wedge and a cone. The results of these computations are compared with the results from a centrally-differenced, fully coupled, nonequilibrium PNS code. The agreement is excellent, except in the vicinity of the shock wave where the present code exhibits superior shock capturing capabilities.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2614
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  • 165
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    The numerical simulation of transonic separated flow about the complete F-16A (1988)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The thin-layer, Reynolds-averaged, Navier-Stokes equations are used to simulate the transonic viscous flow about the complete F-16A fighter aircraft. These computations demonstrate how computational fluid dynamics (CFD) can be used to simulate turbulent viscous flow about realistic aircraft geometries. A zonal grid approach is used to provide adequate viscous grid clustering on all aircraft surfaces. Zonal grids extend inside the F-16A inlet and up to the compressor face while power on conditions are modeled by employing a zonal grid extending from the exhaust nozzle to the far field. A simple solution adaptive grid procedure is used on the wing surface and good agreement with experimental data is obtained. Computations for the F-16A in side slip are also presented.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 88-2506
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  • 166
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    Analytical study of graphite-epoxy tube response to thermal loads (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The thermally-induced stresses and deformations in graphite-epoxy tubes with aluminum foil bonded to both inner and outer surfaces, and to the outer surface only are computed. Tubes fabricated from three material systems, T300/934, P75s/934, and P75s/BP907, and having a 1 inch inner radius and a lamination sequence of (+15/0 + or - 10/0)sub s are studied. Radial, axial, and circumferential stresses in the various layers of the tube, in the foil, and in the adhesive bonding the foil to the tubes are computed using an elasticity solution. The results indicate that the coatings have no detrimental effect on the stress state in the tube, particularly those stresses that lead to microcracking. The addition of the aluminum foil does, however, significantly influence the axial expansion of the T300/934 tube, the tube with the softer graphite fibers. The addition of foil can change the sign of the axial coefficient of thermal expansion. Twist tendencies of the tubes are only slightly affected by the addition of the coatings, but are of second order compared to the axial response.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-CR-183215 , NAS 1.26:183215 , CCMS-88-16 , VPI-E-88-25 , PB89-104160
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  • 167
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    A test for interfacial effects and stress transfer in ceramic matrix composites (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The efforts to fabricate single embedded filament specimens of carbon and SiC fibers were unsuccessful largely due to the thermal stresses resulting from differences in thermal coefficient of expansion. Other factors appear to have been involved including embrittlement of the metal substrate by the H2 gas in the chemical vapor deposition flow stream and reaction layers formed between the silicon carbide and the metal substrate. The carbon fiber may have been attacked by the CVD reactant. It is concluded that these differential stresses are so large as to make the embedded fiber test impractical for the study of interphase effects and stress transfer in fiber ceramic matrix systems.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-CR-182873 , NAS 1.26:182873
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  • 168
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    Recent advances in transonic computational aeroelasticity (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A transonic unsteady aerodynamic and aeroelasticity code called CAP-TSD was developed for application to realistic aircraft configurations. The code permits the calculation of steady and unsteady flows about complete aircraft configurations for aeroelastic analysis in the flutter critical transonic speed range. The CAP-TSD code uses a time accurate approximate factorization algorithm for solution of the unsteady transonic small disturbance potential equation. An overview is given of the CAP-TSD code development effort and results are presented which demonstrate various capabilities of the code. Calculations are presented for several configurations including the General Dynamics 1/9 scale F-16 aircraft model and the ONERA M6 wing. Calculations are also presented from a flutter analysis of a 45 deg sweptback wing which agrees well with the experimental data. Descriptions are presented of the CAP-TSD code and algorithm details along with results and comparisons which demonstrate these recent developments in transonic computational aeroelasticity.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100663 , NAS 1.15:100663
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  • 169
