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  • 201
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    A Navier-Stokes calculation of the airfoil dynamic stall process (1984)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A time dependent Navier-Stokes calculation procedure has been applied to the problem of an NACA 0012 airfoil oscillating in pitch in a low Mach number, high Reynolds number environment. The calculated results show many of the known physical features, including sudden suction surface separation, vortices shed at the leading and trailing edges and the return to attached flow at low incidences. Both the lift and moment coefficient curves show the expected features and the calculated wall pressure coefficients show strong correspondence to measured data.
    Keywords: AERODYNAMICS
    Type: AD-P004163 , AFOSR Proc. of the Workshop on Unsteady Separated Flow; p 82-89
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  • 202
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    An experimental investigation of the parallel blade-vortex interaction (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A scheme for investigating the parallel blade vortex interaction (BVI) has been designed and tested. The scheme involves setting a vortex generator upstream of a nonlifting rotor so that the vortex interacts with the blade at the forward azimuth. The method has revealed two propagation mechanisms: a type C shock propagation from the leading edge induced by the vortex at high tip speeds, and a rapid but continuous pressure pulse associated with the proximity of the vortex to the leading edge. The latter is thought to be the more important source. The effects of Mach number and vortex proximity are discussed.
    Keywords: AERODYNAMICS
    Type: NASA-TM-86005 , A-9850 , NAS 1.15:86005 , USAAVSCOM-TM-84-A-9
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  • 203
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    Comparison of analytical and experimental steadyand unsteady-pressure distributions at Mach number 0.78 for a high-aspect-ratio supercritical wing model with oscillating control surfaces (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The unsteady aerodynamic lifting surface theory, the Doublet Lattice method, with experimental steady and unsteady pressure measurements of a high aspect ratio supercritical wing model at a Mach number of 0.78 were compared. The steady pressure data comparisons were made for incremental changes in angle of attack and control surface deflection. The unsteady pressure data comparisons were made at set angle of attack positions with oscillating control surface deflections. Significant viscous and transonic effects in the experimental aerodynamics which cannot be predicted by the Doublet Lattice method are shown. This study should assist development of empirical correction methods that may be applied to improve Doublet Lattice calculations of lifting surface aerodynamics.
    Keywords: AERODYNAMICS
    Type: NASA-TM-84589 , NAS 1.15:84589
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  • 204
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    Comparison of uniform perturbation solutions and numerical solutions for some potential flows past slender bodies (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Approximate solutions for potential flow past an axisymmetric slender body and past a thin airfoil are calculated using a uniform perturbation method and then compared with either the exact analytical solution or the solution obtained using a purely numerical method. The perturbation method is based upon a representation of the disturbance flow as the superposition of singularities distributed entirely within the body, while the numerical (panel) method is based upon a distribution of singularities on the surface of the body. It is found that the perturbation method provides very good results for small values of the slenderness ratio and for small angles of attack. Moreover, for comparable accuracy, the perturbation method is simpler to implement, requires less computer memory, and generally uses less computation time than the panel method. In particular, the uniform perturbation method yields good resolution near the regions of the leading and trailing edges where other methods fail or require special attention.
    Keywords: AERODYNAMICS
    Type: NASA-CR-172485 , ICASE-84-56 , NAS 1.26:172485
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  • 205
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    Two recent extensions of the vortex method (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The vortex method, coupled to an integral boundary-layer solver, is applied to the numerical simulation of high-Reynolds-number, separated flows in two new cases: a bluff body in a wind tunnel and flow in a cascade. In the bluff body case, the blockage effect of the tunnel walls is included approximately, assuming an inviscid boundary condition at the walls. The resulting increase in drag is computed, and compares well with a small-disturbance theory and with experiments. The results for flow in a cascade are compared with results from two finite-difference codes for a single-blade case, and good agreement is found. When a staggered cascade is treated with five independent blades, the simulation predicts rotating stall, depending on the angle of attack, and the essential features of the flow are correct. The sensitivity of this phenomenon to various parameters is studied and the stall boundary is found.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0343
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  • 206
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    Nonlinear aerodynamic effects on bodies in supersonic flow (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The supersonic flow about generic bodies was analyzed to identify the elments of the nonlinear flow and to determine the influence of geometry and flow conditions on the magnitude of these nonlinearities. The nonlinear effects were attributed to separated-flow nonlinearities and attached-flow nonlinearities. The nonlinear attached-flow contribution was further broken down into large-disturbance effects and entropy effects. Conical, attached-flow bundaries were developed to illustrate the flow regimes where the nonlinear effects are significant, and the use of these boundaries for angle of attack and three-dimensional geometries was indicated. Normal-force and pressure comparisons showed that the large-disturbance and separated-flow effects were the dominant nonlinear effects at low supersonic Mach numbers and that the entropy effects were dominant for high supersonic Mach number flow. The magnitude of all the nonlinear effects increased with increasing angle of attack. A full-potential method, NCOREL, which includes an approximate entropy correction, was shown to provide accurate attached-flow pressure estimates from Mach 1.6 through 4.6.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0231
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  • 207
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    PAN AIR modeling studies. II - Sideslip option, network gaps, three-dimensional forebody flow, and thick-trailing-edge representation (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: PAN AIR is a computer program that calculates linear potential flow about arbitrary configurations at both subsonic and supersonic Mach numbers. This paper is a follow-on of another paper entitled 'PAN AIR Modeling Studies', in which several studies were presented that exhibited PAN AIR's versatility for modeling diverse configurations. Results from four modeling studies of interest in modeling realistic aircraft shapes are presented. The topics addressed are (1) half-geometry option in sideslip, (2) network gaps, (3) three-dimensional forebody flows, and (4) trailing-edge representation.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0220
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  • 208
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    PAN AIR prediction of NASA Ames 12-foot pressure wind-tunnel interference on a fighter configuration (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Models tested in the NASA Ames 12-Foot Pressure Wind Tunnel over an angle of attack range from 0 deg to 90 deg are mounted on a floor strut that protrudes from a fairly large support bump. In high-angle-of-attack tests (angle of attack = 40 deg to 90 deg), for which the floor support was originally designed, the effects of the flow angularities produced by the bump are often negligible. This is not so for low-angle-of-attack tests (0 deg to 40 deg). Since there are no standard means for correcting test data for this bump effect, low-angle-of-attack testing with the bump is not recommended by the Ames wind-tunnel staff. This paper presents an exploratory study of a technique for correcting balance forces and experimental pressures for combined wall and bump effects. This is done by modeling the aircraft, wind-tunngl walls, and bump, with PAN AIR. The wall-and-bump-induced increments in the lift coefficient and pitching-moment coefficient predicted by PAN AIR are compared with increments obtained from the Ames 12-foot tunnel with the bump and an 8 x 12 low speed wind tunnel which has no bump.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0219
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  • 209
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    Flowfield scaling of a swept compression corner interaction A comparison of experiment and computation (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Over the past decade, there has been some important progress in understanding the problems of three-dimensional (3D) shock wave/turbulent boundary layer interactions. However, the problems are by no means solved. The present investigation has the objective to determine the flowfield structure of a swept compression corner interaction and to perform a test of an equation considered by Settles and Bogdonoff (1982). The experimental data obtained in a 20 x 20 cm high Reynolds number supersonic wind tunnel are compared with the predictions of a state-of-the-art numerical solution of the Navier-Stokes equations. It is found that a 3D scaling law for Reynolds number effects, previously established for interaction 'footprints', is equally valid when applied to the present flowfield.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0096
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  • 210
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    Improved finite difference schemes for transonic potential calculations (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Engquist and Osher (1980) have introduced a finite difference scheme for solving the transonic small disturbance equation, taking into account cases in which only compression shocks are admitted. Osher et al. (1983) studied a class of schemes for the full potential equation. It is proved that these schemes satisfy a new discrete 'entropy inequality' which rules out expansion shocks. However, the conducted analysis is restricted to steady two-dimensional flows. The present investigation is concerned with the adoption of a heuristic approach. The full potential equation in conservation form is solved with the aid of a modified artificial density method, based on flux biasing. It is shown that, with the current scheme, expansion shocks are not possible.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0092
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  • 211
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    Second-order-accurate spatial differencing for the transonic small-disturbance equation (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Current methods for calculating transonic flows with the small-disturbance potential equation are typically only first-order accurate in the supersonic regions of the flow. However, calculations using the full-potential equation show significant improvements in accuracy when second-order methods are used instead of first-order methods. In this paper, algorithms for the small-disturbance equations are presented, for both steady and unsteady flows, with spatial differencing that is second-order accurate in both the subsonic and supersonic regions of the flow. These algorithms are stable, simple extensions of implicit monotone algorithms; the former are only first-order accurate spatially in the supersonic regions. Several calculations in one and two dimensions have been made, and the first- and second-order calculations are compared. In one dimension, the calculations include the four types of shock-wave motion. In two dimensions, the comparisons include steady flow over a Korn airfoil and unsteady flow over a pitching airfoil. All the comparisons in two dimensions show that the new methods are just as stable numerically as the old methods, but improve the accuracy of the solutions. The algorithm improvements can be implemented in present computer codes by making minor coding modifications.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0091
