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  • AERODYNAMICS  (12,790)
  • 1
    Publication Date: 2019-06-28
    Description: Wind tunnel tests have been conducted on an NACA 2412 airfoil section at Reynolds number of 2.2 x 10(exp 6) and Mach number of 0.13. Detailed measurements of flow fields associated with turbulent boundary layers have been obtained at angles of attack of 12.4 degrees, 14.4 degrees, and 16.4 degrees. Pre- and post-separated velocity and pressure survey results over the airfoil and in the associated wake are presented. Extensive force, pressure, tuft survey, hot-film survey, local skin friction, and boundary layer data are also included. Pressure distributions and separation point locations show good agreement with theory for the two layer angles of attack. Boundary layer displacement thickness, momentum thickness, and shape factor agree well with theory up to the point of separation. There is considerable disparity between extent of flow reversal in the wake as measured by pressure and hot-film probes. The difference is attributed to the intermittent nature of the flow reversal.
    Keywords: AERODYNAMICS
    Type: NASA-CR-197497 , NAS 1.26:197497 , AR77-3
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  • 2
    Publication Date: 2019-07-13
    Description: The implementation of a two-equation k-omega turbulence model into the NPARC flow solver is described. Motivation for the selection of this model is given, major code modifications are outlined, new imputs to the code are described, and results are presented for several validation cases: an incompressible flow over a smooth flat plate, a subsonic diffuser flow, and a shock-induced separated flow. Comparison of results with the k-epsilon model indicate that the k-omega model predicts simple flows equally well whereas, for adverse pressure gradient flows, the k-omega model outperforms the other turbulence models in NPARC.
    Keywords: AERODYNAMICS
    Type: NASA-TM-107080 , NAS 1.15:107080 , E-9955 , AIAA PAPER 96-0383 , NIPS-96-08118 , Aerospace Sciences Meeting and Exhibit; Jan 15, 1996 - Jan 18, 1996; Reno, NV; United States
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  • 3
    Publication Date: 2019-07-13
    Description: An approach for solving the compressible Euler and Navier-Stokes equations upon meshes composed of nearly arbitrary polyhedra is described. Each polyhedron is constructed from an arbitrary number of triangular and quadrilateral face elements, allowing the unified treatment of tetrahedral, prismatic, pyramidal, and hexahedral cells, as well the general cut cells produced by Cartesian mesh approaches. The basics behind the numerical approach and the resulting data structures are described. The accuracy of the mixed volume grid approach is assessed by performing a grid refinement study upon a series of hexahedral, tetrahedral, prismatic, and Cartesian meshes for an analytic inviscid problem. A series of laminar validation cases are made, comparing the results upon differing grid topologies to each other, to theory, and experimental data. A computation upon a prismatic/tetrahedral mesh is made simulating the laminar flow over a wall/cylinder combination.
    Keywords: AERODYNAMICS
    Type: NASA-TM-107135 , NAS 1.15:107135 , AIAA PAPER 96-0762 , E-10065 , NIPS-96-07909 , Aerospace Sciences Meeting and Exhibit; Jan 15, 1996 - Jan 18, 1996; Reno, NV; United States
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  • 4
    Publication Date: 2011-08-24
    Description: Measurements of wing buffeting, using root strain gages, were made in the NASA Langley 0.3 m cryogenic wind tunnel to refine techniques which will be used in larger cryogenic facilities such as the United States National Transonic Facility (NTF) and the European Transonic Wind Tunnel (ETW). The questions addressed included the relative importance variations in frequency parameter and Reynolds number, the choice of model material (considering both stiffness and damping) and the effects of static aeroelastic distortion. The main series of tests was made on three half models of slender 65 deg delta wings with a sharp leading edge. The three delta wings had the same planform but widely differing bending stiffnesses and frequencies (obtained by varying both the material and the thickness of the wings). It was known that the steady flow on this configuration would be insensitive to variations in Reynolds number. On this wing at vortex breakdown the spectrum of the unsteady excitation is unusual, having a sharp peak at particular frequency parameter. Additional tests were made on one unswept half-wing of aspect ratio 1.5 with an NPL 9510 aerofoil section, known to be sensitive to variations in Reynolds number at transonic speeds. The test Mach numbers were M = 0.21 and 0.35 for the delta wings and to M = 0.30 for the unswept wing. On this wing the unsteady excitation spectrum is fairly flat (as on most wings). Hence correct representation of the frequency parameter is not particularly important.