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    The present status and the future of missile aerodynamics (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Some recent developments in the state of the art in missile aerodynamics are reviewed. Among the subjects covered are: (1) tri-service/NASA data base, (2) wing-body interference, (3) nonlinear controls, (4) hypersonic transition, (5) vortex interference, (6) airbreathers, supersonic inlets, (7) store separation problems, (8) correlation of missile data, (9) CFD codes for complete configurations, (10) engineering prediction methods, and (11) future configurations. Suggestions are made for future research and development to advance the state of the art of missile aerodynamics.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100063 , A-87289 , NAS 1.15:100063
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  • 170
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    Three-dimensional Navier-Stokes simulations of turbine rotor-stator interaction (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Fluid flows within turbomachinery tend to be extremely complex in nature. Understanding such flows is crucial to improving current designs of turbomachinery. The computational approach can be used to great advantage in understanding flows in turbomachinery. A finite difference, unsteady, thin layer, Navier-Stokes approach to calculating the flow within an axial turbine stage is presented. The relative motion between the stator and rotor airfoils is made possible with the use of patched grids that move relative to each other. The calculation includes endwall and tip leakage effects. An introduction to the rotor-stator problem and sample results in the form of time averaged surface pressures are presented. The numerical data are compared with experimental data and the agreement between the two is found to be good.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100081 , A-88106 , NAS 1.15:100081
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  • 171
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    Analysis and Optimization of Truncated Scarf Nozzles Subject to External Flow Conditions (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Results of a calculation of an optimized truncated scarfed nozzle were compared. The truncated scarfed nozzle was designed for an exit Mach number of 6.0, i.e., the Mach number at the last nozzle characteristic is 6.0, with an external flow Mach number of 5.0. The nozzle was designed by the Rao method for optimum thrust nozzles modified for 2-D flow and truncated scarfed nozzle applications. This design was analyzed using a shock-fitting method for 2-D supersonic flows. Excellent agreement was achieved between the design and analysis. Truncation of the lower nozzle wall (cowl) revealed that there is an optimum length for truncating the cowl without degrading the nozzle performance. Truncation of the nozzle cowl past this optimal length should be analyzed in trade-off studies for thrust loss versus gross vehicle weight. Plots of the oblique shock wave equations were also identified which will allow computation of slip line angle, dynamic pressure coefficient, or ambient Mach number for various specific heat ratios.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100955 , E-4146 , NAS 1.15:100955
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  • 172
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    Adaptive mesh generation for viscous flows using Delaunay triangulation (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A method for generating an unstructured triangular mesh in two dimensions, suitable for computing high Reynolds number flows over arbitrary configurations is presented. The method is based on a Delaunay triangulation, which is performed in a locally stretched space, in order to obtain very high aspect ratio triangles in the boundary layer and the wake regions. It is shown how the method can be coupled with an unstructured Navier-Stokes solver to produce a solution adaptive mesh generation procedure for viscous flows.
    Keywords: AERODYNAMICS
    Type: NASA-CR-181699 , ICASE-88-47 , NAS 1.26:181699
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    Space shuttle phase B wind tunnel model and test information. Volume 1: Booster configuration (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Archived wind tunnel test data are available for flyback booster or other alternative recoverable configurations as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle. Considerable wind tunnel data was acquired by the competing contractors and the NASA Centers for an extensive variety of configurations with an array of wing and body planforms. All contractor and NASA wind tunnel test data acquired in the Phase B development have been compiled into a database and are available for application to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Database is structured by vehicle component and configuration type. Basic components include the booster, the orbiter and the launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retroglide and twin body. Orbiter configuration types include straight and delta wings, lifting body, drop tanks, and double delta wings. Launch configurations include booster and orbiter components in various stacked and tandem combinations. This is Volume 1 (Part 1) of the report -- Booster Configuration.