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  • 212
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    Comparison of experimental and computational compressible flow in a S-duct (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: This paper describes experimental measurements of secondary flow in a constant area, circular cross-section 30-30 deg S-duct, and compares the results obtained with the computations performed using the PEPSIG code, a parabolized Navier-Stokes code. The flow entering the duct was turbulent, with entrance Mach number of 0.6, and the boundary layer thickness at the duct entrance was 10 percent of the duct diameter. The duct mean radius of curvature to the duct diameter was 5.077. Flow parameters were measured at six stations along the length of the duct. These measurements were made using a five-port cone probe. At least ten radial traverses were made at each station on both sides of the symmetry plane. Wall static pressures along three azimuth angles of zero, 90, and 180 deg along the duct were measured. Plots presenting the secondary velocity field as well as contour plots of the total and static-pressure fields have been obtained. Strong secondary flows were observed in the first bend, and these continued into the second bend with the formation of new vorticity in the opposite sense in the second bend. The flow exiting the duct contained two pairs of counter-rotating vortices. The computational results are in general agreement with the experiments. However, it appears that the computations underestimate the extent of the pressure distortion, due to simplifications made in the pressure field calculations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0033
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  • 213
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    Aerodynamic canard/wing parametric analysis for general aviation applications (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Vortex panel and vortex lattice methods have been utilized in an analytic study to determine the two- and three-dimensional aerodynamic behavior of canard and wing configurations. The purpose was to generate data useful for the design of general aviation canard aircraft. Essentially no two-dimensional coupling was encountered and the vertical distance between the lifting surfaces was found to be the main contributor to interference effects of the three-dimensional analysis. All canard configurations were less efficient than a forward wing with an aft horizontal tail, but were less sensitive to off-optimum division of total lift between the two surfaces, such that trim drag could be less for canard configurations. For designing a general aviation canard aircraft, results point toward large horizontal and vertical distance between the canard and wing, a large wing-to-canard area ratio, and with the canard at a low incidence angle relative to the wing.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0560
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  • 214
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    Calculation of steady and oscillating airfoil flow fields via the Navier Stokes equations (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A Navier-Stokes calculation procedure is applied to high Reynolds number flow about steady and unsteady airfoils. The procedure solves the ensemble averaged governing equations via a linearized block implicit (LBI) technique which in general converges to a nominal steady flow within 120 time steps. The grid used in the computation is highly stretched thus resolving the turbulent boundary layers. Calculations have been compared with data for both NACA 4412 airfoil at high incidence and an NACA 0012 airfoil oscillating in dynamic stall. In both cases, good agreement is noted between measured and calculated surface pressure distributions.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0525
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  • 215
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    Numerical simulation of the tip vortex off a low-aspect-ratio wing at transonic speed (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The viscous transonic flow around a low-aspect-ratio wing has been computed using an implicit, three-dimensional, 'thin-layer' Navier-Stokes solver. The grid around the geometry of interest is obtained numerically as a solution to a Dirichlet problem for the cube. The geometry chosen for this study is a low-aspect-ratio wing with large sweep, twist, taper, and camber. The topology chosen to wrap the mesh around the wing with good tip resolution is a C-O type mesh. Using this grid, the flow around the wing was computed for a free-stream Mach number of 0.82 at an angle of attack of 5 deg. At this Mach number, an oblique shock forms on the upper surface of the wing, and a tip vortex and three-dimensional flow separation off the wing surface are observed. Particle path lines indicate that the three-dimensional flow separation on the wing surface is part of the roots of the tip-vortex formation. The lifting of the tip vortex before the wing trailing edge is clearly observed by following the trajectory of particles released around the wing tip.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0522
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  • 216
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    Prediction of the vortex wake for noncircular missiles in supersonic flow (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Engineering prediction methods with the capability to calculate induced effects of lee-side separation vorticity associated with circular and noncircular missiles at high angles of attack in supersonic flow are compared. Methods of interest include a discrete vortex cloud technique, concentrated vortex models, and solutions of Euler's equations with specified separation. Comparison of measured and predicted surface pressure distributions and flow field surveys are presented for bodies with circular and elliptic cross sections. Two flow models for computing lee-side vortex-induced effects on control fins in the vicinity of the vortex field are examined, and suggestions regarding the appropriate flow model for specific situations are included.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0504
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  • 217
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    Fine-cut bluntness effects on the hypersonic static stability of a 10 degree cone (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A systematic wind-tunnel investigation has been performed on a series of spherically blunted 10-degree cones to determine Mach number and angle-of-attack effects on the hypersonic static stability. The cones, which range in nose-to-base radius ratios from 0 to 0.5, have been tested from -4 to 20 degrees angle of attack at Mach numbers of 6 and 10 in air and 20 in helium. Relatively large excursions in the center-of-pressure location were observed for small changes in cone nose bluntness. The movement of the center of pressure was also noted to become more pronounced as Mach number increased; increases in angle of attack tended to lessen the magnitude of these excursions. Computational predictions of basic aerodynamic coefficients as well as the center-of-pressure locations using both viscous and inviscid theory compare very well with the experimental data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0503
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  • 218
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    Wake profile measurements of fixed and oscillating flaps (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Although the potential of laser velocimetry for the non-intrusive measurement of complex shear flows has long been recognized, there have been few applications in other small, closely controlled laboratory situations. Measurements in large scale, high speed wind tunnels are still a complex task. To support a study of periodic flows produced by an oscillating edge flap in the Ames eleven foot wind tunnel, this study was done. The potential for laser velocimeter measurements in large scale production facilities are evaluated. The results with hot wire flow field measurements are compared.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166602 , NAS 1.26:166602
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  • 219
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    One-dimensional unsteady modeling of supersonic inlet unstart/restart (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A quasi-one-dimensional unsteady inviscid analysis of mixed-compression supersonic inlet flow is presented with emphasis on modeling of inlet unstart/restart phenomena. Numerical solution of the governing equations of motion is performed using a computationally efficient shock-capturing split-characteristics algorithm. Inlet unstart is modeled using a mass balance method which relates the expelled normal shock position ahead of the inlet cowl to the amount of spilled mass flow over the inlet housing. Comparison of computed results with experimental data for an axisymmetric inlet at a free-stream Mach number of 2.50 shows quite reasonable agreement over an entire unstart/restart transient which includes centerbody translation and retraction as well as bypass mass flow variations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0439
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  • 220
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    Optimization and application of riblets for turbulent drag reduction (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Riblet surfaces have been tested in boundary layers having different upstream histories and at higher Reynolds numbers than previously reported. The drag reduction for the riblet surfaces was found to be dependent on the height and spacing of the riblets in law-of-the-wall variables regardless of the free-stream Reynolds number or upstream boundary-layer history. Micro-photographs of the actual riblet geometries are examined to determine the effect of rib details on the riblet drag-reduction performances. To further increase drag-reduction performance, riblet surfaces are combined with another drag-reduction concept, the large-eddy breakup device (LEBU). In addition, the yaw sensitivity of riblets is evaluated, as well as the characteristics of riblet surfaces manufactured out of a thin vinyl sheet.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0347
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  • 221
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    Trends in computational capabilities for fluid dynamics (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Milestones in the development of computational aerodynamics are reviewed together with past, present, and future computer performance (speed and memory) trends. Factors influencing computer performance requirements for both steady and unsteady flow simulations are identified. Estimates of computer speed and memory that are required to calculate both inviscid and viscous, steady and unsteady flows about airfoils, wings, and simple wing body configurations are presented and compared to computer performance which is either currently available, or is expected to be available before the end of this decade. Finally, estimates of the amounts of computer time that are required to determine flutter boundaries of airfoils and wings at transonic Mach numbers are presented and discussed.
    Keywords: AERODYNAMICS
    Type: NASA-TM-86012 , A-9860 , NAS 1.15:86012
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  • 222
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    High Reynolds number tests of the cast 10-2/DOA 2 airfoil in the Langley 0.3-meter transonic cryogenic tunnel, phase 2 (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Wind tunnel tests of an advanced technology airfoil, the CAST 10-2/DOA 2, were conducted in the Langley 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT). This was the third of a series of tests conducted in a cooperative airfoil research program between the National Aeronautics and Space Administration and the Deutsche Forschungsund Versuchsanstalt fur Luft- und Raumfahrt e. V. For these tests, temperature was varied from 270 K to 110 K at pressures from 1.5 to 5.75 atmospheres. Mach number was varied from 0.60 to 0.80, and the Reynolds number (based on airfoil chord) was varied from 2 to 20 million. The aerodynamic data for the 7.62 cm chord airfoil model used in these tests is presented without analysis. Descriptions of the 0.3-m TCT, the airfoil model, the test instrumentation, and the testing procedures are included.
