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  • AERODYNAMICS  (479)
  • 2015-2019
  • 1980-1984  (479)
  • 1925-1929
  • 1983  (479)
  • 101
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    An experimental investigation of multiple shock wave/turbulent boundary layer interactions in a circular duct (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Detailed pitot, static and wall pressure measurements have been obtained for multiple shock wave/turbulent boundary layer interactions in a circular duct at a free-stream Mach number of 1.49 and at a unit Reynolds number of 4.90 x 10 to the 6th per meter. The details of the flow field show the formation of a series of normal shock waves with successively decreasing strength and with decreasing distance between the successive shock waves. The overall pressure recovery is much lower than the single normal shock pressure recovery at the same free-stream Mach number. A one-dimensional flow model based on the boundary layer displacement buildup is postulated to explain the formation of a series of normal shock waves.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1744
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  • 102
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    Measurements of compressible secondary flow in a circular S-duct (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: This paper presents the results of an experimental study of secondary flow in a circular cross section 30 deg - 30 deg S-duct with entrance Mach number of 0.6. Local flow velocity vectors have been measured along the length of the duct at six stations. These measurements have been made using a five-port cone probe. Static and total pressure profiles in the transverse planes are obtained from the cone probe measurements. Wall static pressure measurements along three azimuth angles of 0 deg, 90 deg, and 180 deg along the duct are also made. Contour plots presenting the three dimensional velocity field as well as the total- and static-pressure fields are obtained. Surface oil flow visualization technique has been used to provide details of the flow on the S-duct boundaries. The experimental observations have been compared with typical computational results.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1739
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  • 103
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    Contouring tunnel walls to achieve free-air flow over a transonic swept wing (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The effects of wind-tunnel walls on the flow over a swept wing were greatly reduced by wall contouring. Significant reductions in spanwise pressure gradients were achieved by shaping all of the walls to conform to the streamlines over the model in free air. Surface pressure and oil-flow data were used to evaluate the effects of Mach and Reynolds numbers on the design. Comparisons of these data with inviscid calculations indicate that free-air flow is established at a Mach number of 0.74 and at Reynolds numbers above 4.7 million.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1725
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  • 104
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    Calculation of afterbody flows with a composite velocity formulation (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A recently developed technique for numerical solution of the Navier-Stokes equations for subsonic, laminar flows is investigated. It is extended here to allow for the computation of transonic and turbulent flows. The basic approach involves a multiplicative composite of the appropriate velocity representations for the inviscid and viscous flow regions. The resulting equations are structured so that far from the surface of the body the momentum equations lead to the Bernoulli equation for the pressure, while the continuity equation reduces to the familiar potential equation. Close to the body surface, the governing equations and solution techniques are characteristic of those describing interacting boundary layers. The velocity components are computed with a coupled strongly implicity procedure. For transonic flows the artificial compressibility method is used to treat supersonic regions. Calculations are made for both laminar and turbulent flows over axisymmetric afterbody configurations. Present results compare favorably with other numerical solutions and/or experimental data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1736
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  • 105
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    Energetics and optimum motion of oscillating lifting surfaces (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Low-frequency, unsteady, lifting-line theory is used to characterize the energetics and optimum motion of an unswept rigid wing oscillating harmonically in an inviscid, incompressible flow. The energetics calculations account for the leading edge suction force, the power absorbed in the wing oscillations, and the energy loss rate produced by vortex shedding. Optimization is achieved by minimizing the average energy loss rate in relation to a given thrust, and a unique solution is found in the three dimensional case for low, reduced frequencies. The two-dimensional solution is nonunique, a condition which is examined in terms of the normal modes of the energy loss rate matrix. An invisible mode with a hydrodynamic efficiency of 100 pct is obtained in the two-dimensional case, causing the nonuniqueness of the solution by yielding no fixed positive thrust through perfect unsteady feathering.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1710
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  • 106
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    Navier-Stokes calculations for the vortex wake of a rotor in hover (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An efficient finite-difference scheme for the solution of the incompressible Navier-Stokes equation is used to study the vortex wake of a rotor in hover. The solution procedure uses a vorticity-stream function formulation and incorporates an asymptotic far-field boundary condition enabling the size of the computational domain to be reduced in comparison to other methods. The results from the present method are compared with experimental data obtained by smoke flow visualization and hot-wire measurements for several rotor blade configurations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1676
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  • 107
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    An experimental study of surface curvature effects on a supersonic turbulent boundary layer (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The mean flow results of an experimental study of compressible turbulent boundary layers in an adverse pressure gradient with and without surface curvature effects are presented. The test was conducted in an axisymmetric flow facility. The upstream Reynolds number based on boundary layer momentum thickness was 5884 and the boundary layer thickness was 0.90 cm. The curvature effects were examined by studying two flows with essentially identical adverse pressure gradients. One flow was along a concave compression surface test section, while the other was along a straight-walled test section. Mean flow measurements included wall static pressure distributions, wall temperatures, pitot pressure profiles and total temperature profiles. The mean flow results indicated that the surface curvature resulted in a definite increase of turbulent mixing in the boundary layer.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1672
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  • 108
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    Experimental studies of the boundary layer on an airfoil at low Reynolds numbers (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An experimental wind tunnel investigation was carried out to study the effect of laminar separation bubbles on a NACA 66(3)-018 airfoil for Reynolds numbers less than 4.0 x 10 to the 5th. Leading edge laminar separation bubbles formed for angles of attack of approximately 7 to 12 deg. To study the leading edge separation bubble more closely, hotwire anemometer measurements were made in the airfoil a Reynolds number of 8.0 x 10 to the 4th. Velocity and turbulence intensity profiles were obtained and boundary layer parameters were calculated. Frequency spectra were also calculated at key points in the airfoil boundary layer for this case. Correlation of the anemometry data with static pressure distributions, and flow visualization data provided insight into laminar separation bubble behavior at low Reynolds numbers.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1671
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  • 109
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    Calculation of viscous supersonic flows over finned bodies (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The parabolized Navier-Stokes (PNS) equations are used to calculate the viscous, supersonic flow fields about a six-finned projectile and a generic four-finned missile at angles of attack. Since current computer speeds and storage preclude a fully three-dimensional calculation using the unsteady, Reynolds-averaged, Navier-Stokes equations, the applicability of the PNS equations to the above flow fields is of considerable interest. Two important aspects of the calculation are grid generation and the type of smoothing used to prevent nonphysical solutions. This paper includes a description of the grid-generation process. Results in the form of density contours and velocity vector plots are presented for the two configurations. The applicability of the PNS equations to the complicated flow fields considered is successfully demonstrated.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1667
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  • 110
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    Topology of vortices in conical flow (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The possibility of the existence of a viscous vortex in a conical flow is discussed. Existence of a viscous vortex is shown to be consistent with a flow-field model where only the velocity components and the total enthalpy behave conically. Existence of a viscous vortex is not consistent with the fully conical, flow-field model; however, far from the nose tip, the physical flow quantities tend to behave asymptotically as a full conical flow. Such a viscous vortex, when it exists, must spiral inwards to the focal point. It can be used as a reasonable model to start numerical solutions of various kinds.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1664
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  • 111
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    Hypersonic flows over biconics using a variable-effective-gamma, Parabolized-Navier-Stokes code (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Hypersonic flows over straight and bent biconics are calculated for a range of freestream conditions in which the gas behind the shock is treated as either perfect or real. The Parabolized-Navier-Stokes (PNS) equations form the basis of the approximation scheme. Good comparisons with experimental data for pressures, forces and moments, heat transfer, and oil-flow patterns serve to validate the perfect-gas version of the code. Circumferential velocity vector plots further aid in the interpretation of leeside oil-flow patterns. A variable-effective-gamma (VEG) option is implemented for the real-gas calculations. Gamma, now defined as the ratio of enthalpy to internal energy, is determined from a locally valid linear relation in enthalpy and pressure at every mesh point which in turn is calculated from a benchmark equilibrium code. The VEG option is easily incorporated into a host PNS code because it uses the underlying perfect-gas structure of the code. Comparisons to experimental data for heat transfer and shock shape in high enthalpy air have been obtained using the VEG option.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1666
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  • 112
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    Three-dimensional transonic nacelle/inlet flowfield computations using an efficient approximate factorization algorithm (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A highly efficient computer analysis has been developed for predicting transonic nacelle/inlet flowfields. This algorithm can compute the three-dimensional transonic flowfield about axisymmetric or asymmetric nacelle/inlet configurations at zero or nonzero incidence. The flowfield is determined by solving the full-potential equation in conservative form on a body-fitted curvilinear computational mesh. The difference equations are solved using the AF2 approximate factorization scheme. The effects of boundary layer viscous entrainment are approximated in the inviscid algorithm by applying a surface transpiration velocity which is determined from the calculated boundary layer growth. Computed results and correlations with existing methods and experiment are presented to illustrate application of the analysis.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1417
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  • 113
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    Three-dimensional compressible viscous analysis of mixer nozzles (1983)
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    Publication Date: 2019-06-28
    Description: An analysis-based design procedure for compound-mixer exhaust nozzles is presented and compared to test data. The design approach is based on two numerical solutions to the 3-D viscous compressible Navier-Stokes equations: an equation splitting technique used for the analysis of the core and bypass flow, and a parabolic marching scheme used in the analysis of the mixing duct. The selection of the analytical methods through test data comparisons and their coupling into an integrated design system are discussed. NASA test data is used to demonstrate the validity of the computations from the exhaust system rating station, upstream of the mixer lobe, to the nozzle throat. An estimate is made of the savings in development time and cost utilizing the new procedure.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1401