    Keywords: AERODYNAMICS
    Type: NASA-CR-178414-VOL-1-PT-1 , NAS 1.26:178414-VOL-1-PT-1 , DMS-DB-02-VOL-1-PT-1
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  • 174
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    Investigation of a SiC/Ti-24Al-11Nb composite (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A summary of ongoing research on the characterization of a continuous fiber reinforced SiC/Ti-24Al-11Nb (at percent) composite is presented. The powder metallurgy fabrication technique is described as are the nondestructive evaluation results of the as-fabricated composite plates. Tensile properties of the SiC fiber, the matrix material, and the 0-deg SiC/Ti-24Al-11Nb composite (fibers oriented unidirectionally, parallel to the loading axis) from room temperature to 1100 C are presented and discussed with regard to the resultant fractography. The as-fabricated fiber-matrix interface has been examined by scanning transmission electron microscopy and the compounds present in the reaction zone have been identified. Fiber-matrix interaction and stability of the matrix near the fiber is characterized at 815, 985, and 1200 C from 1 to 500 hr. Measurements of the fiber-matrix reaction, the loss of C-rich coating from the surface of the SiC fiber, and the growth of the Beta depleted zone in the matrix adjacent to the fiber are presented. These data and the difference in coefficient of thermal expansion between the fiber and the matrix are discussed in terms of their likely effects on mechanical properties.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-TM-100956 , E-4253 , NAS 1.15:100956
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  • 175
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    Finite Element Composite Analysis Program (FECAP) for a microcomputer (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A special purpose finite element composite analysis program for analyzing composite material behavior with a microcomputer is described. The formulation assumes a state of generalized plane strain in a material consisting of two or more orthotropic phases. Loading can be mechanical and/or thermal. The theoretical background, computer implementation, and program users guide are described in detail. A sample program is solved showing the required user input and computer generated output.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-TM-100670 , NAS 1.15:100670
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  • 176
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    The high temperature creep deformation of Si3N4-6Y2O3-2Al2O3 (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The creep properties of silicon nitride containing 6 wt percent yttria and 2 wt percent alumina have been determined in the temperature range 1573 to 1673 K. The stress exponent, n, in the equation epsilon dot varies as sigma sup n, was determined to be 2.00 + or - 0.15 and the true activation energy was found to be 692 + or - 25 kJ/mol. Transmission electron microscopy studies showed that deformation occurred in the grain boundary glassy phase accompanied by microcrack formation and cavitation. The steady state creep results are consistent with a diffusion controlled creep mechanism involving nitrogen diffusion through the grain boundary glassy phase.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-CR-183204 , NAS 1.26:183204
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    Accurate solutions, parameter studies and comparisons for the Euler and potential flow equations (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Parameter studies are conducted using the Euler and potential flow equation models for unsteady and steady flows in both two and three dimensions. The Euler code is an implicit, upwind, finite volume code which uses the Van Leer method of flux-vector-splitting which has been recently extended for use on dynamic meshes and maintain all the properties of the original splitting. The potential flow code is an implicit, finite difference method for solving the transonic small disturbance equations and incorporates both entropy and vorticity corrections into the solution procedures thereby extending its applicability into regimes where shock strength normally precludes its use. Parameter studies resulting in benchmark type calculations include the effects of spatial and temporal refinement, spatial order of accuracy, far field boundary conditions for steady flow, frequency of oscillation, and the use of subiterations at each time step to reduce linearization and factorization errors. Comparisons between Euler and potential flows results are made as well as with experimental data where available.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100664 , NAS 1.15:100664
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  • 178
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    Space shuttle phase B wind tunnel model and test information. Volume 2: Orbiter configuration (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Archived wind tunnel test data are available for flyback booster or other alternate recoverable configurations as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle. Considerable wind tunnel data was acquired by the competing contractors and the NASA centers for an extensive variety of configurations with an array of wing and body planforms. All contractor and NASA wind tunnel test data acquiredin the Phase B development have been compiled into a database and are available for applying to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Database is structured by vehicle component and configuration type. Basic components include the booster, the orbiter, and the launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retroglide, and twin body. Orbiter configuration types include straight and delta wings, lifting body, drop tanks, and double delta wings. Launch configration types include booster and orbiter components in various stacked and tandom combinations. The digital database consists of 220 files of data containing basic tunnel recorded data.