    Keywords: AERODYNAMICS
    Type: NASA-TM-86273 , NAS 1.15:86273
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  • 223
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    Aerodynamic design of the contoured wind-tunnel liner for the NASA supercritical, laminar-flow-control, swept-wing experiment (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An overview is presented of the entire procedure developed for the aerodynamic design of the contoured wind tunnel liner for the NASA supercritical, laminar flow control (LFC), swept wing experiment. This numerical design procedure is based upon the simple idea of streamlining and incorporates several transonic and boundary layer analysis codes. The liner, presently installed in the Langley 8 Foot Transonic Pressure Tunnel, is about 54 ft long and extends from within the existing contraction cone, through the test section, and into the diffuser. LFC model testing has begun and preliminary results indicate that the liner is performing as intended. The liner design results presented in this paper, however, are examples of the calculated requirements and the hardware implementation of them.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2335 , L-15521 , NAS 1.60:2335
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  • 224
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    Refinement and evaluation of helicopter real-time self-adaptive active vibration controller algorithms (1984)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: A Real-Time Self-Adaptive (RTSA) active vibration controller was used as the framework in developing a computer program for a generic controller that can be used to alleviate helicopter vibration. Based upon on-line identification of system parameters, the generic controller minimizes vibration in the fuselage by closed-loop implementation of higher harmonic control in the main rotor system. The new generic controller incorporates a set of improved algorithms that gives the capability to readily define many different configurations by selecting one of three different controller types (deterministic, cautious, and dual), one of two linear system models (local and global), and one or more of several methods of applying limits on control inputs (external and/or internal limits on higher harmonic pitch amplitude and rate). A helicopter rotor simulation analysis was used to evaluate the algorithms associated with the alternative controller types as applied to the four-bladed H-34 rotor mounted on the NASA Ames Rotor Test Apparatus (RTA) which represents the fuselage. After proper tuning all three controllers provide more effective vibration reduction and converge more quickly and smoothly with smaller control inputs than the initial RTSA controller (deterministic with external pitch-rate limiting). It is demonstrated that internal limiting of the control inputs a significantly improves the overall performance of the deterministic controller.
    Keywords: AERODYNAMICS
    Type: NASA-CR-3821 , NAS 1.26:3821 , R83-956149-16
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  • 225
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    An experimental and theoretical investigation of deposition patterns from an agricultural airplane (1984)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: A flight test program has been conducted with a representative agricultural airplane to provide data for validating a computer program model which predicts aerially applied particle deposition. Test procedures and the data from this test are presented and discussed. The computer program features are summarized, and comparisons of predicted and measured particle deposition are presented. Applications of the computer program for spray pattern improvement are illustrated.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2348 , L-15718 , NAS 1.60:2348
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  • 226
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    Mach 6 flow-field and boundary-layer surveys beneath the forebody of an airbreathing missile (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A wind tunnel investigation has been conducted to determine the flow characteristics of the shock layer beneath the forebody of a hypersonic, airbreathing missile incorporating an aft-mounted inlet. In the inviscid part of the flow field, the measured parameters were in agreement with those predicted by a three-dimensional Method of Characteristics theory. At test conditions matching full scale Mach and Reynolds numbers at an altitude for maximum L/D cruise, boundary layer transition occurred downstream of the inlet face. While this means that a full size vehicle would have a more easily separated boundary layer, it is noted that the Reynolds number discrepancy between wind tunnel and actual flight conditions may result in a movement of the transition point to somewhere upstream of the inlet.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0233
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  • 227
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    Multigrid solution of the Euler equations for aircraft configurations (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A multigrid scheme for solving the Euler equations is presented. The method has been successfully applied to two-dimensional airfoil calculations on both O-type and C-type meshes. In three dimensions the scheme has proved equally effective and calclations of flows over wing/body combinations are possible with convergence achieved in less than 100 cycles.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0093
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    Algebraic grid generation for wing-fuselage bodies (1984)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: An algebraic procedure for the generation of boundary-fitted grids about wing-fuselage configurations is presented. A wing-fuselage configuration is specified by cross sections and mathematically represented by Coons' patches. A configuration is divided into sections so that several grid blocks that either adjoin each other or partially overlap each other can be generated, and each grid has six surfaces that map into a computational cube. Grids are first determined on the six boundary surfaces and then in the interior. Grid curves that are on the surface of the configuration are derived using plane-patch intersections, and single-valued functions relating approximate arc lengths along the curves to computational coordinates define the distribution of grid points. The two-boundary technique and transfinite interpolation are used to determine the boundary surface grids that are not on the configuration, and transfinite interpolation with linear blending functions is used to determine the interior grids.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0002
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  • 229
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    An Analysis of Wave Interactions in Swept-Wing Flows (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Crossflow instabilities dominate disturbance growth in the leading-edge region of swept wings. Streamwise vortices in a boundary layer strongly influence the behavior of other disturbances. Amplification of crossflow vortices near the leading edge produces a residual spanwise nonuniformity in the mid-chord regions where Tollmien-Schlichting (T-S) waves are strongly amplified. Should the T-S wave undergo double-exponential growth because of this effect, the usual transition prediction methods would fail. The crossflow/Tollmien-Schlichting wave interaction was modeled as a secondary instability. The effects of suction are included, and different stability criteria are examined. The results are applied to laminar flow control wings characteristic of energy-efficient aircraft designs.
    Keywords: AERODYNAMICS
    Type: NASA-CR-173723 , NAS 1.26:173723
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  • 230
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    Computation of imaginary-side pressure distributions over the flexible walls of the test section insert for the 0.3-M transonic cryogenic tunnel (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Two dimensional airfoil testing in an adaptive wall test-section wind tunnel requires the computation of the imaginary flow fields extending outward from the top and bottom test section walls. A computer program was developed to compute the flow field which would be associated with an arbitrary test section wall shape. The program is based on incompressible flow theory with a Prandtl-Glauert compressibility correction. The program was validated by comparing the streamline and the pressure field generated by a source in uniform flow with the results from the computer program. A listing of the program, the validation test results, and a sample program are included.
    Keywords: AERODYNAMICS
    Type: NASA-CR-172363 , NAS 1.26:172363
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  • 231
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    Natural laminar flow experiments on modern airplane surfaces (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Flight and wind-tunnel natural laminar flow experiments have been conducted on various lifting and nonlifting surfaces of several airplanes at unit Reynolds numbers between 0.63 x 10 to the 6th power/ft and 3.08 x 10 to the 6th power/ft, at Mach numbers from 0.1 to 0.7, and at lifting surface leading-edge sweep angles from 0 deg to 63 deg. The airplanes tested were selected to provide relatively stiff skin conditions, free from significant roughness and waviness, on smooth modern production-type airframes. The observed transition locations typically occurred downstream of the measured or calculated pressure peak locations for the test conditions involved. No discernible effects on transition due to surface waviness were observed on any of the surfaces tested. None of the measured heights of surface waviness exceeded the empirically predicted allowable surface waviness. Experimental results consistent with spanwise contamination criteria were observed. Large changes in flight-measured performance and stability and control resulted from loss of laminar flow by forced transition. Rain effects on the laminar boundary layer caused stick-fixed nose-down pitch-trim changes in two of the airplanes tested. No effect on transition was observed for flight through low-altitude liquid-phase clouds. These observations indicate the importance of fixed-transition tests as a standard flight testing procedure for modern smooth airframes.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2256 , L-15552 , NAS 1.60:2256
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  • 232
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    Low-speed investigation of effects of wing leading- and trailing-edge flap deflections and canard incidence on a fighter configuration equipped with a forward-swept wing (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An advanced fighter configuration with a forward-swept wing of aspect ratio 3.28 is tested in the Langley 7 by 10 Foot High Speed Tunnel at a Mach number of 0.3. The wing has 29.5 degrees of forward sweep of the quarter chord line and is equipped with 15 percent chord leading edge and 30 percent chord trailing edge flaps. The canard is sweptback 45 degrees. Tests were made through a range of angle of attack from about -2 degrees to 22 degrees. Deflecting the flaps significantly improves the lift drag characteristics at the higher angles of attack. The canard is able to trim the configurations with different flap deflections over most of the range of angle of attack. The penalty in maximum lift coefficient due to trimming is about 0.10.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85795 , L-15746 , NAS 1.15:85795
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  • 233
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    A solution procedure for three-dimensional incompressible Navier-Stokes equation and its application (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An implicit, finite-difference procedure is presented for numerically solving viscous incompressible flows. For convenience of applying the present method to three-dimensional problems, primitive variables, namely the pressure and velocities, are used. One of the major difficulties in solving incompressible flows that use primitive variables is caused by the pressure field solution method which is used as a mapping procedure to obtain a divergence-free velocity field. The present method is designed to accelerate the pressure-field solution procedure. This is achieved by the method of pseudocompressibility in which the time derivative pressure term is introduced into the mass conservation equation. The pressure wave propagation and the spreading of the viscous effect is investigated using simple test problems. The present study clarifies physical and numerical characteristics of the pseudo-compressible approach in simulating incompressible flows. Computed results for external and internal flows are presented to verify the present procedure. The present algorithm has been shown to be very robust and accurate if the selection of the pseudo-compressibility parameter has been made according to the guidelines given.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85978 , A-9801 , NAS 1.15:85978
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  • 234
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    A computational method for viscous incompressible flows (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An implicit, finite-difference procedure for numerically solving viscous incompressible flows is presented. The pressure-field solution is based on the pseudocompressibility method in which a time-derivative pressure term is introduced into the mass-conservation equation to form a set of hyperbolic equations. The pressure-wave propagation and the spreading of the viscous effect is investigated using simple test problems. Computed results for external and internal flows are presented to verify the present method which has proved to be very robust in simulating incompressible flows.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85958 , A-9748 , NAS 1.15:85958
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  • 235
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    Measurement of local connective heat transfer coefficients of four ice accretion shapes (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: In the analytical study of ice accretions that form on aerodynamic surfaces (airfoils, engine inlets, etc.) it is often necessary to be able to calculate convective heat transfer rates. In order to do this, local convective heat transfer coefficients for the ice accretion shapes must be known. In the past, coefficients obtained for circular cylinders were used as an approximation to the actual coefficients since no better information existed. The purpose of this experimental study was to provide local convective heat transfer coefficients for four shapes that represent ice accretions. The shapes were tested with smooth and rough surfaces. The experimental method chosen was the thin-skin heat rate technique. Using this method local Nusselt numbers were determined for the ice shapes. In general it was found that the convective heat transfer was higher in regions where the model's surfaces were convex and lower in regions where the model's surfaces were concave. The effect of roughness was to increase the heat transfer in the high heat transfer regions by approximately 100% while little change was apparent in the low heat transfer regions.