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  • 114
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    An integral analysis of transonic normal shock wave/turbulent boundary layer interactions in internal flow (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An approximate integral viscous-inviscid interaction method is presented for calculating the development of a turbulent boundary layer subjected to a normal shock wave induced adverse pressure gradient in an internal axisymmetric flow. The inflow conditions and the downstream pressure are provided for the computation. In the supersonic region of shock pressure rise, the Prandtl-Meyer function is used to couple the viscous and inviscid flows. An analytical model for the coupling process is postulated and appropriate equations are defined. Downstream of the sonic point, one-dimensional inviscid flow is assumed for coupling with the viscous flow. The turbulent boundary layer is calculated using Green's integral lag-entrainment method. Comparisons of the solutions with the experimental data are made for interactions which are unseparated, near separation and separated. For comparison purposes, solutions to the time-dependent, mass-averaged, Navier-Stokes equations incorporating a two-equation, Wilcox-Rubesin turbulence model are also shown. The computed results from the integral method show good agreement with experimental data for unseparated interactions and reasonable agreement with the trend of the viscous effects when the interaction becomes increasingly separated.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1402
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  • 115
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    Progress toward the analysis of supersonic inlet flows (1983)
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    Publication Date: 2019-06-28
    Description: A zonal flow analysis procedure was developed to predict the flow through the supersonic diffuser of an axisymmetric mixed compression inlet at angle-of-attack. In this analysis, the inlet flow is divided into three types of regions, each with different dominant flow phenomena. These are the inviscid supersonic core, boundary layer, and shock/boundary layer interaction flows. An appropriate analysis was selected or developed for the three-dimensional flow in each type of region. Procedures were developed to interface these analyses for the overall inlet flow analysis. This analysis was applied to an inlet operating at M = 2.58 at several angle-of-attack conditions. Comparisons are presented between computed and measured flow properties for the inlet and for the component analysis flows. Extensions of the present procedure to include the terminal shock and subsonic diffuser flows are recommended. Desirable experiments for evaluation of the inlet analysis procedure or the component analyses and to support improved modeling or extension of the inlet analysis are defined and recommended.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1371
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  • 116
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    Progress toward the analysis of complex propulsion installation flow phenomenon (1983)
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    Publication Date: 2019-06-28
    Description: A trend toward replacement of parametric model testing with parametric analysis for the design of aircraft is driven by the rapidly escalating cost of wind tunnel testing, the increasing availability of large fast computers, and powerful numerical flow algorithms. In connection with the complex flow phenomena characteristic of propulsion installations, it is now necessary to employ both parametric analysis and testing for design procedures. Powerful flow analysis techniques are available to predict local flow phenomena. However, the employment of these techniques is very expensive. It is, therefore, necessary to link these analyses with less powerful and less expensive procedures for an accurate analysis of propulsion installation flowfields. However, the interfacing and coupling processes needed are not available. The present investigation is concerned with progress made regarding the development of suitable linking methods. Attention is given to methods of analysis for predicting the flow around a nacelle coupled to a highly swept wing.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1367
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  • 117
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    A static investigation of yaw vectoring concepts on two-dimensional convergent-divergent nozzles (1983)
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    Publication Date: 2019-06-28
    Description: The flow-turning capability and nozzle internal performance of yaw-vectoring nozzle geometries were tested in the NASA Langley 16-ft Transonic wind tunnel. The concept was investigated as a means of enhancing fighter jet performance. Five two-dimensional convergent-divergent nozzles were equipped for yaw-vectoring and examined. The configurations included a translating left sidewall, left and right sidewall flaps downstream of the nozzle throat, left sidewall flaps or port located upstream of the nozzle throat, and a powered rudder. Trials were also run with 20 deg of pitch thrust vectoring added. The feasibility of providing yaw-thrust vectoring was demonstrated, with the largest yaw vector angles being obtained with sidewall flaps downstream of the nozzle primary throat. It was concluded that yaw vector designs that scoop or capture internal nozzle flow provide the largest yaw-vector capability, but decrease the thrust the most.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1288
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  • 118
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    STOL wind tunnel test results for a tactical supercruiser (1983)
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    Publication Date: 2019-06-28
    Description: Attention is given to the NASA Langley Research Center's testing of a 10.5 percent-scale supersonic cruiser (supercruiser) aircraft model in its V/STOL wind tunnel, in order to investigate the low speed aerodynamic characteristics of STOL enhancement devices. The STOL devices employed by the supercruiser configuration are high vector angle ramp nozzles, working in conjunction with a remote augmented lift system (RALS), in addition to a canard trim system. Also investigated were thrust reverser/ground plane interaction effects, for the evaluation of landing characteristics. It is noted that STOL approach thrust management requires the use of a partially reversing RALS nozzle which develops approximately 31 percent of main nozzle thrust, and that strong nose-up interactions during ground roll, with reverser operation, may limit dry power engine thrust for braking assistance to about 50 percent of maximum dry power.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1224
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  • 119
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    Method for calculating effects of a propfan on aircraft aerodynamics at subsonic speeds (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A subsonic-flow panel code has been modified to handle the effects of a propeller wake. The effects of the propeller were modelled by a system of ring vortices of constant strength. Principles based on the blade element theory and the momentum theory were used to evaluate the swirl velocity and the pressure increase, across the propeller. Theoretical calculations are compared to experimental results at a Mach number of 0.50. The discrepancies between the theory and the experimental results are analysed. Suggestions for improvements to enhance the accuracy of the theoretical prediction are indicated.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1216
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  • 120
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    Calculation of a simulated 3-D high speed inlet using the Navier-Stokes equations (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A hybrid numerical algorithm, developed to solve the full three-dimensional Navier-Stokes equations, is applied to the computation of the flowfield in a simulated three-dimensional high speed aircraft inlet at a Mach number of 2.5 and Reynolds number of 1.4 x 10 to the 7th based on inlet length. The numerical algorithm incorporates a coordinate transformation in order to handle general flow geometries, and utilizes the algebraic turbulent eddy viscosity model of Baldwin and Lomax. The hybrid algorithm has been vectorized on the CDC CYBER 203 computer using the SL/1 vector programming language developed at NASA Langley. The computed results are compared with experimental measurements of the ramp and cowl static pressures, and boundary layer pitot profiles. The results are also compared with a previous two-dimensional Navier-Stokes computation of the same configuration. The agreement with the experimental data is generally good; however, additional improvements in turbulence modeling are needed.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1165
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  • 121
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    Empennage/afterbody integration for single and twin-engine fighter aircraft (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An extensive experimental program to determine the effects of empennage surfaces on single and twin-engine afterbody/nozzle drag has been conducted by the Propulsion Aerodynamics Branch at the NASA Langley Research Center. Empennage interference drag was obtained by using experimental values of afterbody/nozzle drag and computed values of empennage drag. The effects of tail location, span, number (single versus twin), toe angle, cant angle, camber and root chord length are discussed. The magnitude of empennage interference drag on single and twin engine configurations is examined.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1126
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  • 122
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    Three-dimensional flow studies on a slotted transonic wind tunnel wall (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Previously cited in issue 06, p. 799, Accession no. A82-17855
    Keywords: AERODYNAMICS
    Type: (ISSN 0001-1452)
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  • 123
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    Improved relaxation schemes for transonic potential calculations (1983)
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    Publication Date: 2019-06-28
    Description: A block relaxation scheme, grouped in a red-black ordering, is applied to transonic airfoil calculations using body fitted coordinates. The scheme is simple and is easily vectorizable. Detailed comparisons with Approximate Factorization Method (AF2) are presented and it is shown that the improved relaxation scheme is competitive in all cases considered. Transonic results, of engineering accuracy, on an 0-type grid of 149 x 30 points, are ususally obtained within two hundred iterations (approximately 40 seconds on Cyber 175).
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-0372
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  • 124
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    Entrainment and mixing in thrust augmenting ejectors (1983)
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    Publication Date: 2019-06-28
    Description: An experimental investigation of two-dimensional thrust augmenting ejector flows has been conducted. Measurements of the shroud surface pressure distribution, mean velocity, turbulent intensities and Reynolds stresses were made in two shroud geometries at various primary nozzle pressure ratios. The effects of shroud geometry and primary nozzle pressure ratio on the shroud surface pressure distribution, mean flow field and turbulent field were determined. From these measurements the evolution of mixing within the shroud of the primary flow and entrained fluid was obtained. The relationship between the mean flow field, the turbulent field and the shroud surface pressure distribution is discussed.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-0172
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  • 125
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    Transonic flow calculations using the Euler equations (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Previously cited in issue 06, p. 796, Accession no. A82-17786
    Keywords: AERODYNAMICS
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  • 126
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    Aerodynamic theory for wing with side edge passing subsonically through a gust (1983)
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    Publication Date: 2019-06-28
    Description: An approximate solution for the unsteady loading near the square-shape tip of a wing passing through an oblique gust is obtained in closed form. The aerodynamic theory developed can be used to predict airloads felt by a helicopter blade experiencing a blade/vortex interaction for high blade tip speed and/or for small vertical blade/vortex separation. Under these conditions one can show that the blade's trailing edge has little influence on the character of the chordwise loading at all spanwise sections; thus, the chord may be allowed to extend to infinity in the downstream direction. Therefore, the model considered here is that of a quarter-infinite flat plate wing with side edge passing subsonically through an oblique gust.