    Keywords: AERODYNAMICS
    Type: NASA-CR-178415-VOL-2-PT-1 , NAS 1.26:178415-VOL-2-PT-1 , DMS-DB-02-VOL-2-PT-1
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  • 179
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    Probabilistic structural analysis to quantify uncertainties associated with turbopump blades (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A probabilistic study of turbopump blades has been in progress at NASA Lewis Research Center for over the last two years. The objectives of this study are to evaluate the effects of uncertainties in geometry and material properties on the structural response of the turbopump blades to evaluate the tolerance limits on the design. A methodology based on probabilistic approach was developed to quantify the effects of the random uncertainties. The results indicate that only the variations in geometry have significant effects.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-TM-100278 , E-3919 , NAS 1.15:100278
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  • 180
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    As-received microstructure of a SiC/Ti-15-3 composite (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A silicon carbide fiber reinforced titanium (Ti-15V-3Cr-3Sn-3Al) composite is metallographically examined. Several methods for examining composite materials are investigated and documented. Polishing techniques for this material are described. An interference layering method is developed to reveal the structure of the fiber, the reaction zone, and various phases within the matrix. Microprobe and transmission electron microscope (TEM) analyses are performed on the fiber/matrix interface. A detailed description of the fiber distribution as well as the microstructure of the fiber and matrix are presented.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-TM-100938 , E-4218 , NAS 1.15:100938
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  • 181
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    Numerical computation of space shuttle orbiter flow field (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A new parabolized Navier-Stokes (PNS) code has been developed to compute the hypersonic, viscous chemically reacting flow fields around 3-D bodies. The flow medium is assumed to be a multicomponent mixture of thermally perfect but calorically imperfect gases. The new PNS code solves the gas dynamic and species conservation equations in a coupled manner using a noniterative, implicit, approximately factored, finite difference algorithm. The space-marching method is made well-posed by special treatment of the streamwise pressure gradient term. The code has been used to compute hypersonic laminar flow of chemically reacting air over cones at angle of attack. The results of the computations are compared with the results of reacting boundary-layer computations and show excellent agreement.
    Keywords: AERODYNAMICS
    Type: NASA-CR-183193 , NAS 1.26:183193 , ISU-ERI-AMES-89403 , CFD-19
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  • 182
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    Program user's manual for an unsteady helicopter rotor-fuselage aerodynamic analysis (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The Rotor-Fuselage Analysis is a method of calculating the aerodynamic reaction between a helicopter rotor and fuselage. This manual describes the structure and operation of the computer programs that make up the Rotor-Fuselage Analysis, programs which prepare the input and programs which display the output.
    Keywords: AERODYNAMICS
    Type: NASA-CR-181701 , NAS 1.26:181701 , UTRC/R88-956977
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  • 183
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    Three-dimensional zonal grids about arbitrary shapes by Poisson's equation (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A method for generating 3-D finite difference grids about or within arbitrary shapes is presented. The 3-D Poisson equations are solved numerically, with values for the inhomogeneous terms found automatically by the algorithm. Those inhomogeneous terms have the effect near boundaries of reducing cell skewness and imposing arbitrary cell height. The method allows the region of interest to be divided into zones (blocks), allowing the method to be applicable to almost any physical domain. A FORTRAN program called 3DGRAPE has been written to implement the algorithm. Lastly, a method for redistributing grid points along lines normal to boundaries will be described.