    Keywords: AERODYNAMICS
    Type: NASA-CR-174680 , NAS 1.26:174680
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    Assessment of aerodynamic performance of V/STOL and STOVL fighter aircraft (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The aerodynamic performance of V/STOL and STOVL fighter/attack aircraft was assessed. Aerodynamic and propulsion/airframe integration activities are described and small and large scale research programs are considered. Uncertainties affecting aerodynamic performance that are associated with special configuration features resulting from the V/STOL requirement are addressed. Example uncertainties relate to minimum drag, wave drag, high angle of attack characteristics, and power induced effects.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85937 , A-9690 , NAS 1.15:85937
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  • 237
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    Transonic calculation of airfoil stability and response with active controls (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Transonic aeroelastic stability and response analyses are performed for the MBB A-3 supercritical airfoil. Three degrees of freedom are considered: plunge, pitch, and aileron pitch. The control of airfoil stability and response in transonic flow are studied. Stability analyses are performed using a Pade aeroelastic model based on the use of LTRAN2-NLR transonic small disturbance finite difference computer code. Response analyses are performed by coupling the structural equations of motion to the unsteady aerodynamic forces of LTRAN2-NLR. The focus is on transonic time marching transient response solutions using modal identification to determine stability. Frequency and damping of these modes are directly compared in the complex s-plane with Pade model eigenvalues. Transonic stability and response characteristics of 2-D airfoils are discussed and comparisons are made. Application of the Pade aeroelastic model and time marching analyses to flutter suppression using active controls is demonstrated.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85770 , NAS 1.15:85770
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  • 238
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    Trailing edge flow conditions as a factor in airfoil design (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Some new developments relevant to the design of single-element airfoils using potential flow methods are presented. In particular, the role played by the non-dimensional trailing edge velocity in design is considered and the relationship between the specified value and the resulting airfoil geometry is explored. In addition, the ramifications of the unbounded trailing edge pressure gradients generally present in the potential flow solution of the flow over an airfoil are examined, and the conditions necessary to obtain a class of airfoils having finite trailing edge pressure gradients developed. The incorporation of these conditions into the inverse method of Eppler is presented and the modified scheme employed to generate a number of airfoils for consideration. The detailed viscous analysis of airfoils having finite trailing edge pressure gradients demonstrates a reduction in the strong inviscid-viscid interactions generally present near the trailing edge of an airfoil.
    Keywords: AERODYNAMICS
    Type: NASA-CR-173294 , NAS 1.26:173294 , AAE-84-1 , UILU-ENG-84-0501
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  • 239
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    A laser velocimeter system for large-scale aerodynamic testing (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A unique laser velocimeter was developed that is capable of sensing two orthogonal velocity components from a variable remote distance of 2.6 to 10 m for use in the 40- by 80-Foot and 80- by 120-Foot Wind Tunnels and the Outdoor Aerodynamic Research Facility at Ames Research Center. The system hardware, positioning instrumentation, and data acquisition equipment are described in detail; system capabilities and limitations are discussed; and expressions for systematic and statistical accuracy are developed. Direct and coupled laboratory measurements taken with the system are compared with measurements taken with a laser velocimeter of higher spatial resolution, and sample data taken in the open circuit exhaust flow of a 1/50-scale model of the 80- by 120-Foot Wind Tunnel are presented.
    Keywords: AERODYNAMICS
    Type: NASA-TM-84393 , A-9524 , NAS 1.15:84393
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    Numerical solution of Space Shuttle Orbiter flow field including real gas effects (1984)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The hypersonic, laminar flow around the Space Shuttle Orbiter has been computed for both an ideal gas (gamma = 1.2) and equilibrium air using a real-gas, parabolized Navier-Stokes code. This code employs a generalized coordinate transformation; hence, it places no restrictions on the orientation of the solution surfaces. The initial solution in the nose region was computed using a 3-D, real-gas, time-dependent Navier-Stokes code. The thermodynamic and transport properties of equilibrium air were obtained from either approximate curve fits or a table look-up procedure. Numerical results are presented for flight conditions corresponding to the STS-3 trajectory. The computed surface pressures and convective heating rates are compared with data from the STS-3 flight.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-1747
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  • 241
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    Problems of rate chemistry in the flight regimes of aeroassisted orbital transfer vehicles (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The dissociating and ionizing nonequilibrium flows behind a normal shock wave are calculated for the density and vehicle regimes appropriate for aeroassisted orbital transfer vehicles; the departure of vibrational and electron temperatures from the gas temperature as well as viscous transport phenomena are accounted for. From the thermodynamic properties so determined, radiative power emission is calculated using an existing code. The resulting radiation characteristics are compared with the available experimental data. Chemical parameters are varied to investigate their effect on the radiation characteristics. It is concluded that the current knowledge of rate chemistry leads to a factor-of-4 uncertainty in nonequilibrium radiation intensities. The chemical parameters that must be studied to improve the accuracy are identified.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-1730
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    Interactive phenomena in supersonic jet mixing problems. I Phenomenology and numerical modeling techniques (1984)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The interactive phenomena that occur in supersonic jet mixing flowfields, and numerical modeling techniques developed to analyze such phenomena are discussed. A spatial marching procedure based on solving the parabolized Navier-Stokes jet mixing equations is presented. This procedure combines shock-capturing methodology for the analysis of supersonic mixing regions with pressure-split methodology for the analysis of subsonic mixing regions. The two regions are coupled at viscous sonic lines utilizing a viscous-characteristic coupling procedure. Specialized techniques for the treatment of jet boundary growth, strong discontinuties (Mach disks), and small embedded subsonic zones (behind Mach disks) are presented. Turbulent processes are represented by two-equation turbulence model formulations. In Part II of this article, numerical studies are presented for a variety of supersonic jet interactive phenomena.
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 905-913
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  • 243
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    Effects of small-scale, high intensity inlet turbulence on flow in a two-dimensional diffuser (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The flow through a 2D experimental diffuser with channel width 2.60 cm and divergence angle (2 theta) 9 or 20 deg is investigated experimentally for inlet Reynolds number 78,300 and velocity 43.9 m/s, with and without vertical rods to generate inlet turbulence in excess of the limits defined by Hoffmann (1981) and Hoffmann and Gonzales (1983). Measurements are obtained using a thermal wall-flow-direction probe and a single hot-wire velocity probe, and the results are presented graphically. Significant increases in the pressure-recovery coefficient of the diffuser (10 percent at 9 deg and 22 percent at 20 deg) are attributed to the action of turbulence to reduce distortion and delay separation, thus creating an altered flow condition with symmetrical velocity profiles.