    Keywords: AERODYNAMICS
    Type: Aeronautical Journal (ISSN 0001-9240); 99; 981; p. 1-14
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  • 5
    Publication Date: 2013-08-31
    Description: Analytical investigation of dynamic stall on HAWT (horizontal-axis wind turbines) rotor loads was conducted. Dynamic stall was modeled using the Gormont approach on the MOD-2 rotor, treating the blade as a rigid body teetering about a fixed axis. Blade flapwise bending moments at station 370 were determined with and without dynamic stall for spatial variations in local wind speed due to wind shear and yaw. The predicted mean flapwise bending moments were found to be in good agreement with test results. Results obtained with and without dynamic stall showed no significant difference for the mean flapwise bending moment. The cyclic bending moments calculated with and without dynamic stall effects were substantially the same. None of the calculated cyclic loads reached the level of the cyclic loads measured on the MOD-2 using the Boeing five-minute-average technique.
    Keywords: AERODYNAMICS
    Type: DASCON Engineering, Collected Papers on Wind Turbine Technology; p 41-46
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  • 6
    Publication Date: 2013-08-31
    Description: A coordinated effort has been underway over the past four years to elevate unstructured-grid methodology to a mature level. The goal of this endeavor is to provide a validated capability to non-expert users for performing rapid aerodynamic analysis and design of complex configurations. The Euler component of the system is well developed, and is impacting a broad spectrum of engineering needs with capabilities such as rapid grid generation and inviscid flow analysis, inverse design, interactive boundary layers, and propulsion effects. Progress is also being made in the more tenuous Navier-Stokes component of the system. A robust grid generator is under development for constructing quality thin-layer tetrahedral grids, along with a companion Navier-Stokes flow solver. This paper presents an overview of this effort, along with a perspective on the present and future status of the methodology.
    Keywords: AERODYNAMICS
    Type: NASA. Lewis Research Center, Surface Modeling, Grid Generation, and Related Issues in Computational Fluid Dynamic (CFD) Solutions; p 289-308
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  • 7
    facet.materialart.
    Unknown
    In:  Other Sources
    Publication Date: 2019-01-25
    Description: Grid generation plays an integral part in the solution of computational fluid dynamics problems for aerodynamics applications. A major difficulty with standard structured grid generation, which produces quadrilateral (or hexahedral) elements with implicit connectivity, has been the requirement for a great deal of human intervention in developing grids around complex configurations. This has led to investigations into unstructured grids with explicit connectivities, which are primarily composed of triangular (or tetrahedral) elements, although other subdivisions of convex cells may be used. The existence of large gradients in the solution of aerodynamic problems may be exploited to reduce the computational effort by using high aspect ratio elements in high gradient regions. However, the heuristic approaches currently in use do not adequately address this need for high aspect ratio unstructured grids. High aspect ratio triangulations very often produce the large angles that are to be avoided. Point generation techniques based on contour or front generation are judged to be the most promising in terms of being able to handle complicated multiple body objects, with this technique lending itself well to adaptivity. The eventual goal encompasses several phases: first, a partitioning phase, in which the Voronoi diagram of a set of points and line segments (the input set) will be generated to partition the input domain; second, a contour generation phase in which body-conforming contours are used to subdivide the partition further as well as introduce the foundation for aspect ratio control, and; third, a Steiner triangulation phase in which points are added to the partition to enable triangulation while controlling angle bounds and aspect ratio. This provides a combination of the advancing front/contour techniques and refinement. By using a front, aspect ratio can be better controlled. By using refinement, bounds on angles can be maintained, while attempting to minimize the number of Steiner points.