    Keywords: AERODYNAMICS
    Type: NASA-TM-101018 , A-88258 , NAS 1.15:101018
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  • 184
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    An improved version of NCOREL: A computer program for 3-D nonlinear supersonic potential flow computations (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A computer code called NCOREL (for Nonconical Relaxation) has been developed to solve for supersonic full potential flows over complex geometries. The method first solves for the conical at the apex and then marches downstream in a spherical coordinate system. Implicit relaxation techniques are used to numerically solve the full potential equation at each subsequent crossflow plane. Many improvements have been made to the original code including more reliable numerics for computing wing-body flows with multiple embedded shocks, inlet flow through simulation, wake model and entropy corrections. Line relaxation or approximate factorization schemes are optionally available. Improved internal grid generation using analytic conformal mappings, supported by a simple geometric Harris wave drag input that was originally developed for panel methods and internal geometry package are some of the new features.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4165 , NAS 1.26:4165 , RE-744
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  • 185
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    Euler equation computations for the flow over a hovering helicopter rotor (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A numerical solution technique is developed for computing the flow field around an isolated helicopter rotor in hover. The flow is governed by the compressible Euler equations which are integrated using a finite volume approach. The Euler equations are coupled to a free wake model of the rotary wing vortical wake. This wake model is incorporated into the finite volume solver using a prescribed flow, or perturbation, technique which eliminates the numerical diffusion of vorticity due to the artificial viscosity of the scheme. The work is divided into three major parts: (1) comparisons of Euler solutions to experimental data for the flow around isolated wings show good agreement with the surface pressures, but poor agreement with the vortical wake structure; (2) the perturbation method is developed and used to compute the interaction of a streamwise vortex with a semispan wing. The rapid diffusion of the vortex when only the basic Euler solver is used is illustrated, and excellent agreement with experimental section lift coefficients is demonstrated when using the perturbation approach; and (3) the free wake solution technique is described and the coupling of the wake to the Euler solver for an isolated rotor is presented. Comparisons with experimental blade load data for several cases show good agreement, with discrepancies largely attributable to the neglect of viscous effects. The computed wake geometries agree less well with experiment, the primary difference being that too rapid a wake contraction is predicted for all the cases.
    Keywords: AERODYNAMICS
    Type: NASA-CR-177493 , NAS 1.26:177493
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  • 186
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    Dual adaptive control: Design principles and applications (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The design of an actively adaptive dual controller based on an approximation of the stochastic dynamic programming equation for a multi-step horizon is presented. A dual controller that can enhance identification of the system while controlling it at the same time is derived for multi-dimensional problems. This dual controller uses sensitivity functions of the expected future cost with respect to the parameter uncertainties. A passively adaptive cautious controller and the actively adaptive dual controller are examined. In many instances, the cautious controller is seen to turn off while the latter avoids the turn-off of the control and the slow convergence of the parameter estimates, characteristic of the cautious controller. The algorithms have been applied to a multi-variable static model which represents a simplified linear version of the relationship between the vibration output and the higher harmonic control input for a helicopter. Monte Carlo comparisons based on parametric and nonparametric statistical analysis indicate the superiority of the dual controller over the baseline controller.
    Keywords: AERODYNAMICS
    Type: NASA-CR-177485 , NAS 1.26:177485
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  • 187
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    Investigation of oscillating cascade aerodynamics by an experimental influence coefficient technique (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Fundamental experiments are performed in the NASA Lewis Transonic Oscillating Cascade Facility to investigate the torsion mode unsteady aerodynamics of a biconvex airfoil cascade at realistic values of the reduced frequency for all interblade phase angles at a specified mean flow condition. In particular, an unsteady aerodynamic influence coefficient technique is developed and utilized in which only one airfoil in the cascade is oscillated at a time and the resulting airfoil surface unsteady pressure distribution measured on one dynamically instrumented airfoil. The unsteady aerodynamics of an equivalent cascade with all airfoils oscillating at a specified interblade phase angle are then determined through a vector summation of these data. These influence coefficient determined oscillation cascade data are correlated with data obtained in this cascade with all airfoils oscillating at several interblade phase angle values. The influence coefficients are then utilized to determine the unsteady aerodynamics of the cascade for all interblade phase angles, with these unique data subsequently correlated with predictions from a linearized unsteady cascade model.