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 21; June 198
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  • 127
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    Analysis of aerothermal loads on spherical dome protuberances (1983)
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    Publication Date: 2019-06-28
    Description: Hypersonic flow over spherical dome protuberances was investigated to determine increased pressure and heating loads to the surface. The configuration was mathematically modeled in a time-dependant three-dimensional analysis of the conservation of mass, momentum (Navier-Stokes), and energy equations. A boundary mapping techique was used to obtain a rectangular parallelepiped computational domain, and a MacCormack explicit time-split predictor-corrector finite difference algorithm was used to obtain solutions. Results show local pressures and heating rates for domes one-half, one, and two boundary layer thicknesses high were increased by factors on the order of 1.4, 2, and 6, respectively. However, because lee-side pressure and thermal loads were reduced the two lower height domes did not experience any net increase in total loads. The total loads on the higher dome were increased by twenty-five percent. Flow over the lower dome was everywhere attached while flow over the intermediate dome had small windward and leeside separations. The higher dome had an unsteady windward separation region and a large leeside separation region. Trailing vortices form on all domes with intensity increasing with dome height. Discussions of applying the results to a thermally bowed thermal protection system are presented.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1557
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  • 128
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    Aerodynamic characteristics of generalized bent biconic bodies for aero-assisted, orbital-transfer vehicles (1983)
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    Publication Date: 2019-06-28
    Description: A method was developed to generate the surface coordinates of body shapes suitable for aero-assisted, orbital-transfer vehicles (AOTVs) by extending bent biconic geometries. Lift, drag, and longitudinal moments were calculated for the bodies using Newtonian flow theory. These techniques were applied to symmetric and asymmetric aerobraking vehicles, and to an aeromaneuvering vehicle with high L/D. Results for aerobraking applications indicate that a 70-deg, fore half cone angle with spherically blunted nose, rounded edges, and a slight asymmetry would be appropriate. Moreover, results show that an aeromaneuvering vehicle with L/D greater than 2.0, and with sufficient stability, is feasible.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1512
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  • 129
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    Viscous real gas flowfields about three dimensional configurations (1983)
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    Publication Date: 2019-06-28
    Description: Laminar, real gas hypersonic flowfields over a three dimensional configuration are computed using an unsteady, factored implicit scheme. Local chemical and thermodynamic properties are evaluated by an equilibrium composition method. Transport properties are obtained from individual species properties and application of a mixture rule. Numerical solutions are presented for an ideal gas and equilibrium air for free-stream Mach numbers of 13 and 15 and at various angles of attack. The effect of real gas is to decrease the shock-layer thickness resulting from decreased shock-layer temperatures and corresponding increased density. The combined effects of viscosity and real gas are to increase the subsonic layer near the wall.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1511
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    Low L/D aerobrake test at Mach 10 (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An aerobraking Orbital transfer vehicle may be used to increase the Space Shuttle mission capability to and from high orbits. A Mach 10 wind-tunnel test was performed for a low lift-drag aerobrake, to define a preliminary aerothermal environment for this candidate concept. Test hardware simulated the ribs and stretched fabric of conceptual flight hardware. Pressures, paint-melting histories, and oilflow data were measured on the brake. Pressure and thermocouple heating rate data were measured on the payload. Brake peak heating is at the edge at all angles of attack, although the stagnation point is not outboard of 75 percent radius even at 20 degrees angle of attack. Brake ribs show slightly higher heating than flats, although pressures are essentially constant. Payload peak heating occurs near 12 degrees angle of attack, and is 30 percent of the sphere stagnation point heating (for a sphere of brake diameter). Payload pressure distributions follow the heating pattern. Reynolds number effects are small on the brake and large on the payload, for the range of test conditions: 0.4-1.0 million/foot.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1509
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  • 131
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    Aerothermal environment in control surface gaps in hypersonic flow - An overview (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Experimental and analytical research by the NASA Langley Research Center to develop an understanding of the fluid and thermal environment in control surface gaps such as the spanwise gap of the wing elevon and chordwise gap of a split elevon configuration typical of the Space Shuttle are summarized. Although the experimental and analytical studies were initiated too late to significantly impact the basic Shuttle design they do provide a fundamental understanding of the basic fluid/thermal environment in control surface gaps and help to establish a firm data base for future vehicle design.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1483
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  • 132
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    Computation of nonlinear supersonic potential flow over three-dimensional surfaces (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Previously cited in issue 12, p. 1850, Accession no. A82-27090
    Keywords: AERODYNAMICS
    Type: (ISSN 0021-8669)
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  • 133
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    Aerodynamic design using numerical optimization (1983)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: The procedure of using numerical optimization methods coupled with computational fluid dynamic (CFD) codes for the development of an aerodynamic design is examined. Several approaches that replace wind tunnel tests, develop pressure distributions and derive designs, or fulfill preset design criteria are presented. The method of Aerodynamic Design by Numerical Optimization (ADNO) is described and illustrated with examples.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85550 , NAS 1.15:85550 , CFDL-TR-83-2
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  • 134
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    Reynolds number effects on pressure loss and turbulence characteristics of four tube-bundle heat exchangers (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The aerodynamic characteristics of pressure loss and turbulence on four tube-bundle configurations representing heat-exchanger geometries with nominally the same heat capacity were measured as a function of Reynolds numbers from about 4000 to 400,000 based on tube hydraulic diameter. Two configurations had elliptical tubes, the other two had round tubes, and all four had plate fins. The elliptical-tube configurations had lower pressure loss and turbulence characteristics than the round-tube configurations over the entire Reynolds number range.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85807 , L-15721 , NAS 1.15:85807
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  • 135
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    Unsteady laminar boundary-layer calculations on oscillating configurations including backflow. Part 2: Airfoil in high-amplitude pitching motion. Dynamic stall (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A previously developed finite-difference procedure for calculating unsteady, incompressible, laminar boundary layers on an oscillating flat plate is applied to a wing section undergoing high-amplitude pitching oscillations about various mean incidences. To start the entire boundary-layer calculation, appropriate initial conditions and outer boundary conditions are specified, using a stagnation-point fixed frame of reference. The breakdown of the numerical calculation procedure in the x,t-domain is interpreted to coincide with unsteady separation. Details of the boundary-layer behavior in the vicinity of separation are investigated, and a close analogy between the present results and those for a three-dimensional steady separation is found.
    Keywords: AERODYNAMICS
    Type: NASA-TM-84319-PT-2 , A-9403 , NAS 1.15:84319-PT-2
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  • 136
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    Demonstration of the coast-down technique for determining train resistances (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Full-scale measurement or validation of the various factors of train running resistance is an essential step in decreasing train energy consumption. Such a measurement capability would enable railroads to evaluate the cost benefits of operational and train consistent configuration changes, and new vehicle and truck designs for decreasing aerodynamic drag and rolling resistance. A decrease in the rolling resistance affects more than just a decrease in energy consumption; it also will result in decreased mechanical wear, hence less wheel and rail maintenance and replacement costs. A demonstration of a simple coast-down technique (based on computer-reduction of distance history) was accomplished using specially configured trains on main line rail provided by the Atchison, Topeka and Sante Fe Railway Co. This demonstration test shows that this distance-history coast-down technique for trains is easy to execute in the field. The total running resistance history was accurately determined and subsequently separated into rolling resistance (mechanical friction) and aerodynamic drag.