    Keywords: AERODYNAMICS
    Type: ASME, Transactions, Journal of Fluids Engineering (ISSN 0098-2202); 106; 121-124
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  • 244
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    Three-dimensional unsteady Euler equations solution using flux vector splitting (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A method for numerically solving the three-dimensional unsteady Euler equations using flux vector splitting is developed. The equations are cast in curvilinear coordinates and a finite volume discretization is used. An explicit upwind second-order predictor-corrector scheme is used to solve the discretized equations. The scheme is stable for a CFL number of 2 and local time stepping is used to accelerate convergence for steady-state problems. Characteristic variable boundary conditions are developed and used in the far-field and at surfaces. No additional dissipation terms are included in the scheme. Numerical results are compared with results from an existing three-dimensional Euler code and experimental data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-1552
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  • 245
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    Aerothermal environment of a blunted three-dimensional nonaxisymmetric body at Mach 6.8 (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Measurements of surface pressure and heat flux at Mach 6.8 for a nonaxisymmetric body, typical of the forward fuselage of an entry vehicle, have been obtained in the Langley 8-ft high-temperature tunnel. Comparisons of these data with predictions obtained from two numerical methods, which treat the test medium as an ideal gas with a laminar boundary layer, are presented. In general, the numerical results are consistent with each other and follow the data in regions of laminar flow; however, the data for angle of attack 14.9 deg indicate transitional flow on most of the windward surface. Also, the two numerical methods disagree slightly in regions of zero or nearly zero pressure gradients.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-1698
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  • 246
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    Applications of a conservative zonal scheme to transient and geometrically complex problems (1984)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A conservative zoning technique, wherein the flow field for a finite-difference calculation is divided into several regions to simplify grid generation, is discussed and is applied in the solution of a two-dimensional problem of complex topology. Calculations are performed on two zonal, or patched, grid systems for the supersonic flow over a double-airfoil configuration. The solution is smooth and continuous across the zonal interfaces, and shock waves pass through the boundaries without distortion. In addition, the time-accuracy of the zonal-boundary method is verified by a two-zone cyclinder calculation with a stationary inner and a rotating outer mesh.The feasibility of the zonal approach for use in the solution of geometrically complex and unsteady problems is thus demonstrated.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-1532
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  • 247
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    A method for measuring skin friction drag on a flat plate in contaminated gas flows (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A technique for measuring friction drag in turbulent gas and gas/particle flows over flat plates is presented, and preliminary results are reported. A 0.25-in.-thick 72 x 6-in. Al plate is suspended by six horizontal support air bearings and four vertical alignment air bearings between fixed dummy plates and leading-edge and trailing-edge fairings in the 32-in.-high 48-in.-wide 11-ft-long test section of a closed-circuit atmospheric wind tunnel operating at 50-150 ft/sec. Particles of Fe and Al oxides of diameter 20-150 microns and density up to 0.3 lb particles per lb air are injected via a 6 x 0.167-in. nozzle; turbulence is induced by a roughened section of the leading-edge fairing; and friction drag is measured using a load-cell pressure transducer. Sample results are shown in a graph, demonstrating good agreement with theoretical drag calculations.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 0021-8669); 21; 543
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  • 248
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    Complete flowfields over low and wide angle AOTV conceptual configurations (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Flowfields over two conceptual aeroassisted orbital transfer vehicles, a moderate lift-to-drag ratio biconic, and an axisymmetric zero lift aerobrake are investigated. The solution procedures employ a compressible Navier-Stokes time-asymptotic technique for blunt body flow and a Parabolized Navier-Stokes spatial marching technique for the primarily supersonic shock layer flow over low-angle bodies. Emphasis is placed on the laminar convective heating predictions and comparisons to experimental data where available. Code robustness and solution times are discussed and guidelines are given for the specification of various numerical parameters. The Parabolized Navier-Stokes technique has a high probability of solution success for the range of applications considered herein, and very good comparisons with experimental heat-transfer data have been obtained. The blunt body Navier-Stokes code is more sensitive to parameter specification for the aerobrake calculations. Still, coarse grid solutions can be obtained quickly to get a good first cut at flowfield definition. Fair to good comparisons have been obtained between experimental heating and pressure data and coarse and fine grid solutions.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-1695
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  • 249
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    Computation of three-dimensional viscous flows using a space-marching method (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A space-marching method, developed to compute three-dimensional flows for internal geometries, has been utilized to predict viscous flows through a curved duct and over a swept wing. The Navier-Stokes equations have been posed as an initial value problem by neglecting the streamwise viscous diffusion terms and by treating the pressure gradient as a known source term. The resulting equations have been solved by a non-iterative (single pass) algorithm at each streamwise step. The results are compared with earlier computations (based on iterative methods) and the experimental data. The agreement between the present predictions, the experimental data, and the earlier predictions is good for the cases computed. The computation time is only a fraction of the iterative methods.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-1298
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  • 250
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    A fast viscous correction method for unsteady transonic flow about airfoils (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Previously cited in issue 05, p. 583, Accession no. A83-16621
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 746
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  • 251
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    Flow-field investigation of a supercruise fighter model (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A NASA Langley investigation was conducted in the 16-foot Transonic Tunnel to survey the flow field around a model of a Supersonic cruise fighter configuration. In this investigation, a model of a supersonic cruise fighter configuration formerly utilized in afterbody-nozzle performance investigations was surveyed with a single, multiholed probe to determine local values of angle of attack, side flow, and Mach number. The investigation was conducted at Mach numbers of 0.6, 0.9, and 1.2 at angles of attack from 0 to 10 deg. The purpose of the investigation was to provide a data base of experimental data for use in verification of theoretical methods, and to compare the experimental data with predictions from currently available theoretical techniques. Results from this investigation show that local angles of attack were generally greater than free stream above the wing and generally less than free stream below the wing. Also there were large spanwise gradients above the wing at the higher angles of attack. The comparisons of experimental data with theoretical predictions show that the theoretical techniques give a qualitative estimate of the flow-field but will require much work to give good quantitative results.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-1331
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    Numerical simulation of flow through inlets/diffusers with terminal shocks (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Two-dimensional Euler and Navier-Stokes solutions of the flow through three inlet/diffuser configurations with terminal shock systems are reported. Calculations without bleed indicate that the terminal shock location is very sensitive to the outflow back pressure. For cases where there are no available experimental results, it becomes difficult to estimate the back pressure that will result in a terminal shock. Estimates based on quasi-one-dimensional analysis are not found adequate for complex two-dimensional flows. It is found that since the flow downstream of the terminal shock is subsonic, and what happens at the outflow boundary affects the flow inside the inlet, enough of the subsonic diffuser must be modeled to accurately predict the terminal shock region. The diffuser portion should be fairly long with the outflow boundary occurring in a region of more or less uniform flow to be able to prescribe a uniform back pressure. The third configuration studied was investigated with and without incorporating bleed in the code. It is found that the use of bleed stabilizes the shock location and allows solutions which without bleed result in unstarting of the inlet. Comparisons are made with available experimental data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-1362
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  • 253
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    Flow improvements in the NASA Langley 4- by 7-meter tunnel circuit (1984)
    In:  Other Sources
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    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. Previously announced in STAR as N84-15117
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0603
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    Potential flow past axisymmetric bodies at angle of attack (1984)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The Karamcheti (1966) suggestion concerning the use of higher order singularity techniques has been developed for the calculation of incompressible flow past an axisymmetric body at angle of attack. Attention is given to the results of a convergence study using this axial singularity method, where solution accuracy has been investigated for ellipsoids of slenderness ratio in the 1-10 range for both axial and inclined flow. Effects of singularity type, element number and size distribution, and singularity line inset distance, are noted, and a paneling scheme is developed which yields accurate results for the class of axisymmetric bodies having continuous body slopes with discontinuous curvature jumps.
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 0021-8669); 21; 218-220
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    A nonlinear structural concept for drag-reducing compliant walls (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Two mechanisms of drag reduction for flow over flat plates were investigated. The first mechanism employs Bushnell's hypothesis that compliant walls produce drag reduction by interfering with the formation of the turbulent spots in a turbulent boundary layer. It is shown that the amplitudes and frequencies of compliant wall motions for drag reduction might be achieved by using slightly curved walls and the resulting large amplitude motions of snap buckling. A simple structural model of an arch is used in the analysis, and an asymptotic method is developed. The required wall motions can be obtained by using materials like mylar. In addition, the delay of transition from laminar to turbulent flow by driven walls was studied for Poiseuille channel flow. The walls are driven by a periodic traveling wave. A significant increase in the transitional Reynolds number is obtained by appropriately prescribing the wavelength and phase velocity of the wall motion. Previously developed asymptotic methods are used in the analysis. Previously announced in STAR as N83-11061
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 399-402;
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  • 256
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    Lateral aerodynamics of delta wings with leading-edge separation (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Previously cited in issue 19, p. 2972, Accession no. A82-39142
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 323-328
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  • 257
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    Nonlinear computation of wing-body-vertical tail-wake flows at low supersonic speeds (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A numerical method based on the conservation form of the full potential equation has been applied to the problem of three-dimensional supersonic flows with embedded subsonic regions. The governing equation is cast in a nonorthogonal coordinate system, and the theory of characteristics is used to accurately monitor the type-dependent flow field. A conservative switching scheme is employed to transition from the supersonic marching procedure to a subsonic relaxation algorithm and vice versa. The newly developed computer program can handle arbitrary geometries with fuselage, wing, vertical tail and wake components at combined angle of attack and sideslip. Results are presented for a low supersonic Mach numbers flow over the Shuttle orbiter (including the OMS pods and vertical tail), and for flows over a realistic fighter type configuration. Comparisons with experimental data are shown to be in good agreement for various cases.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0427
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  • 258
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    Laser velocimeter measurements of dynamic stall (1984)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: Laser velocimeter measurements were made during the study of a two-dimensional NACA 0012 airfoil undergoing conditions of dynamic stall. The measurements, which were obtained in the Ames 2 foot wind tunnel at reduced frequencies of 0.12 and 1.2, show significant flow field hysteresis around the static stall angle. Comparisons were also made with dual-plate interferograms and good agreement was found for the attached flow cases. For separated flow, characteristic vortex shedding caused poor agreement and significantly increased the measured Reynolds shear stresses.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166603 , NAS 1.26:166603