    Keywords: AERODYNAMICS
    Type: NASA-TM-101313 , E-4308 , NAS 1.15:101313 , AIAA PAPER 88-2815
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  • 188
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    Method to predict external store carriage characteristics at transonic speeds (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Development of a computational method for prediction of external store carriage characteristics at transonic speeds is described. The geometric flexibility required for treatment of pylon-mounted stores is achieved by computing finite difference solutions on a five-level embedded grid arrangement. A completely automated grid generation procedure facilitates applications. Store modeling capability consists of bodies of revolution with multiple fore and aft fins. A body-conforming grid improves the accuracy of the computed store body flow field. A nonlinear relaxation scheme developed specifically for modified transonic small disturbance flow equations enhances the method's numerical stability and accuracy. As a result, treatment of lower aspect ratio, more highly swept and tapered wings is possible. A limited supersonic freestream capability is also provided. Pressure, load distribution, and force/moment correlations show good agreement with experimental data for several test cases. A detailed computer program description for the Transonic Store Carriage Loads Prediction (TSCLP) Code is included.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4170 , NAS 1.26:4170
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  • 189
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    VSAERO analysis of tip planforms for the free-tip rotor (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The results of a numerical analysis of two interacting lifting surfaces separated in the spanwise direction by a narrow gap are presented. The configuration consists of a semispan wing with the last 32 percent of the span structurally separated from the inboard section. The angle of attack of the outboard section is set independently from that of the inboard section. In the present study, the three-dimensional panel code VSAERO is used to perform the analysis. Computed values of tip surface lift and pitching moment coefficients are correlated with experimental data to determine the proper approach to model the gap region between the surfaces. Pitching moment data for various tip planforms are also presented to show how the variation of tip pitching moment with angle of attack may be increased easily in incompressible flow. Calculated three-dimensional characteristics in compressible flow at Mach numbers of 0.5 and 0.7 are presented for new tip planform designs. An analysis of sectional aerodynamic center shift as a function of Mach number is also included for a representative tip planform. It is also shown that the induced drag of the tip surface is reduced for negative incidence angles relative to the inboard section. The results indicate that this local drag reduction overcomes the associated increase in wing induced drag at high wing lift coefficients.
    Keywords: AERODYNAMICS
    Type: NASA-CR-177487 , NAS 1.26:177487
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  • 190
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    Characterization of elastic-plastic properties of AS4/APC-2 thermoplastic composite (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Elastic and inelastic properties of AS4/APC-2 composites were characterized with respect to temperature variation by using a one-parameter orthotropic plasticity model and a one parameter failure criterion. Simple uniaxial off-axis tension tests were performed on coupon specimens of unidirectional AS4/APC-2 thermoplastic composite at various temperatures. To avoid the complication caused by the extension-shear coupling effect in off-axis testing, new tabs were designed and used on the test specimens. The experimental results showed that the nonlinear behavior of constitutive relations and the failure strengths can be characterized quite well using the one parameter plasticity model and the failure criterion, respectively.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-CR-183145 , NAS 1.26:183145 , CML-88-3
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  • 191
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    Thermo-viscoelastic analysis of composite materials, volume 1 (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Advanced composite materials, especially graphite/epoxy, are being applied to aircraft structures in order to improve performance and save weight. An important consideration in composite design is the residual strength of a structure containing holes, delaminations, or interlaminar damage when subjected to compressive loads. Recent studies have revealed the importance of viscoelastic effects in polymer-based composites. The viscoelastic effect is particularly significant at elevated temperature/moisture conditions since the matrix material is strongly affected by the environment. The solution of viscoelastic problems in composites was limited to special cases which can be solved by classical lamination theory. A finite element procedure is presented for calculating time-dependent stresses and strains in composite structures with general configurations and complicated boundary conditions. Using this procedure the in-plane and interlaminar stress distributions and histories in notched and unnotched composites were obtained for mechanical and thermal loads. Both two-dimensional and three-dimensional viscoelastic problems are analyzed. The effects of layup orientation and load spectrum on creep response and stress relaxation were also studied.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-CR-183048-VOL-1 , NAS 1.26:183048-VOL-1
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  • 192
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    A computational analysis of under-expanded jets in the hypersonic regime (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Underexpanded axisymmetric jets are studied numerically using a full Navier-Stokes solver. Emphasis has been given to supersonic and hypersonic jets in supersonic and hypersonic ambient flows, a phenomenon previously overlooked. It is demonstrated that the shear layers and shock patterns in a jet plume can be captured without complicated viscous/inviscid and subsonic/supersonic coupling schemes. In addition, a supersonic pressure relief effect has been identified for underexpanded jets in supersonic ambient flows. While it is well known that an underexpanded jet in a quiescent ambience (or subsonic ambience) contains multiple shock cells, the present study shows that because of the supersonic pressure relief effect, an underexpanded jet in a supersonic or hypersonic ambience contains only one major shock cell.