    Keywords: AERODYNAMICS
    Type: NASA-CR-173468 , JPL-PUB-83-85 , NAS 1.26:173468
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  • 137
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    Experimental study of the separating confluent boundary-layer. Volume 2: Experimental data (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental low speed study of the separating confluent boundary layer on a NASA GAW-1 high lift airfoil is described. The airfoil was tested in a variety of high lift configurations comprised of leading edge slat and trailing edge flap combinations. The primary test instrumentation was a two dimensional laser velocimeter (LV) system operating in a backscatter mode. Surface pressures and corresponding LV derived boundary layer profiles are given in terms of velocity components, turbulence intensities and Reynolds shear stresses as characterizing confluent boundary layer behavior up to and beyond stall. LV derived profiles and associated boundary layer parameters and those obtained from more conventional instrumentation such as pitot static transverse, Preston tube measurements and hot-wire surveys are compared.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166018 , NAS 1.26:166018 , LG82ER0184
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  • 138
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    Numerical solution of three-dimensional unsteady transonic flow over wings including inviscid/viscous interactions (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A numerical procedure is presented for computing the unsteady transonic flow field about three dimensional swept wings undergoing general time dependent motion. The outer inviscid portion of the flow is assumed to be governed by the modified unsteady transonic small disturbance potential equation which is integrated in the time domain by means of an efficient alternating direction implicit approximate factorization algorithm. Gross dominant effects of the shock boundary layer interaction are accounted for by a simple empirically defined model. Viscous flow regions adjacent to the wing surface and in the trailing wake are described by a set of integral equations appropriate for compressible turbulent shear layers. The two dimensional boundary layer equations are applied quasi-statically stripwise across the span. Coupling with the outer inviscid flow is implemented through use of the displacement thickness concept within the limitations of small disturbance theory. Validity of the assumptions underlying the method is established by comparison with experimental data for the flow about a high aspect ratio transport wing having an advanced airfoil section.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166561 , NAS 1.26:166561
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  • 139
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    A wing concept for supersonic maneuvering (1983)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: A theoretical and experimental program in which a wing concept for supersonic maneuvering was developed and then demonstrated experimentally in a series of wind tunnel tests is described. For the typical fighter wing, the problem of obtaining efficient lift at supersonic maneuvering C sub 's occurs due to development of a strong crossflow shock, and boundary layer separation. A natural means of achieving efficient supersonic maneuvering is based on controlling the non-linear inviscid crossflow on the wing in a manner analogous to the supercritical aerodynamic methods developed for transonic speeds. The application of supercritical aerodynamics to supersonic speeds is carried out using Supercritical Conical Camber (SC3). This report provides an aerodynamic analysis of the effort, with emphasis on wing design using non-linear aerodynamics. The substantial experimental data base is described in three separate wind tunnel reports, while two of the computer programs used in the work are also described in a separate report. Based on the development program it appears that a controlled supercritical crossflow can be obtained reliably on fighter-type wing planforms, with an associated drag due to lift reduction of about 20% projected using this concept.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3763 , NAS 1.26:3763
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  • 140
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    Three-dimensional implicit lambda methods (1983)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: This paper derives the three dimensional lambda-formulation equations for a general orthogonal curvilinear coordinate system and provides various block-explicit and block-implicit methods for solving them, numerically. Three model problems, characterized by subsonic, supersonic and transonic flow conditions, are used to assess the reliability and compare the efficiency of the proposed methods.
    Keywords: AERODYNAMICS
    Type: NASA-CR-172264 , NAS 1.26:172264 , REPT-83-62
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  • 141
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    Multicomponent velocity measurement in a pison-cylinder configuration using laser velocimetry (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Because of its nonintrusive nature, Laser Doppler Velocimetry (LDV) has become a popular tool for velocity measurements in internal combustion engines. This work shows how one can use an on-axis measurement technique, in conjunction with the standard two channel LDV technique, to make simultaneous three-component measurements using a single focusing lens. Simultaneous measurement of two of these three components in a piston-cylinder configuration is demonstrated.
    Keywords: AERODYNAMICS
    Type: NASA-TM-83534 , E-1835 , NAS 1.15:83534
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  • 142
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    Aerodynamic characteristics, including effect of body shape, of a Mach 6 aircraft concept (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Longitudinal aerodynamic characteristics for a hydrogen-fueled hypersonic transport concept at Mach 6 are presented. The model components consist of four bodies with identical longitudinal area distributions but different cross-sectional shapes and widths, a wing, horizontal and vertical tails, and a set of wing-mounted nacelles simulated by slid bodies on the wing upper surface. Lift-drag ratios were found to be only sightly affected by fuselage planform width or cross sectional shape. Relative distribution of fuselage volume above and below the wing was found to have an effect on the lift-drag ratio, with a higher lift drag ratio produced by the higher wing position.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2235 , L-15675 , NAS 1.60:2235
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  • 143
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    Effect of thrust reverser operation on the lateral-directional characteristics of a three-surface F-15 model at transonic speeds (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the Langley 16 Foot Transonic Tunnel to determine the lateral directional aerodynamic characteristics of a fully metric 0.04 scale model of the F-15 three surface configuration (canards, horizontal tails) with twin two dimensional nozzles and twin axisymmetric nozzles installed. The effects of two dimensional nozzle in flight thrust reversing and rudder deflection were also determined. Test data were obtained at static conditions and at Mach numbers from 0.60 to 1.20 over an angle of attack range from -2 deg to 15 deg. Reynolds number varied from 2.6 million to 3.8 million. Angle of sideslip was set at approximately 0 deg and -5 deg for all configurations and at -10 deg for selected configurations.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2234 , L-15648 , NAS 1.60:2234
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  • 144
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    Thrust-induced effects on subsonic longitudinal aerodynamic characteristics of a vectored-engine-over-wing configuration (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the Langley 4 by 7 Meter Tunnel of the thrust induced effects on the longitudinal aerodynamic characteristics of a vectored-engine-over-wing fighter aircraft. The investigation was conducted at Mach numbers from 0.14 to 0.17 over an angle-of-attack range from -2 deg to 26 deg. The major model variables were the spanwise blowing nozzle sweep angle and main nozzle vector angle along with trailing edge, flap deflections. The overall thrust coefficient (main and spanwise nozzles) was varied from 0 (jet off) to 2.0. The results indicate that the thrust-induced effects from the main nozzle alone were small and mainly due to boundary-layer control affecting a small area behind the nozzle. When the spanwise blowing nozzles were included, the induced effects were larger than the main nozzle alone and were due to both boundary layer control and induced circulation lift. No leading edge vortex effects were evident.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2228 , L-15629 , NAS 1.60:2228
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  • 145
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    Aerodynamic force measurements with a strain-gage balance in a cryogenic wind tunnel (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Aerodynamic force measurements on a generalized 75 deg delta wing model with sharp leading edges were made with a three component internal strain gage balance in a cryogenic wind tunnel at stagnation temperatures of 300 K, 200 K, and 110 K. The feasibility of using a strain gage balance without thermal control in a cryogenic environment as well as the use of electrical resistance heaters, an insulator between the model and the balance, and a convection shield on the balance was investigated. Force and moment data on the delta wing model as measured by the balance are compared at the different temperatures while holding constant either the Reynolds number or the tunnel stagnation pressure. Tests were made at Mach numbers of 0.3 and 0.5 and at angles of attack up to 29 deg. The results indicate that it is feasible to acquire accurate force and moment data while operating at steady state thermal conditions in a cryogenic wind tunnel, either with or without electrical heaters on the balance. Within the limits of the balance accuracy, there were no apparent Reynolds number effects on the aerodynamic results for the delta wind model.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2251 , L-15685 , NAS 1.60:2251
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  • 146
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    Internal pressure distributions for a two-dimensional thrust-reversing nozzle operating at a free-stream Mach number of zero (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the static test facility of the Langley 16-Foot Transonic Tunnel to measure static pressure distributions inside a nonaxisymmetric thrust reversing nozzle. The tests were made at nozzle total pressures ranging from ambient to about eight times ambient pressure at a free stream Mach number of zero. Tabulated pressure data are presented.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85655 , L-15582 , NAS 1.15:85655
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  • 147
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    Vortex flap flow reattachment line and subsonic longitudinal aerodynamic data on 50 deg to 74 deg Delta wings on common fuselage (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Positions of the primary vortex flow reattachment line and longitudinal aerodynamic data were obtained at Mach number 0.3 for a systematic series of vortex flaps on delta wing body configurations with leading edge sweeps of 50, 58, 66, and 74 deg. The investigation was performed to study the parametric effects of wing sweep, vortex flap geometry and deflection, canards, and trailing edge flaps on the location of the primary vortex reattachment line relative to the flap hinge line. The vortex reattachment line was located via surface oil flow photographs taken at selected angles of attack. Force and moment measurements were taken over an angle of attack range of -1 deg to 22 deg at zero sideslip angle for many configurations to further establish the data base and to assess the aforementioned parametric effects on longitudinal aerodynamics. Both the flow reattachment and aerodynamic data are presented.