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    The computation of rotational conical flows (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A method based on the Clebsch velocity decomposition is presented to solve the steady, inviscid, supersonic flow field about arbitrary conical geometries. The system of equations developed, although formally equivalent to the Euler equations, retains the computational efficiency of type-dependent potential flow solutions. Accurate rotational-flow solutions are developed using shock-fitting procedures at the bow and imbedded waves along with special treatment of the vortical layer. Solutions are presented for several high Mach number conical flows and compared with existing Euler solutions and experimental data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0258
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    Analysis of entry aerodynamic heat-transfer data for the orbiter wing lower surface (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Flight-derived aerodynamic heat-transfer data for the orbiter wing lower surface, from STS-2, -3, and -5, are presented and compared with both ground-based experimental results and state-of-the-art computational flowfield results for a nominal angle of attack of 40 degrees. The flight data clearly show the development of the interference heat-transfer region on the wing lower surface resulting from the downstream effects of the bow-shock/wing-shock interaction. The location of the interference heating region is well correlated with a region of minimum static enthalpy near the boundary-layer edge as predicted by a 3-dimensional, inviscid flowfield computation. The magnitude of the interference heat transfer is no greater than the undisturbed laminar heat transfer which occurs during the 'peak aerodynamic heating' portion of entry.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0227
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    Comparison of viscous-shock-layer heating analysis with Shuttle flight data in slip flow regime (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Comparison of STS-2 Shuttle flight heating data along the windward centerline has been made with two-dimensional nonequilibrium viscous shock-layer solutions obtained with shock and wall-slip conditions at an altitude range of 90 to 110 km. The shock slip condition used is the modified Rankine-Hugoniot relations of Cheng as used by Davis, and the wall-slip conditions are based on the first order consideration derived from kinetic theory as given by Scott and Hendricks. The results indicate that the calculated heating distributions with slip boundary conditions agree better with the flight data than those without slip conditions. The agreement improves when the accommodation coefficient or freestream density is decreased to one-half, suggesting the possibility of less than full accommodation for the tile surface and (or) an overestimate of freestream density using the Jacchia-Roberts model. Heating reduction due to the slip effect becomes very pronounced as the flow becomes more rarefied, and the effect is more significant for the stagnation region than the aft region of the vehicle.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0226
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    Direct simulation of transitional flow for hypersonic reentry conditions (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: This paper presents results of flowfield calculations for typical hypersonic reentry conditions encountered by the nose region of the Space Shuttle Orbiter. Most of the transitional flow regime is covered by the altitude range of 150 to 92 km. Calculations were made with the Direct Simulation Monte Carlo (DSMC) method that accounts for translational, rotational, vibrational, and chemical nonequilibrium effects. Comparison of the DSMC heating results with both Shuttle flight data and continuum predictions showed good agreement at the lowest altitude considered. However, as the altitude increased, the continuum predictions, which did not include slip effects, departed rapidly from the DSMC results by overpredicting both heating and drag. The results demonstrate the effects of rarefaction on the shock and the shock layer, along with the extent of the slip and temperature jump at the surface. Also, the sensitivity of the flow structure to the gas-surface interaction model, thermal accommodation, and surface catalysis are studied.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0223
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  • 263
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    Subsonic/transonic prediction capabilities for nozzle/afterbody configurations (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Prediction methods for several nozzle/afterbody flow problems at subsonic and transonic speeds are presented. These methods range from viscous-inviscid interaction methods to solutions for the Navier-Stokes equations in two and three dimensions. The problems addressed are the flow around isolated axisymmetric nozzles, isolated nonaxisymmetric nozzles, and axisymmetric nozzles with empennage. An assessment of the state of development of the methods via comparisons with experimental data is presented.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0192
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  • 264
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    The effect of a short region of concave curvature on a supersonic turbulent boundary layer (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Mean flow measurements have been made in a supersonic turbulent boundary layer experiencing a short region of concave curvature and an adverse pressure gradient. The incoming boundary layer had a Mach number of 2.87 and a unit Reynolds number of 6.3 x 10 to the 7th/m. Flow over constant radii curvatures of delta (0)/R = 0.10 and 0.02 with a fixed turning angle of 8 deg were investigated. Results indicate that both flows remain two-dimensional; length scales increase over their equilibrium values; the mean flow behaves similarly to subsonic flows experiencing concave curvature and lateral divergence; and boundary layer behavior is strongly similar to a corresponding 8 deg ramp flow.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0169
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    An implict LU scheme for the Euler equations applied to arbitrary cascades (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An implicit scheme for solving the Euler equations is derived and demonstrated. The alternating-direction implicit (ADI) technique is modified, using two implicit-operator factors corresponding to lower-block-diagonal (L) or upper-block-diagonal (U) algebraic systems which can be easily inverted. The resulting LU scheme is implemented in finite-volume mode and applied to 2D subsonic and transonic cascade flows with differing degrees of geometric complexity. The results are presented graphically and found to be in good agreement with those of other numerical and analytical approaches. The LU method is also 2.0-3.4 times faster than ADI, suggesting its value in calculating 3D problems.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0167
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    Experimental studies on two dimensional shock boundary layer interactions (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Experiments have been performed on the interaction of oblique shock waves with flat plate boundary layers in the 30.48 cm x 30.48 cm (1 ft. x 1 ft.) supersonic wind tunnel at NASA Lewis Research Center. High accuracy measurements of the plate surface static pressure and shear stress distributions as well as boundary layer velocity profiles were obtained through the interaction region. Documentation was also performed of the tunnel test section flow field and of the two-dimensionality of the interaction regions. The findings provide detailed description of two-dimensional interaction with initially laminar boundary layers over the Mach number range 2.0 to 4.0. Additional information with regard to interactions involving initially transitional boundary layers is presented over the Mach number range 2.0 to 3.0 and those for initially turbulent boundary layers at Mach 2.0. These experiments were directed toward providing well documented information of high accuracy useful as test cases for analytic and numerical calculations. Flow conditions encompassed a Reynolds number range of 4.72E6 to 2.95E7 per meter. The shock boundary layer interaction results were found to be generally in good agreement with the experimental work of previous authors both in terms of direct numerical comparison and in support of correlations establishing laminar separation characteristics.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0099
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    Transonic flow calculations using a flux vector splitting method for the Euler equations (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A study of the flux vector splitting method of Steger and Warming for the solution of the time dependent Euler equations in strong conservation law form for arbitrary two-dimensional geometries is presented. The procedure employed here differs from that of Buning and Steger in that it uses a different algorithm and employs implicit boundary conditions. Moreover, the method, as implemented here, does not contain any explicit smoothing or any adjustable parameters. Calculations were carried out for an NACA 0012 airfoil at various Mach numbers and angles of attack, and cylinders. Steady symmetric solutions were obtained for the full cylinder at a freestream Mach number of .5 without imposing a symmetry condition. In general, good agreement with other methods was obtained.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0090
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    An implicit form for the Osher upwind scheme (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Conservative upwind schemes for the Euler equations, such as the Osher scheme, accurately resolve flow discontinuities and correctly model the physics of the problem. However, these schemes require many more arithmetic operations per integration step than simple central-difference schemes and hence result in large computing times. An implicit version of the first-order- and second-order-accurate Osher schemes in two spatial dimensions and generalized coordinates is developed in this study. Because implicit schemes permit the use of large integration steps, in many cases they require fewer integration steps to reach steady-state (especially in calculations on grids with widely varying mesh-cell sizes). The implicit scheme developed in this study accelerated convergence speeds by almost an order of magnitude in the problems considered. Test cases include quasi-one-dimensional nozzle flow and supersonic flow past a cylinder.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0088
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    The application of vortex theory to the optimum swept propeller (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: It is shown that an optimum propeller generating a swept wake must satisfy the Betz condition, at least to a first order. A numerical solution for swept propellers generating a rigid helicoidal wake is formulated and some results are presented. These results indicate that sweep has a significant effect on Goldstein's kappa factor, particularly at high advance ratios typical of those at which advanced turboprops operate.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-0036
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    Analysis of three-dimensional transonic compressors (1984)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: A method for computing the three-dimensional transonic flow around the blades of a compressor or of a propeller is given. The method is based on the use of the velocity potential, on the hypothesis that the flow is inviscid, irrotational and isentropic. The equation of the potential is solved in a transformed space such that the surface of the blade is mapped into a plane where the periodicity is implicit. This equation is in a nonconservative form and is solved with the help of a finite difference method using artificial time. A computer code is provided and some sample results are given in order to demonstrate the influence of three-dimensional effects and the blade's rotation.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166580 , NAS 1.26:166580
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    Fast solutions to the steady state compressible and incompressible fluid dynamic equations (1984)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: For low speed flows the use of the compressible fluid dynamic equations is inefficient. The use of an explicit scheme requires delta t to be bounded by 1/c. However, the physical parameters change over time scales of order 1/u which is much larger. Hence, it is not appropriate to use explicit schemes for very subsonic flows. Implicit schemes are hard to vectorize and frequently do not converge quickly for very subsonic flows. If one is only interested in the steady state then a minor change to an existing code can greatly increase the efficiency of an explicit method. Even when using an implicit method the proposed changes increase the efficiency of the scheme. The Euler equations for low speed flows will be considered first and then incompressible flows. The method is generalized to include viscous effects. Supersonic flow is accelerated by essentially decoupling the equations.
    Keywords: AERODYNAMICS
    Type: NASA-CR-172416 , ICASE-84-28 , NAS 1.26:172416
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    Bibliography of Supersonic Cruise Research (SCR) program from 1980 to 1983 (1984)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: A bibliography for the Supersonic Cruise Research (SCR) and Variable Cycle Engine (VCE) Programs is presented. An annotated bibliography for the last 123 formal reports and a listing of titles for 44 articles and presentations is included. The studies identifies technologies for producing efficient supersonic commercial jet transports for cruise Mach numbers from 2.0 to 2.7.
    Keywords: AERODYNAMICS
    Type: NASA-RP-1117 , L-15740 , NAS 1.61:1117
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    A numerical study of the controlled flow tunnel for a high lift model (1984)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: A controlled flow tunnel employs active control of flow through the walls of the wind tunnel so that the model is in approximately free air conditions during the test. This improves the wind tunnel test environment, enhancing the validity of the experimentally obtained test data. This concept is applied to a three dimensional jet flapped wing with full span jet flap. It is shown that a special treatment is required for the high energy wake associated with this and other V/STOL models. An iterative numerical scheme is developed to describe the working of an actual controlled flow tunnel and comparisons are shown with other available results. It is shown that control need be exerted over only part of the tunnel walls to closely approximate free air flow conditions. It is concluded that such a tunnel is able to produce a nearly interference free test environment even with a high lift model in the tunnel.