    Keywords: AERODYNAMICS
    Type: NASA-TM-101319 , E-4317 , NAS 1.15:101319
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  • 193
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    Space shuttle phase B wind tunnel model and test information. Volume 2: Orbiter configuration (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Archived wind tunnel test data are available for flyback booster or other alternative recoverable configurations as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle. Considerable wind tunnel data was acquired by the competing contractors and the NASA centers for an extensive variety of configurations with an array of wing and body planforms. All contractor and NASA wind tunnel test data acquired in the Phase B development have been compiled into a data base and are available for applying to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Data Base is structured by vehicle component and configuration type. Basic components include the booster, the orbiter, and the launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retro-glide and twin body. Orbiter configuration types include straight and delta wings, lifting body, drop tanks, and double delta wings. Launch configuration types include booster and orbiter components in various stacked and tandem combinations.
    Keywords: AERODYNAMICS
    Type: NASA-CR-178415-VOL-2-PT-2 , NAS 1.26:178415-VOL-2-PT-2 , DMS-DB-02-VOL-2
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  • 194
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    A micromechanics model of the stiffness and strength of laminates with fiber waviness (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A mathematical model based on the Euler-Bernoulli beam theory is proposed for predicting the effective Young's moduli of piecewise isotropic composite laminates with local ply curvatures in the main load-carrying layers. Strains in corrugated layers, In-Phase layers, Out-of-Phase layers are predicted for various geometries and material configurations by assuming matrix layers as elastic foundations of different spring constants. The effective Young's moduli measured from corrugated aluminum specimens and aluminum/epoxy specimens with In-Phase and Out-of-Phase wavy patterns coincide very well with the model predictions. Moire fringe analysis of an In-Phase specimen and an Out-of-Phase specimen are also presented confirming the main assumption of the model related to the elastic constraint due to the matrix layers. The present model is also compared with the experimental results and other models, including the micro-buckling models, published in the literature. The results of the present study show that even a very small scale local ply curvature produces a noticeable effect on the mechanical constitutive behavior of a laminated composite.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-CR-181670 , NAS 1.26:181670
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  • 195
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    The determination of the topological structure of skin friction lines on a rectangular wing-body combination (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A short tutorial in the application of topological ideas to the intepretation of oil flow patterns is presented. Topological concepts such as critical points, phase portraits, topological stability, and indexing are discussed. These concepts are used in an ordered procedure to construct phase portraits of skin friction lines with oil flow patterns for a wing-body combination and two angles of attack. The relationship between the skin friction phase portrait and planar cuts of the velocity field is also discussed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4168 , NAS 1.26:4168
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  • 196
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    Vortex methods for separated flows (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The numerical solution of the Euler or Navier-Stokes equations by Lagrangian vortex methods is discussed. The mathematical background is presented in an elementary fashion and includes the relationship with traditional point-vortex studies, the convergence to smooth solutions of the Euler equations, and the essential differences between two- and three-dimensional cases. The difficulties in extending the method to viscous or compressible flows are explained. The overlap with the excellent review articles available is kept to a minimum and more emphasis is placed on the area of expertise, namely two-dimensional flows around bluff bodies. When solid walls are present, complete mathematical models are not available and a more heuristic attitude must be adopted. The imposition of inviscid and viscous boundary conditions without conformal mappings or image vortices and the creation of vorticity along solid walls are examined in detail. Methods for boundary-layer treatment and the question of the Kutta condition are discussed. Practical aspects and tips helpful in creating a method that really works are explained. The topics include the robustness of the method and the assessment of accuracy, vortex-core profiles, timemarching schemes, numerical dissipation, and efficient programming. Calculations of flows past streamlined or bluff bodies are used as examples when appropriate.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100068 , A-88097 , NAS 1.15:100068
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  • 197
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    A diagonally inverted LU implicit multigrid scheme (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A new Diagonally Inverted LU Implicit scheme is developed within the framework of the multigrid method for the 3-D unsteady Euler equations. The matrix systems that are to be inverted in the LU scheme are treated by local diagonalizing transformations that decouple them into systems of scalar equations. Unlike the Diagonalized ADI method, the time accuracy of the LU scheme is not reduced since the diagonalization procedure does not destroy time conservation. Even more importantly, this diagonalization significantly reduces the computational effort required to solve the LU approximation and therefore transforms it into a more efficient method of numerically solving the 3-D Euler equations.