    Keywords: AERODYNAMICS
    Type: NASA-TM-84618 , L-15702 , NAS 1.15:84618
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  • 148
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    A Computer Program for the Calculation of Three-Dimensional Transonic Nacelle/Inlet Flowfields (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A highly efficient computer analysis was developed for predicting transonic nacelle/inlet flowfields. This algorithm can compute the three dimensional transonic flowfield about axisymmetric (or asymmetric) nacelle/inlet configurations at zero or nonzero incidence. The flowfield is determined by solving the full-potential equation in conservative form on a body-fitted curvilinear computational mesh. The difference equations are solved using the AF2 approximate factorization scheme. This report presents a discussion of the computational methods used to both generate the body-fitted curvilinear mesh and to obtain the inviscid flow solution. Computed results and correlations with existing methods and experiment are presented. Also presented are discussions on the organization of the grid generation (NGRIDA) computer program and the flow solution (NACELLE) computer program, descriptions of the respective subroutines, definitions of the required input parameters for both algorithms, a brief discussion on interpretation of the output, and sample cases to illustrate application of the analysis.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166528 , NAS 1.26:166528 , LG83ER0163
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  • 149
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    A grid-embedding transonic flow analysis computer program for wing/nacelle configurations (1983)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: An efficient grid-interfacing zonal algorithm was developed for computing the three-dimensional transonic flow field about wing/nacelle configurations. the algorithm uses the full-potential formulation and the AF2 approximate factorization scheme. The flow field solution is computed using a component-adaptive grid approach in which separate grids are employed for the individual components in the multi-component configuration, where each component grid is optimized for a particular geometry such as the wing or nacelle. The wing and nacelle component grids are allowed to overlap, and flow field information is transmitted from one grid to another through the overlap region using trivariate interpolation. This report represents a discussion of the computational methods used to generate both the wing and nacelle component grids, the technique used to interface the component grids, and the method used to obtain the inviscid flow solution. Computed results and correlations with experiment are presented. also presented are discussions on the organization of the wing grid generation (GRGEN3) and nacelle grid generation (NGRIDA) computer programs, the grid interface (LK) computer program, and the wing/nacelle flow solution (TWN) computer program. Descriptions of the respective subroutines, definitions of the required input parameters, a discussion on interpretation of the output, and the sample cases illustrating application of the analysis are provided for each of the four computer programs.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166529 , NAS 1.26:166529 , LG83ER0164
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  • 150
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    Analytical model of rotor wake aerodynamics in ground effect (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The model and the computer program developed provides the velocity, location, and circulation of the tip vortices of a two-blade helicopter in and out of the ground effect. Comparison of the theoretical results with some experimental measurements for the location of the wake indicate that there is excellent accuracy in the vicinity of the rotor and fair amount of accuracy far from it. Having the location of the wake at all times enables us to compute the history of the velocity and the location of any point in the flow. The main goal of out study, induced velocity at the rotor, can also be calculated in addition to stream lines and streak lines. Since the wake location close to the rotor is known more accurately than at other places, the calculated induced velocity over the disc should be a good estimate of the real induced velocity, with the exception of the blade location, because each blade was replaced only by a vortex line. Because no experimental measurements of the wake close to the ground were available to us, quantitative evaluation of the theoretical wake was not possible. But qualitatively we have been able to show excellent agreement. Comparison of flow visualization with out results has indicated the location of the ground vortex is estimated excellently. Also the flow field in hover is well represented.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166533 , NAS 1.26:166533
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  • 151
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    Investigation of installation effects on twin-engine convergent-divergent nozzles (1983)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine installation effects on convergent-divergent nozzles applicable to twin-engine reduced-power supersonic cruise aircraft. Tests were conducted at Mach numbers from 0.50 to 1.20, angles of attack from -5 deg to 9 deg, and at nozzle pressure ratios from jet off (1.0) to 8.0. The effects of empennage arrangement, nozzle length, and afterbody closure on total and component drag coefficients were investigated.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2205 , L-15609 , NAS 1.60:2205
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  • 152
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    Specification for a Program for an Interative Aeroelastic Solution (PIAS) (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An engineering and software specification which was written for a computer program to calculate aeroelastic structural loads including the effects of nonlinear aerodynamics is presented. The procedure used in the program for an iterative aeroelastic solution (PIAS) is to alternately execute two computer codes: one to calculate aerodynamic loads for a specific wing shape, and another to calculate the deflected shape caused by this loading. A significant advantage to the design of PIAS is that the initial aerodynamic module can be replaced with others. The leading edge vortex (LEV) program is used as the aerodynamic module in PIAS. This provides the capability to calculate aeroelastic loads, including the effects of a separation induced leading edge vortex. The finite element method available in ATLAS Integrated structural analysis and design system is used to determine the deflected wing shape for the applied aerodynamics and inertia loads. The data management capabilities in ATLAS are used by the execution control monitor (ECM) of PIAS to control the solution process.
    Keywords: AERODYNAMICS
    Type: NASA-CR-172200 , NAS 1.26:172200 , D6-52134
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  • 153
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    Tabulation of data from tests of an NPL 9510 airfoil in the Langley 0.3-meter transonic cryogenic tunnel (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The tabulated data from tests of a six inch chord NPL 9510 airfoil in the Langley 0.3-Meter Transonic Cryogenic Tunnel. The tests were performed over the following range of conditions: Mach numbers of 0.35 to 0.82, total temperature of 94 K to 300 K, total pressure of 1.20 to 5.81 atm, Reynolds number based on chord of 1.34 x 10 to the 6th to 48.23 x 10 to the 6th, and angle of attack of 0 deg to 6 deg. The NPL 9510 airfoil was observed to have decreasing drag coefficient up to the highest test Reynolds number.
    Keywords: AERODYNAMICS
    Type: NASA-TM-84579 , NAS 1.15:84579
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  • 154
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    XTRAN2L: A program for solving the general-frequency unsteady transonic small disturbance equation (1983)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: A program, XTRAN2L, for solving the general-frequency unsteady transonic small disturbance potential equation was developed. It is a modification of the LTRAN2-NLR code. The alternating-direction-implicit (ADI) method of Rizzetta and Chin is used to advance solutions of the potential equation in time Engquist-Osher monotone spatial differencing is used in the ADI solution algorithm. As a result, the XTRAN2L code is more robust and more efficient than similar codes that use Murman-Cole type-dependent spatial differencing. Nonreflecting boundary conditions that are consistent with the general-frequency equation have been developed and implemented at the far-field boundaries. Use of those conditions allow the computational boundaries to be moved closer to the airfoil with no loss of accuracy. This makes the XTRAN2L code more economical to use.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85723 , NAS 1.15:85723
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  • 155
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    Predicting aerodynamic characteristics of vortical flows on three-dimensional configurations using a surface-singularity panel method (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A general low-order surface-singularity panel method is used to predict the aerodynamic characteristics of a problem where a wing-tip vortex from one wing closely interacts with an aft mounted wing in a low Reynolds Number flow; i.e., 125,000. Nonlinear effects due to wake roll-up and the influence of the wings on the vortex path are included in the calculation by using a coupled iterative wake relaxation scheme. The interaction also affects the wing pressures and boundary layer characteristics: these effects are also considered using coupled integral boundary layer codes and preliminary calculations using free vortex sheet separation modelling are included. Calculated results are compared with water tunnel experimental data with generally remarkably good agreement.
    Keywords: AERODYNAMICS
    Type: AGARD Aerodyn. of Vortical Type Flows in Three Dimensions; 12 p
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  • 156
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    Helicopter rotor wake geometry and its influence in forward flight. Volume 2: Wake geometry charts (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Isometric and projection view plots, inflow ratio nomographs, undistorted axial displacement nomographs, undistorted longitudinal and lateral coordinates, generalized axial distortion nomographs, blade/vortex passage charts, blade/vortex intersection angle nomographs, and fore and aft wake boundary charts are discussed. Example condition, in flow ratio, undistorted axial location, longitudinal and lateral coordinates, axial coordinates distortions, blade/tip vortex intersections, angle of intersection, and fore and aft wake boundaries are also discussed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3727 , NAS 1.26:3727 , R83-912666-58
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  • 157
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    Experimental investigation of forebody and wing leading-edge vortex interactions at high angles of attack (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental investigation was conducted to assess the vortex flow-field interactions on an advanced, twin-jet fighter aircraft configuration at high angles of attack. Flow-field surveys were conducted on a small-scale model in the Northrop 0.41 - by 0.60-meter water tunnel and, where appropriate, the qualitative observations were correlated with low-speed wind tunnel data trends obtained on a large-scale model of the advanced fighter in the NASA Langley Research Center 30- by 60-foot (9.1- by 18.3-meter) facility. Emphasis was placed on understanding the interactions of the forebody and LEX-wing vortical flows, defining the effects on rolling moment variation with sideslip, and identifying modifications to control or regulate the vortex interactions at high angles of attack. The water tunnel flow visualization results and wind tunnel data trend analysis revealed the potential for strong interactions between the forebody and LEX vortices at high angles of attack. In particular, the forebody flow development near the nose could be controlled by means of carefully-positioned radome strakes. The resultant strake-induced flow-field changes were amplified downstream by the more powerful LEX vortical motions with subsequent large effects on wing flow separation characteristics.