    Keywords: AERODYNAMICS
    Type: NASA-CR-166572 , NAS 1.26:166572
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  • 274
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    Navier-Stokes calculations for the vortex of a rotor in hover (1984)
    In:  CASI
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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: NASA-TM-85894 , A-9611 , NAS 1.15:85894 , USAAVSCOM-TR-84-A-3
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    Aeropropulsive characteristics of nonaxisymmetric-nozzle thrust reversers at Mach numbers from 0 to 1.20 (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the performance of nonaxisymmetric-nozzle thrust reversers installed on a generic twin-engine fighter aircraft model. Test data were obtained at static conditions and at Mach numbers from 0.15 to 1.20 with jet exhaust simulated by high pressure air. Results showed that reverse-thrust levels of greater than 50 percent at static conditions and greater than 30 percent at in-flight conditions could be achieved. Internal reverser-port passage length was found to be very important in improving reverser performance. Increasing the reverser-port passage length improved reverse-thrust performance by as much as 28 percent at static conditions and by as much as 17 percent at Mach 1.20.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2306 , L-15724 , NAS 1.60:2306
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    Physical aspects of computing the flow of a viscous fluid (1984)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: One of the main themes in fluid dynamics at present and in the future is going to be computational fluid dynamics with the primary focus on the determination of drag, flow separation, vortex flows, and unsteady flows. A computation of the flow of a viscous fluid requires an understanding and consideration of the physical aspects of the flow. This is done by identifying the flow regimes and the scales of fluid motion, and the sources of vorticity. Discussions of flow regimes deal with conditions of incompressibility, transitional and turbulent flows, Navier-Stokes and non-Navier-Stokes regimes, shock waves, and strain fields. Discussions of the scales of fluid motion consider transitional and turbulent flows, thin- and slender-shear layers, triple- and four-deck regions, viscous-inviscid interactions, shock waves, strain rates, and temporal scales. In addition, the significance and generation of vorticity are discussed. These physical aspects mainly guide computations of the flow of a viscous fluid.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85893 , A-9650 , NAS 1.15:85893
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    Water-tunnel study of transition flow around circular cylinders (1984)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: The recently reported phenomenon of asymmetric flow separation from a circular cylinder in the critical Reynolds number regime has been confirmed in a water-tunnel experiment. For the first time, an attempt was made to visualize the wake of the cylinder during the transition from subcritical to critical flow and to correlate the visualizations with lift and drag measurements. The occurrence of a dominant asymmetric-flow state was quite repeatable, both when increasing and decreasing the Reynolds number, resulting in a mean lift coefficient of C sub L approx 1.2 and a shift in the angle of the wake by about 12 deg. A distinctive step change in the drag and shedding frequency was also found to occur. A hysteresis was confirmed to exist in this region as the Reynolds number was cycled over the transition range. Both boundaries of the asymmetry appear to be supercritical bifurcations in the flow. The asymmetry was normally steady in the mean; however, there were instances when the direction of the asymmetry reversed and remained so for the duration of the Reynolds number sweep through this transition region. A second asymmetry was observed at a higher Reynolds number; however, the mean lift coefficient was much lower, and the direction of the asymmetry was not observed to reverse. Introducing a small local disturbance into the boundary layer was found to prevent the critical asymmetry from developing along the entire span of the cylinder.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85879 , A-9606 , NAS 1.15:85879 , AVSCOM-TM-84-A-1
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    A recontoured, upper surface designed to increase the maximum lift coefficient of a modified NACA 65 (0.82) (9.9) airfoil section (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A recontoured upper surface was designed to increase the maximum lift coefficient of a modified NACA 65 (0.82)(9.9) airfoil section which was tested at Mach numbers of 0.3 and 0.4 and Reynolds numbers of 2.3x10(6) and 4.3x10(6). The original 6-series section was tested for comparison with the recontoured section. The recontoured profile was found to have a higher maximum lift coefficient at all test conditions than the original airfoil. The recontoured airfoil showed less drag and nearly the same pitching moment characteristics as the original 6-series airfoil at all test conditions. The improvements found for the recontoured airfoil of the present study are similar to those found during previous investigations of recontoured 6-series airfoils with less camber.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85855 , A-9514 , NAS 1.15:85855
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    Investigation of flow phenomena in a transonic fan rotor using laser anemometry (1984)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: Several flow phenomena including flowfield periodicity, rotor shock oscillation, and rotor shock system geometry were investigated in a transonic low aspect ratio fan rotor using laser anemometry. Flow periodicity is found to increase with increasing rotor pressure rise, and to correlate with blade geometry variations. Analysis of time-accurate laser anemometer data indicates that the rotor shock oscillates about its mean location with an amplitude of 3 to 4 percent of rotor chord. The shock surface is nearly two-dimensional or levels of rotor pressure rise at and above the peak efficiency level but becomes more complex for lower levels of pressure rise. Spanwise shock lean generates radial flows due to streamline deflection in the hub-to-shroud streamsurface.
    Keywords: AERODYNAMICS
    Type: NASA-TM-83555 , E-1934 , NAS 1.15:83555
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    Loads and aeroelasticity division research and technology accomplishments for FY 1983 and plans for FY 1984 (1984)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: Research was done in the following areas: development and validation of solution algorithms, modeling techniques, integrated finite elements for flow-thermal-structural analysis and design, optimization of aircraft and spacecraft for the best performance, reduction of loads and increase in the dynamic structural stability of flexible airframes by the use of active control, methods for predicting steady and unsteady aerodynamic loads and aeroelastic characteristics of flight vehicles with emphasis on the transonic range, and methods for predicting and reducing helicoper vibrations.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85740 , NAS 1.15:85740
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    Aerodynamic characteristics of a Sparrow 3 missile model in the flow field of a generalized parent body at Mach 2.86 (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Longitudinal aerodynamic characteristics of a Sparrow 3 wing control missile model were measured through a range of separation distances relative to a flat plate surface that represented the parent-body configuration. Measurements were obtained with and without two dimensional circular arc protuberances attached to the flat plate surface. The tests were conducted at a Mach number of 2.86 and a Reynolds number per meter of 6.56 million. The behavior of these longitudinal characteristics with varying separation distance in the flow field created by the flat plate and protuberance was generally as would be expected on the basis of flow field boundaries determined from the second order approximation of Friedrich. In general, varying roll angle from 0 deg to 45 deg caused no significant effect on the store separation characteristics.
    Keywords: AERODYNAMICS
    Type: NASA-TM-85713 , L-15705 , NAS 1.15:85713
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    Wave drag as the objective function in transonic fighter wing optimization (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The original computational method for determining wave drag in a three dimensional transonic analysis method was replaced by a wave drag formula based on the loss in momentum across an isentropic shock. This formula was used as the objective function in a numerical optimization procedure to reduce the wave drag of a fighter wing at transonic maneuver conditions. The optimization procedure minimized wave drag through modifications to the wing section contours defined by a wing profile shape function. A significant reduction in wave drag was achieved while maintaining a high lift coefficient. Comparisons of the pressure distributions for the initial and optimized wing geometries showed significant reductions in the leading-edge peaks and shock strength across the span.
    Keywords: AERODYNAMICS
    Type: NASA-TP-2265 , L-15687 , NAS 1.60:2265
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    A linearized unsteady aerodynamic analysis for transonic cascades (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The unsteady airloads generated by the vibrations of turbomachine blades operating at transonic Mach numbers are predicted by a linearized potential flow analysis whose unsteady aerodynamics model encompasses the effects of blade geometry, nonzero mean pressure variation across the blade row, high frequency blade motion, and shock motion, all within the framework of a linearized frequency-domain formulation. A numerical solution for the entire unsteady flow field is determined by matching a solution covering an extended blade passage region to another covering, and extending beyond, the supersonic region(s) adjacent to a blade surface. Results are given for cascades of double circular arc and flat plate blades, in order to demonstrate the unsteady analysis and to partially illustrate the effects of blade geometry, inlet Mach number, vibration frequency and shock motion on unsteady response.
    Keywords: AERODYNAMICS
    Type: Journal of Fluid Mechanics (ISSN 0022-1120); 149; 403-429
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  • 284
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    Acoustic pressures emanating from a turbomachine stage (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A knowledge of the acoustic energy emission of each blade row of a turbomachine is useful for estimating the overall noise level of the machine and for determining its discrete frequency noise content. Because of the close spacing between the rotor and stator of a compressor stage, the strong aerodynamic interactions between them have to be included in obtaining the resultant flow field. This paper outlines a three-dimensional theory for determining the discrete frequency noise content of an axial compressor consisting of a rotor and a stator each with a finite number of blades. The lifting surface theory and the linearized equation of an ideal, nonsteady compressible fluid motion are used for thin blades of arbitrary cross section. The combined pressure field at a point of the fluid is constructed by linear addition of the rotor and stator solutions together with an interference factor obtained by matching them for net zero vorticity behind the stage.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-2325
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    Two dimensional blade-vortex interaction flow visualization study (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Blade-vortex interaction occurs when a rotor blade encounters the tip vortex from a previous rotor blade. To obtain details of the close encounter process, the results from a flow visualization study of an airfoil representing a rotor blade in the wake of an oscillating airfoil serving as a vortex generator are described. A distinguishing feature of this study is that the vortex filament is oriented parallel to the blade span, orthogonal to the test section free stream velocity. This orientation simulates the case of two-dimensional blade-vortex interaction, which is known to produce the most impulsive and most intensive BVI noise. Photographic data are examined to deduce qualitative and quantitative details of the close encounter interaction process with emphasis on structural changes in the vortex filament and its trajectory.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-2307
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  • 286
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    Development of a rotor wake-vortex model, volume 1 (1984)
    In:  CASI
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    Publication Date: 2019-06-28
    Description: Certain empirical rotor wake and turbulence relationships were developed using existing low speed rotor wave data. A tip vortex model was developed by replacing the annulus wall with a row of image vortices. An axisymmetric turbulence spectrum model, developed in the context of rotor inflow turbulence, was adapted to predicting the turbulence spectrum of the stator gust upwash.