    Keywords: AERODYNAMICS
    Type: NASA-TM-100911 , E-4163 , NAS 1.15:100911 , AIAA PAPER 88-3567
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  • 198
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    High speed inlet calculations with real gas effects (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A 2-D steady-state Navier-Stokes solver has been upgraded to include the effects of frozen and equilibrium air chemistry for applications to high speed flight vehicles. To provide a computationally economical first order approximation to the high temperature physics, variable thermodynamic data is used for the chemically frozen mode to allow for a variation with temperature of the air specific heats and enthalpy. For calculations involving air in chemical equilibrium, a specially modified version of the NASA Lewis Chemical Equilibrium Code, CEC, is used to compute the chemical composition and resultant thermochemical properties. The upgraded solver is demonstrated by comparing results from calorically perfect (C sub p=constant), thermally perfect (frozen) and equilibrium air calculations for a variety of geometries, and flight Mach numbers.
    Keywords: AERODYNAMICS
    Type: NASA-CR-182167 , E-4287 , NAS 1.26:182167 , AIAA PAPER 88-3076
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  • 199
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    Numerical study of delta wing leading edge blowing (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Spanwise and tangential leading edge blowing as a means of controlling the position and strength of the leading edge vortices are studied by numerical solution of the three-dimensional Navier-Stokes equations. The leading edge jet is simulated by defining a permeable boundary, corresponding to the jet slot, where suitable boundary conditions are implemented. Numerical results are shown to compare favorably with experimental measurements. It is found that the use of spanwise leading edge blowing at moderate angle of attack magnifies the size and strength of the leading edge vortices, and moves the vortex cores outboard and upward. The increase in lift primarily comes from the greater nonlinear vortex lift. However, spanwise blowing causes earlier vortex breakdown, thus decreasing the stall angle. The effects of tangential blowing at low to moderate angles of attack tend to reduce the pressure peaks associated with leading edge vortices and to increase the suction peak around the leading edge, so that the integrated value of the surface pressure remains about the same. Tangential leading edge blowing in post-stall conditions is shown to re-establish vortical flow and delay vortex bursting, thus increasing C sub L sub max and stall angle.
    Keywords: AERODYNAMICS
    Type: NASA-CR-183101 , NAS 1.26:183101 , JIAA-TR-86
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  • 200
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    Loading rate sensitivity of open hole composites in compression (1988)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The results are reported of an experimental study on the compressive, time-dependent behavior of graphite fiber reinforced polymer composite laminates with open holes. The effect of loading rate on compressive strength was determined for six material systems ranging from brittle epoxies to thermoplastics at both 75 F and 220 F. Specimens were loaded to failure using different loading rates. The slope of the strength versus elapsed time-to-failure curve was used to rank the materials' loading rate sensitivity. All of the materials had greater strength at 75 F than at 220 F. All the materials showed loading rate effects in the form of reduced failure strength for longer elapsed-time-to-failure. Loading rate sensitivity was less at 220 F than the same material at 70 F. However, C12000/ULTEM and IM7/8551-7 were more sensitive to loading rate than the other materials at 220 F. AS4/APC2 laminates with 24, 32, and 48 plies and 1/16 and 1/4 inch diameter holes were tested. The sensitivity to loading rate was less for either increasing number of plies or larger hole size. The failure of the specimens made from brittle resins was accompanied by extensive delaminations while the failure of the roughened systems was predominantly by shear crippling. Fewer delamination failures were observed at the higher temperature.
    Keywords: COMPOSITE MATERIALS
    Type: NASA-TM-100634 , NAS 1.15:100634 , AVSCOM-TM-88-B-012
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