    Keywords: AERODYNAMICS
    Type: AGARD Aerodyn. of Vortical Type Flows in Three Dimensions; 20 p
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  • 158
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    A computer program for estimating the aerodynamic characteristics of NACA 16-series airfoils (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A computer program written in a table ""look-up'' format, is presented which provides a comprehensive data base on NACA 16-series airfoils. The geometry covered is limited to cambers for a design-lift coefficient from 0.0 to 0.7 and thickness ratios from 4 to 21%. The data include Mach numbers from 0.3 to 1.6, angles of attack from -4 to 8 degrees, and lift coefficients from 0.0 to 0.8. Extrapolation is used to obtain data from Mach numbers, angles of attack, and lift coefficients beyond those for which data are available. A routine to adjust the lift and drag coefficients beyond stall is included. The uses and limitations of the program are also discussed.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85696 , NAS 1.15:85696
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  • 159
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    Effects of Nacelle configuration/position on performance of subsonic transport (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental study was conducted to explore possible reductions in installed propulsion system drag due to underwing aft nacelle locations. Both circular (C) and D inlet cross section nacelles were tested. The primary objectives were: to determine the relative installed drag of the C and D nacelle installations; and, to compare the drag of each aft nacelle installation with that of a conventional underwing forward, drag of each aft nacelle installation with that of a conventional underwing forward, pylon mounted (UTW) nacelle installation. The tests were performed in the NASA-Langley Research Center 16-Foot Transonic Wind Tunnel at Mach numbers from 0.70 to 0.85, airplane angles of attack from -2.5 to 4.1 degrees, and Reynolds numbers per foot from 3.4 to 4.0 million. The nacelles were installed on the NASA USB full span transonic transport model with horizontal tail on. The D nacelle installation had the smallest drag of those tested. The UTW nacelle installation had the largest drag, at 6.8 percent larger than the D at Mach number 0.80 and lift coefficient (C sub L) 0.45. Each tested configuration still had some interference drag, however. The effect of the aft nacelles on airplane lift was to increase C sub L at a fixed angle of attack relative to the wing body. There was higher lift on the inboard wing sections because of higher pressures on the wing lower surface. The effects of the UTW installation on lift were opposite to those of the aft nacelles.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3743 , NAS 1.26:3743 , LR-30436
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  • 160
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    Rotor/body aerodynamic interactions (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A wind tunnel investigation was conducted in which independent, steady state aerodynamic forces and moments were measured on a 2.24 m diam. two bladed helicopter rotor and on several different bodies. The mutual interaction effects for variations in velocity, thrust, tip-path-plane angle of attack, body angle of attack, rotor/body position, and body geometry were determined. The results show that the body longitudinal aerodynamic characteristics are significantly affected by the presence of a rotor and hub, and that the hub interference may be a major part of such interaction. The effects of the body on the rotor performance are presented.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85844 , A-9500 , NAS 1.15:85844 , USAAVRADCOM-TR-83-A-12
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  • 161
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    Reynolds number tests of an NPL 9510 airfoil in the Langley 0.3-meter transonic cryogenic tunnel (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation of the NPL 9510 airfoil was conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel over the following ranges of test conditions: Mach number of 0.35 to 0.82, total temperature of 94 K to 300 K, total pressure of 1.20 to 5.81 atm, Reynolds number based on airfoil chord of 1.34 x 10 to the 6th power to 48.23 x 10 to the 6th power, and angle of attack of 0 deg to 6 deg. The drag creep previously reported by the British National Physics Laboratory at low Reynolds numbers was also found to be present at high Reynolds numbers; the section drag coefficient continued to decrease even at the highest Reynolds number tested. Tests made close to free-stream saturation did not produce altered aerodynamic coefficients due to condensation effects.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85663 , L-15585 , NAS 1.15:85663
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  • 162
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    Flight measurements of hinged-plate wing-spoiler hinge moments (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Hinge moment of hinged-plate wing spoilers were measured during flight of a twin turboprop airplane modified by the addition of upper and lower wing-surface spoilers. The spoiler-actuating hydraulic cylinders were instrumented to measure the forces required to extend the spoiler panels. Those measurements were converted to moment coefficient form, and are presented as a function of spoiler deployment angle. The hinge-moment data were collected at three flight conditions: with flaps extended at approach speed; with flaps retracted at a low speed; and with flaps retracted at a high speed (C sub L = 1.4, 1.0, and 0.5). In general, the magnitude of measured spoiler hinge moments were lower than predicted. Furthermore, for upper surface spoilers with flaps extended, the hinge moments increased in a discontinuous manner between spoiler deflection 10 and 10.
    Keywords: AERODYNAMICS
    Type: NASA-TM-84343 , A-9282 , NAS 1.15:84343
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  • 163
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    Subsonic high-angle-of-attack aerodynamic characteristics of a cone and cylinder with triangular cross sections and a cone with a square cross section (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Experiments were conducted in the 12-Foot Pressure Wind Tunnel at Ames Research Center on three models with noncircular cross sections: a cone having a square cross section with rounded corners and a cone and cylinder with triangular cross sections and rounded vertices. The cones were tested with both sharp and blunt noses. Surface pressures and force and moment measurements were obtained over an angle of attack range from 30 deg to 90 deg and selected oil-flow experiments were conducted to visualize surface flow patterns. Unit Reynolds numbers ranged from 0.8x1,000,000/m to 13.0x1,000,000/m at a Mach number of 0.25, except for a few low-Reynolds-number runs at a Mach number of 0.17. Pressure data, as well as force data and oil-flow photographs, reveal that the three dimensional flow structure at angles of attack up to 75 deg is very complex and is highly dependent on nose bluntness and Reynolds number. For angles of attack from 75 deg to 90 deg the sectional aerodynamic characteristics are similar to those of a two dimensional cylinder with the same cross section.
    Keywords: AERODYNAMICS
    Type: NASA-TM-84377 , A-9392 , NAS 1.15:84377
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  • 164
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    Effect of thrust vectoring and wing maneuver devices on transonic aeropropulsive characteristics of a supersonic fighter (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The aeropropulsive characteristics of an advanced fighter designed for supersonic cruise were determined in the Langley 16-Foot Transonic Tunnel. The objectives of this investigation were to evaluate the interactive effects of thrust vectoring and wing maneuver devices on lift and drag and to determine trim characteristics. The wing maneuver devices consisted of a drooped leading edge and a trailing-edge flap. Thrust vectoring was accomplished with two dimensional (nonaxisymmetric) convergent-divergent nozzles located below the wing in two single-engine podded nacelles. A canard was utilized for trim. Thrust vector angles of 0 deg, 15 deg, and 30 deg were tested in combination with a drooped wing leading edge and with wing trailing-edge flap deflections up to 30 deg. This investigation was conducted at Mach numbers from 0.60 to 1.20, at angles of attack from 0 deg to 20 deg, and at nozzle pressure ratios from about 1 (jet off) to 10. Reynolds number based on mean aerodynamic chord varied from 9.24 x 10 to the 6th to 10.56 x 10 to the 6th.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2119 , L-15526 , NAS 1.60:2119
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  • 165
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    An experimental investigation of a free-tip rotor configuration in a forward flight wind-tunnel test (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Results from an experimental evaluation of a free-tip rotor are presented. The evaluation included whirl tests and wind tunnel tests up to advance ratios of 0.4. The free tip extended over the outer 5% of the rotor blade and included a passive mechanical controller whose output characteristics were varied. The controller configuration combined with the free tip aerodynamics resulted in higher power requirements, because the tip's pitch angle was 5 to 10 degrees greater than that of the inboard portion of the blade, and its pitching motion capability was considered to be inhibited by frictional forces. Recommendations are included for design features for a follow-on test.
    Keywords: AERODYNAMICS
    Type: NASA-TM-84409 , A-9485 , NAS 1.15:84409
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  • 166
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    Magnus effects at high angles of attack and critical Reynolds numbers (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The Magnus force and moment experienced by a yawed, spinning cylinder were studied experimentally in low speed and subsonic flows at high angles of attack and critical Reynolds numbers. Flow-field visualization aided in describing a flow model that divides the Magnus phenomenon into a subcritical region, where reverse Magnus loads are experienced, and a supercritical region where these loads are not encountered. The roles of the spin rate, angle of attack, and crossflow Reynolds number in determining the boundaries of the subcritical region and the variations of the Magnus loads were studied.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-2145
    Format: text
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  • 167
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    Computational aerodynamics applications to transport aircraft design (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Examples are cited in assessing the effect that computational aerodynamics has had on the design of transport aircraft. The application of computational potential flow methods to wing design and to high-lift system design is discussed. The benefits offered by computational aerodynamics in reducing design cost, time, and risk are shown to be substantial.These aerodynamic methods have proved to be particularly effective in exposing inferior or poor aerodynamic designs. Particular attention is given to wing design, where the results have been dramatic.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-2061
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  • 168
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    SC3 - A wing concept for supersonic maneuvering (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A supersonic wing design concept that overcomes the problems associated with nonlinear inviscid pressure distribution is presented together with a summary of the supporting theoretical and experimental development program. The key design feature is a conical panel supporting a controlled supercritical conical camber (SC3) crossflow expansion and recompression which permits lift on the upper surface to be obtained without producing an adverse pressure gradient or crossflow shock wave strong enough to separate the boundary layer. The role of aerodynamics in the design implementation is discussed, as are the concept development, the initial design validation on a fabricated and tested SC3 design, and the body and canard interaction effect. The design and testing of an isolated wing with a planform and thickness distribution representative of a practical application are described. Wind-tunnel results show that significant performance gains can be obtained for fighter aircraft using this concept.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1858
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  • 169