    Keywords: AERODYNAMICS
    Type: NASA-CR-174849 , NAS 1.26:174849
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    Airfoil trailing edge flow measurements and comparison with theory, incorporating open wind tunnel corrections (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: Trailing edge data for boundary layer-near wake thickness parameters are given for airfoils and flat plates. Reynolds number effects are examined as a function of model size, velocity and boundary layer tripping. These data expand that presented previously by the authors particularly for airfoil non-zero angles of attack. Comparisons are made here with boundary layer calculations using potential flow modeling and a well documented two-dimensional finite-difference method for laminar and turbulent boundary layers. Open wind tunnel corrections to angle of attack and camber are developed and are incorporated in the potential flow modeling to assure correct comparisons for non-zero angles of attack. It was found that although the open tunnel flow turbulence affected boundary layer transition for the higher velocities the theory successfully 'brackets' the data. Comparisons demonstrate the degree of accuracy one might expect for the prediction of boundary layer thickness parameters when given only geometry and nominal flow conditions as input to boundary layer codes.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-2266
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  • 288
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    Feedback in separated flows over symmetric airfoils (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: For a flow over an airfoil with laminar separation, a feedback cycle may exist whereby a Kelvin-Helmholtz instability wave emanating from the separation point on the airfoil surface grows along the shear layer and is diffracted as it interacts with the sharp trailing edge of the airfoil, causing acoustic radiation which , in turn, propagates upstream and regenerates the initial instability wave. The analysis is restricted to the high frequency limit. Solutions to the boundary-value problem are obtained using the slowly varying approximation and the method of matched asymptotic expansions. Resonant solutions exist for certain discrete values of the Reynolds and Strouhal numbers. The results are discussed and compared with available data.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-2297
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  • 289
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    Applicability of a panel method, which includes nonlinear effects, to a forward-swept-wing aircraft (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The ability of a lower order panel method VSAERO, to accurately predict the lift and pitching moment of a complete forward-swept-wing/canard configuration was investigated. The program can simulate nonlinear effects including boundary-layer displacement thickness, wake roll up, and to a limited extent, separated wakes. The predictions were compared with experimental data obtained using a small-scale model in the 7- by 10- Foot Wind Tunnel at NASA Ames Research Center. For the particular configuration under investigation, wake roll up had only a small effect on the force and moment predictions. The effect of the displacement thickness modeling was to reduce the lift curve slope slightly, thus bringing the predicted lift into good agreement with the measured value. Pitching moment predictions were also improved by the boundary-layer simulation. The separation modeling was found to be sensitive to user inputs, but appears to give a reasonable representation of a separated wake. In general, the nonlinear capabilities of the code were found to improve the agreement with experimental data. The usefullness of the code would be enhanced by improving the reliability of the separated wake modeling and by the addition of a leading edge separation model.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-2402
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  • 290
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    Application of the implicit MacCormack scheme to the parabolized Navier-Stokes equations (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 1755-176
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  • 291
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    Analysis of airfoil leading-edge separation bubbles (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: A local inviscid-viscous interaction technique was developed for the analysis of low speed airfoil leading edge transitional separation bubbles. In this analysis an inverse boundary layer finite difference analysis is solved iteratively with a Cauchy integral representation of the inviscid flow which is assumed to be a linear perturbation to a known global viscous airfoil analysis. Favorable comparisons with data indicate the overall validity of the present localized interaction approach. In addition numerical tests were performed to test the sensitivity of the computed results to the mesh size, limits on the Cauchy integral, and the location of the transition region.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 83-0300 , AIAA Journal (ISSN 0001-1452); 22; 1697-170
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  • 292
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    Effect of initial conditions on turbulent reattachment downstream of a backward-facing step (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The reattachment of a fully turbulent, two-dimentional shear layer downstream of a backward-facing step has been studied experimentally. The work examines the effect of variations in inlet conditions on the process of reattachment. A series of experiments was conducted in a low-speed wind tunnel using specialized instrumentation suited to the highly turbulent reversing flow near reattachment. Accurate characterization of the time-mean features of the reattaching flows was possible. Assuming linear scaling normalized on distance from reattachment, distributions of normalized pressure coefficient and forward flow fraction, and time-averaged skin friction coefficient appear universal for two-dimensional reattachment, independent of initial conditions and step height, for given duct geometry (area ratio) and for high step-height Reyolds numbers with thin separating boundary layers. The results suggest universal flow structure in the reattachment zone.
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 1727-173
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    Numerical simulation of the transonic flowfield for wing/nacelle configurations (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: An efficient grid-interfacing zonal algorithm has been developed for computing the three-dimensional transonic flow field about wing/nacelle multicomponent configurations. The algorithm uses the full-potential formulation and the AF2 fully-implicit approximate factorization scheme. The flow field position is computed using a component-adaptive grid approach in which separate grids are employed for the individual components in the multicomponent 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 paper presents 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 multicomponent flow field solution. Computed results and correlations with experiment are presented to illustrate application of the analysis.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-2430
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  • 294
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    Calculation of unsteady fan rotor response caused by downstream flow distortions (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The present model for fan rotor/support strut airfoil interaction uses a time-marching code for the rotor flow, coupled with a potential flow model for the stator-strut region. Study of the effect of strut design variables indicates that rotor flow disturbance is increased by the primary variables of larger strut thickness and circumferential spacing, while decreasing exponentially with increased rotor-strut separation. The time-marching code predicts local rotor pressure and flow perturbations in response to an unsteady downstream boundary condition.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-2282
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    On the stability of an infinite swept attachment line boundary layer (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The instability of an infinite swept attachment line boundary layer is considered in the linear regime. The basic three dimensional flow is shown to be susceptible to travelling wave disturbances which propagate along the attachment line. The effect of suction on the instability is discussed and the results suggest that the attachment in boundary layer on a swept wing can be significantly stabilized by extremely small amounts of suction. The results obtained are in excellent agreement with the available experimental observations.
    Keywords: AERODYNAMICS
    Type: Proceedings, Series A - Mathematical and Physical Sciences (ISSN 0080-4630); 395; 1809; 229-245
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    Design parameters for flow energizers (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: One of the most significant uses of flow energizers, which are small highly swept strakes mounted immediately above a lifting surface, is in flow control over regions where a lifting surface is joined to another body, such as a fuselage or nacelle. In the presently reported systematic wind tunnel test study of flow energizers, 14 different geometric configurations using a 75-deg sweep flow energizer were tested on a light twin-engine general aviation aircraft model. It is found that cambered flow energizers perform better than their flat counterparts. All but two of the energizer installations developed lower L/D at cruise angles of attack, lower maximum lift coefficients, and lower stall angles of attack than the baseline model.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-2499
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    The aerodynamic characteristics of a propulsive wing/canard concept in STOL (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Description: The potential of the propulsive wing in developing very high lift coefficients for STOL operation has been investigated with several nozzle aspect ratios. The use of the propulsive wing/canard appears to offer an approach to managing the large negative pitching moments associated with trailing-edge blowing. A full-span model of a wing/canard concept representing a fighter configuration has been tested at STOL conditions in the Langley 4 by 7 Meter Tunnel. The results of this test are presented, and comparisons are made to previous tests of the same configuration tested as a semispan model (Stewrt, 1983). Also presented are data showing the effects of large flap deflection and the effect of nozzle span. Comparisons of the test results with jet-flap theory are made and indicate good agreement.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 84-2396
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  • 298
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    Wavelength selection and growth of Goertler vortices (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Keywords: AERODYNAMICS
    Type: AIAA Journal (ISSN 0001-1452); 22; 1529-153
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  • 299
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    Computational transonic analysis of canted winglets (1984)
    In:  Other Sources
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    Publication Date: 2019-06-28
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft (ISSN 0021-8669); 21; 873-878
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  • 300
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    Investigation to optimize the passive shock wave/boundary layer control for supercritical airfoil drag reduction (1984)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The passive shock wave/boundary layer control for reducing the drag of 14%-thick supercritical airfoil was investigated in the 3 in. x 15.4 in. RPI Transonic Wind Tunnel with and without the top wall insert at transonic Mach numbers. Top wall insert was installed to increase the flow Mach number to 0.90 with the model mounted on the test section bottom wall. Various porous surfaces with a cavity underneath were positioned on the area of the airfoil where the shock wave occurs. The higher pressure behind the shock wave circulates flow through the cavity to the lower pressure ahead of the shock wave. The effects from this circulation prevent boundary layer separation and enthropy increase hrough the shock wave. The static pressure distributions over the airfoil, the wake impact pressure survey for determining the profile drag and the Schlieren photographs for porous surfaces are presented and compared with the results for solid surface airfoil. With a 2.8% uniform porosity the normal shock wave for the solid surface was changed to a lambda shock wave, and the wake impact pressure data indicate a drag coefficient reduction as much as 45% lower than for the solid surface airfoil at high transonic Mach numbers.
    Keywords: AERODYNAMICS
    Type: NASA-CR-175788 , NAS 1.26:175788
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