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    Analysis of surface pressure distributions on two elliptic missile configurations (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The state-of-the-art methods for predicting missile aerodynamic characteristics do not accurately predict the loads of missile configurations with bodies of elliptic cross section. An investigation of this problem found significant nonlinear flow disturbance on the windward surface of a 3:1 elliptic body at Mach 2.50 in addition to the nonlinear vortical flows which develop on the leeside. A nonlinear full-potential flow method (NCOREL) was found to provide extremely accurate pressure estimates for attached-flow conditions and the vortex prediction method contained in the state-of-the-art method (NOSEVTX) was shown to accurately calculate body vortices and leeside pressures.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1841
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  • 170
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    Evaluation of factors determining the accuracy of linearized subsonic panel methods (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A systematic evaluation of the factors determining the accuracy of linearized subsonic panel methods is presented. In particular, the constant and quadratically varying doublet panel methods are compared for thin and thick surface modelings in two and three dimensions. The sensitivity of results to panel edge and control point locations is studied for both of the methods. The first order convergence of the quadratic doublet method near network edges and the subsequent effect on the Kutta condition is investigated. Results from a quadratic doublet method specifically designed for a vector processing computer are shown.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1826
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  • 171
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    Numerical investigation of unsteady flow development in a nozzle-duct configuration (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A computational approach has been used to resolve the transient phenomena caused by a rocket engine starting up and exhausting into a short duct. The results are obtained from the finite-difference Navier-Stokes, continuity, and energy equations for a single component gas. Using a recently developed numerical technique to account for the time-varying boundary conditions at the nozzle throat and the duct openings, the numerical simulation of the idealized ignition flow model has yielded some insight into the complex interactions between the wave patterns and the vortical flow. The present findings verify the hypothesis that a lateral expansion of the jetstream outside the nozzle will generate an over-pressure pulse on the duct wall if the jetstream is restricted within the nozzle by flow separation. The numerical simulations of this complex flow problem should be useful for improving simple correlation techniques used for engineering purposes.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1714
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  • 172
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    Heat-transfer distributions on biconics at incidence in hypersonic-hypervelocity He, N2, air, and CO2 flows (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Laminar heat-transfer rates were measured on spherically blunted, 13 degrees/F degrees on-axis and bent biconics (fore cone bent 7 degrees upward relative to aft cone) at hypersonic-hypervelocity flow conditions in the Langley Expansion Tube. Freestream velocities from 4.5 to 6.9 km/sec and Mach numbers from 6 to 9 were generated using helium, nitrogen, air, and carbon dioxide test gases, resulting in normal shock density ratios from 4 to 19. Angle of attack, referenced to the axis of the aft cone, was varied from zero to 20 degrees in 4 degree increments. The effect of nose bend, angle of attack, and real-gas phenomena on heating distributions are presented along with comparisons of measurement to prediction from a code which solves the three-dimensional 'parabolized Navier-Stokes' equations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1508
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  • 173
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    Heating analysis of bent-nose biconics at high angles of attack using the parabolized Navier-Stokes equations (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A method based on the Parabolized Navier-Stokes equations is used to calculate the flow field and heat transfer of lifting entry vehicles. The method is based on the Bean and Warming implicit algorithm and uses a new procedure for preventing departure solutions. Calculations are carried out for blunt on-axis and bent biconics, assuming a perfect gas and laminar flow, and compared with available heat transfer, surface pressure and shock shape measurements for a range of Mach numbers and angles of attack. In all calculations presented here, the starting solution is obtained from available inviscid and boundary layer codes. Good agreement with experiment is indicated. Thus, the method provides an accurate and rather inexpensive procedure for calculating three-dimensional flows at supersonic Mach numbers.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1507
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  • 174
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    Viscous shock-layer flowfield analysis by an explicit-implicit method (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The present work extends the recently reported implicit analogue of MacCormack's earlier widely-used explicit method to external axisymmetric laminar flows with strong entropy gradients. The details of the 'numerics' of the implicit part are provided in a body-oriented coordinate system with a moving outer (shock) boundary during the transient part of the solutions. The limiting values of the Courant number are obtained when the shock boundary is treated explicitly. The solution algorithm outlined includes the treatment of the source term associated with the equations in weak conservation form. From the results obtained for two sample problems, it becomes clear that accuracy of predictions is, indeed, very good at higher values of the Courant number. There is a significant saving in overall computing time, depending on the Courant number used and the flow Reynolds number. These properties combined with the simplicity of programming the implicit analogue may appeal to researchers for using it in the analysis of three-dimensional flow problems.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1423
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  • 175
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    Prediction of high speed propeller flow fields using a three-dimensional Euler analysis (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Previously announced in STAR as N83-19710
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-0188
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  • 176
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    Zero-length inlets for subsonic V/STOL aircraft (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Previously cited in issue 19, p. 3257, Accession no. A81-42186
    Keywords: AERODYNAMICS
    Type: (ISSN 0021-8669)
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  • 177
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    Investigation of the tip clearance flow inside and at the exit of a compressor rotor passage. II - Turbulence properties (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The flow in the tip clearance region of a compressor rotor is highly turbulent due to the strong interaction of the leakage flow with the annulus wall boundary layer. This paper deals with the turbulence properties of the flow in the tip clearance region of a moderately loaded compressor rotor. The experimental results reported in this paper were obtained using a two-sensor hot-wire probe in combination with an ensemble averaging technique. Blade-to-blade distribution of the axial and tangential turbulence intensities at various radial locations and ten axial locations (four inside the blade passage and the remaining six outside the passage) were derived from this data. Isointensity contours in the clearance region at various radial locations were also obtained from the experimental data. A region of very high turbulence intensities was indicated at the half-chord location from these results. The turbulence intensity profiles also indicated that the leakage flow travels toward the midpassage before rolling up. The turbulence is almost isotropic beyond three-quarter chord downstream of the trailing edge.
    Keywords: AERODYNAMICS
    Type: ASME, Transactions, Journal of Engineering for Power (ISSN 0022-0825); 105; Jan. 198
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  • 178
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    Blade loading and rotation effects on compressor rotor wake near end walls (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: (Previously cited in issue 08, p. 1179, Accession no. A82-22063)
    Keywords: AERODYNAMICS
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  • 179
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    Calculation of symmetric vortex separation affecting subsonic bodies at high incidence (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A method for calculating the longitudinal aerodynamic coefficients and the pressure distributions on a body at reasonably high angles of attack is presented. The body is represented by a combination of source elements and vortex-lattice elements, including separation of the vortices at increasing angles of attack. The method is self-consistent in that the body and the separated vortex wake are treated as an integrated interacting system. The location of the separation line can be included as an arbitrary input from experimental data or can be evaluated approximately by a pressure-dependent criterion. Calculated values of the aerodynamic coefficients and pressure distributions on cone-cylinder and ogive-cylinder bodies compare well, qualitatively and quantitatively, with experimental data, including simulation of the dependence on Reynolds number.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 21; Mar. 198
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  • 180
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    Turbulence measurements in two shock-wave/shear-layer interactions (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Measurements of the longitudinal component of the mass-flow fluctuations have been made in an 8-deg compression corner flow and a reattaching shear layer at a Mach number of approximately 2.9 and unit Reynolds number of about 7 x 10 to the 7th per m. The data include turbulence intensities and probability density distributions. Significant turbulence amplification occurs in both interactions. A qualitative explanation in terms of direct shock effects, extra strain rates, and their relative contribution, is suggested.
    Keywords: AERODYNAMICS
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  • 181
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    Comparison of calculated and measured pressures on straight and swept-tip model rotor blades (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Using the quasi-steady, full potential code, ROT22, pressures were calculated on straight and swept tip model helicopter rotor blades at advance ratios of 0.40 and 0.45, and into the transonic tip speed range. The calculated pressures were compared with values measured in the tip regions of the model blades. Good agreement was found over a wide range of azimuth angles when the shocks on the blade were not too strong. However, strong shocks persisted longer than predicted by ROT22 when the blade was in the second quadrant. Since the unsteady flow effects present at high advance ratios primarily affect shock waves, the underprediction of shock strengths is attributed to the simplifying, quasi-steady, assumption made in ROT22.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85872 , A-9584 , NAS 1.15:85872
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  • 182
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    Loads and aeroelasticity division research and technology accomplishments for FY 1982 and plans for FY 1983 (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Accomplishments of the past year and plans for the coming year are highlighted as they relate to five year plans and the objectives of the following technical areas: aerothermal loads; multidisciplinary analysis and optimization; unsteady aerodynamics; and configuration aeroelasticity. Areas of interest include thermal protection system concepts, active control, nonlinear aeroelastic analysis, aircraft aeroelasticity, and rotorcraft aeroelasticity and vibrations.
    Keywords: AERODYNAMICS
    Type: NASA-TM-84594 , NAS 1.15:84594
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  • 183
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    Flow improvements in the circuit of the Langley 4- by 7-meter tunnel (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The mean velocity profiles in both the horizontal and vertical planes of symmetry at specific locations throughout the tunnel circuit to identify the most promising means for improving the flow in the 4 by 7 meter wind tunnel were measured. In the base line tunnel flow surveys, the flow patterns near the end of the test section indicate a uniform mean velocity distribution. Downstream of the test section, unsymmetrical flow patterns result in low velocities along the inner walls and in flow separation along the inner wall of the diffuser upstream of the drive fan and along the outer wall of the large diffuser downstream of the drive fan. A set of trailing-edge flaps attached to the five flow-control vanes located just downstream of the first corner were installed. These flaps are successful in making the tunnel flow more symmetrical and in eliminating the regions of separation in the diffusers upstream and downstream of the drive fan.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85662 , L-15631 , NAS 1.15:85662
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  • 184
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    Nonequilibrium radiative heating of an ablating Jovian entry probe (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Previously cited in issue 08, p. 1337, Accession no. A80-23948
    Keywords: AERODYNAMICS
    Type: (ISSN 0001-1452)
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  • 185
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    A grid overlapping scheme for flowfield computations about multicomponent configurations (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Previously cited in issue 18, p. 2841, Accession no. A82-37477
    Keywords: AERODYNAMICS
    Type: (ISSN 0001-1452)
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  • 186
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    Aerodynamic characteristics of missile control fins in nonlinear flow fields (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Recent experimental results show that the control effectiveness of a missile fin in supersonic flow at moderate-to-high angles of attack is a strong nonlinear function of free-stream Mach number, body incidence angle, fin bank angle and fin deflection angle. Analysis of the experimental results using an Euler finite-difference computer code with flow separation together with the equivalent angle-of-attack concept indicates that the observed nonlinearities are due to the variation of local dynamic pressure and local Mach number around the missile body alone. The nonlinearities are shown to be a strong source of control cross-coupling for high Mach number, high angle-of-attack combinations. The analysis suggests a relatively simple yet comprehensive approach for accurately accounting for these nonlinear effects.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-2083
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  • 187
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    Influence of airplane components on rotational aerodynamic data for a typical single-engine airplane (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The influence of airplane components, as well as wing location and tail length, on the rotational flow aerodynamics is discussed for a 1/6-scale general aviation airplane model. Examination of the individual component effects provides insight as to the source of the moment and forces that determine spin characteristics. For the subject airplane, it is seen that the presence of the horizontal tail adversely effects the yaw damping of the vertical tail such that at certain angles of attack the vertical actually becomes a propelling component in spinning motion. It is shown that the level of this adverse interaction is a function of the relative positions of the wing and the tail surfaces, e.g., relocating the wing to the high wing position or shortening the tail moment arm from the basic configuration both reduced or eliminated the adverse tail interference for certain angles of attack. The influence of wing location on the damping characteristics of the horizontal-vertical tail combination would not be addressed in any existing tail design criterion for spinning and could only be discerned through rotary balance testing.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-2135
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  • 188
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    Interactions of airfoils with gusts and concentrated vortices in unsteady transonic flow (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Unsteady interactions of concentrated vortices and distributed free-stream gusts with a stationary airfoil have been analyzed in two-dimensional transonic flow. A simple method of introducing such disturbances has been implemented numerically in the well-known transonic small-disturbance code LTRAN2, and calculations have been performed for two important classes of current aerodynamic problems. The first, which demonstrates many of the essential features of the interactions between helicopter rotor blades and their trailing-vortex wakes, is that of a discrete potential vortex convecting past an airfoil. The second is the response of a transonic airfoil to a transverse periodic gust, with and without the alleviation that can be achieved by the proper active control motion of a trailing-edge flap. In both cases, unsteady effects are found to play important roles in the shock-wave motion, in the overall flow-field development, and consequently, in the air loads on the airfoil.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1691
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  • 189
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    Computational aerodynamic design methodology (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The transonic wing design process has been vastly improved at Lockheed-Georgia. The revised design procedure enhances useability and reliability by combining numerical optimization and inverse design into a single wing design code with transonic analysis provided by a modified version of FLO22. A more versatile set of geometric decision variables has been integrated into the optimization portion for geometry perturbations. An automatic restart feature permits the interchangeability of solutions between optimization and inverse design as the design progresses. In combination, these improvements enable practical utilization of a VAX 11/780 computer and significantly reduce the elapsed time required to complete a transonic wing design.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1865
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  • 190
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    Improved method for transonic airfoil design-by-optimization (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: An improved method for use of optimization techniques in transonic airfoil design is demonstrated. FLO6QNM incorporates a modified quasi-Newton optimization package, and is shown to be more reliable and efficient than the method developed previously at NASA-Ames, which used the COPES/CONMIN optimization problem. The design codes are compared on a series of test cases with known solutions, and the effects of problem scaling, proximity of initial point to solution, and objective function precision are studied. In contrast to the older method, well-converged solutions are shown to be attainable in the context of engineering design using computational fluid dynamics tools, a new result. The improvements are due to better performance by the optimization routine and to the use of problem-adaptive finite difference step sizes for gradient evaluation.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1864
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  • 191
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    High angle of attack inviscid flow calculations over a Shuttle-like vehicle with comparisons to flight data (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Recent advances in computational techniques and computer hardware have made it possible to attack the problem of computing steady-state inviscid flowfields over complex three-dimensional bodies such as the Space Shuttle Orbiter. Certain problems have arisen in connection with cases involving high angles of attack. Weilmuenster et al. (1981, 1982) have, therefore, presented a time asymptotic computational method, HALIS, designed to handle the high angle-of-attack problem for the inviscid flow over the Space Shuttle vehicle. The code was modified to include all of the Shuttle vehicle forward of the wing root. Body surface pressures generated by the HALIS code were shown to be in excellent agreement with flight pressures measured on the Shuttle vehicle. In the present investigation, results from the HALIS code, which has been extended to cover the first 1212 inches of the Shuttle vehicle, are compared with flight surface pressures. Shock shapes at several angles of attack are also presented.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1798
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  • 192
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    Axisymmetric bluff-body drag reduction using circumferential grooves (1983)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The effect of shoulder radiusing and circumferentially grooving the afterbodies of bluff bodies to reduce the base drag at low speeds is discussed. Shoulder radii as large as 2.75 body diameters are examined. Reynolds number based on body diameter varied from 20,000 to 200,000. Results indicate that increasing the shoulder radius to 2.75 body diameters can reduce the drag levels to those of a streamline body having 67 percent greater fineness ratio. For zero shoulder radius, circumferential grooves were found to be effective in reducing body drag for zero shoulder radius in both laminar and tripped flow. Circumferential grooves on the afterbody with a shoulder radius of one-half the body diameter were only effective in reducing drag for laminar flow.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-1788
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  • 193
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    Axial wavenumber measurements in axisymmetric jets (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Previously cited in issue 06, p. 941, Accession no. A82-17752
    Keywords: AERODYNAMICS
    Type: (ISSN 0001-1452)
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  • 194
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    An interaction algorithm for three-dimensional turbulent subsonic aerodynamic juncture region flow (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: (Previously cited in issue 06, p. 796, Accession no. A82-17782)
    Keywords: AERODYNAMICS
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  • 195
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    A finite element formulation for steady transonic Euler equations (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: (Previously cited in issue 06, p. 795, Accession no. A82-17759)
    Keywords: AERODYNAMICS
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  • 196
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    Unsteady transonic flow over wings including inviscid/viscous interaction (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: (Previously cited in issue 07, p. 966, Accession no. A82-19797)
    Keywords: AERODYNAMICS
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  • 197
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    Lifting surface theory for a helicopter rotor in forward flight (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A lifting surface theory has been developed for a helicopter rotor in forward flight for incompressible flow. The method utilized the concept of the linearized acceleration potential and make use of the vortex lattice procedures. Results in terms of lift coefficient slope for several forward flight conditions are given.
    Keywords: AERODYNAMICS
    Type: NASA-CR-169997 , NAS 1.26:169997
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  • 198
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    Flap effectiveness on subsonic longitudinal aerodynamic characteristics of a modified arrow wing (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation of the subsonic longitudinal aerodynamic characteristics of a modified arrow-wing model was conducted in the Langley 4- by 7-Meter Tunnel. This investigation addressed the effectiveness of the leading and trailing edge flap deflections of this model. The arrow wing was tested at a Mach number of 0.02 and at an angle-of-attack range from -4 deg to 24 deg. The results of the investigation showed that deflecting the leading edge and trailing edge in combination could promote an attached-flow condition at the wing leading edge. Also, the leading edge suction could be maximized over the complete lift-coefficient range by scheduling a combination of leading and trailing edge flap deflections.
    Keywords: AERODYNAMICS
    Type: NASA-TM-84582 , L-15239 , NAS 1.15:84582
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  • 199
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    Calculation of potential flow past non-lifting bodies at angle of attack using axial and surface singularity methods (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Two different singularity methods have been utilized to calculate the potential flow past a three dimensional non-lifting body. Two separate FORTRAN computer programs have been developed to implement these theoretical models, which will in the future allow inclusion of the fuselage effect in a pair of existing subcritical wing design computer programs. The first method uses higher order axial singularity distributions to model axisymmetric bodies of revolution in an either axial or inclined uniform potential flow. Use of inset of the singularity line away from the body for blunt noses, and cosine-type element distributions have been applied to obtain the optimal results. Excellent agreement to five significant figures with the exact solution pressure coefficient value has been found for a series of ellipsoids at different angles of attack. Solutions obtained for other axisymmetric bodies compare well with available experimental data. The second method utilizes distributions of singularities on the body surface, in the form of a discrete vortex lattice. This program is capable of modeling arbitrary three dimensional non-lifting bodies. Much effort has been devoted to finding the optimal method of calculating the tangential velocity on the body surface, extending techniques previously developed by other workers.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166058 , NAS 1.26:166058
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  • 200
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    Pressure investigation of NASA leading edge vortex flaps on a 60 deg Delta wing (1983)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Pressure distributions on a 60 deg Delta Wing with NASA designed leading edge vortex flaps (LEVF) were found in order to provide more pressure data for LEVF and to help verify NASA computer codes used in designing these flaps. These flaps were intended to be optimized designs based on these computer codes. However, the pressure distributions show that the flaps wre not optimum for the size and deflection specified. A second drag-producing vortex forming over the wing indicated that the flap was too large for the specified deflection. Also, it became apparent that flap thickness has a possible effect on the reattachment location of the vortex. Research is continuing to determine proper flap size and deflection relationships that provide well-behaved flowfields and acceptable hinge-moment characteristics.
    Keywords: AERODYNAMICS
    Type: NASA-CR-169984 , NAS 1.26:169984
    Format: application/pdf
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