    Keywords: AERODYNAMICS
    Type: NASA. Lewis Research Center, Surface Modeling, Grid Generation, and Related Issues in Computational Fluid Dynamic (CFD) Solutions; p 88
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  • 8
    Publication Date: 2019-06-28
    Description: A direct numerical simulation (DNS) algorithm has been developed and validated for use in the investigation of crossflow instability on supersonic swept wings, an application of potential relevance to the design of the High-Speed Civil Transport (HSCT). The algorithm is applied to the investigation of stationary crossflow instability on an infinitely long 77-degree swept wing in Mach 3.5 flow. The results of the DNS are compared with the predictions of linear parabolized stability equation (PSE) methodology. In-general, the DNS and PSE results agree closely in terms of modal growth rate, structure, and orientation angle. Although further validation is needed for large-amplitude (nonlinear) disturbances, the close agreement between independently derived methods offers preliminary validation of both DNS and PSE approaches.
    Keywords: AERODYNAMICS
    Type: NASA-CR-198267 , NAS 1.26:198267 , NIPS-96-08486
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  • 9
    Publication Date: 2019-06-28
    Description: Two F-18 aircraft were flown, one above the other, in two formations, in order for the shock systems of the two aircraft to merge and propagate to the ground. The first formation had the canopy of the lower F-18 in the inlet shock of the upper F-18 (called inlet-canopy). The flight conditions were Mach 1.22 and an altitude of 23,500 ft. An array of five sonic boom recorders was used on the ground to record the sonic boom signatures. This paper describes the flight test technique and the ground level sonic boom signatures. The tail-canopy formation resulted in two, separated, N-wave signatures. Such signatures probably resulted from aircraft positioning error. The inlet-canopy formation yielded a single modified signature; two recorders measured an approximate flattop signature. Loudness calculations indicated that the single inlet-canopy signatures were quieter than the two, separated tail-canopy signatures. Significant loudness occurs after a sonic boom signature. Such loudness probably comes from the aircraft engines.
    Keywords: AERODYNAMICS
    Type: NASA-TM-104312 , H-2067 , NAS 1.15:104312
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  • 10
    Publication Date: 2019-06-28
    Description: A multiblock, discrete sensitivity analysis method is used to couple a direct optimization method and a flow analysis method. The domain is divided into smaller subdomains for which the sensitivities are obtained separately. Then, an effective sensitivity equation is solved to complete the coupling of all the sensitivity information. The flow analysis is based on the thin-layer Navier-Stokes equations solved by an implicit, upwind-biased, finite-volume method. The method of feasible directions is used for the present gradient-based optimization approach. First, a transonic airfoil is optimized to investigate the behavior of the method in highly nonlinear flows as well as the effect of different blocking strategies on the procedure. A supercritical airfoil is produced from an initially symmetric airfoil with multiblocking affecting the path but not the final shape. Secondly, a two-element airfoil is shape optimized in subsonic flow to demonstrate the present method's capability of shaping aerodynamically interfering elements simultaneously. For a very low and a very high Reynolds number cases, the shape of the main airfoil and the flap are optimized to yield improved lift-to-drag ratios.
    Keywords: AERODYNAMICS
    Type: NASA-CR-199785 , NAS 1.26:199785 , AIAA PAPER 94-4273 , NIPS-95-06444
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  • 11
    Publication Date: 2019-06-28
    Description: This report summarizes some NASA Lewis (i.e., government owned) computer codes capable of being used for airbreathing propulsion system studies to determine the design geometry and to predict the design/off-design performance of compressors and turbines. These are not CFD codes; velocity-diagram energy and continuity computations are performed fore and aft of the blade rows using meanline, spanline, or streamline analyses. Losses are provided by empirical methods. Both axial-flow and radial-flow configurations are included.
    Keywords: AERODYNAMICS
    Type: NASA-CR-198433 , NAS 1.26:198433 , E-10041 , NIPS-95-06493
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  • 12
    Publication Date: 2019-06-28
    Description: A far-wing theory in which the validity of the detailed balance principle is maintained in each step of the derivation is presented. The role of the total density matrix including the initial correlations is analyzed rigorously. By factoring out the rapidly varying terms in the complex-time development operator in the interaction representation, better approximate expressions can be obtained. As a result, the spectral density can be expressed in terms of the line-coupling functions in which two coupled lines are arranged symmetrically and whose frequency detunings are omega - 1/2(omega(sub ji) + omega (sub j'i'). Using the approximate values omega - omega(sub ji) results in expressions that do not satisfy the detailed balance principle. However, this principle remains satisfied for the symmetrized spectral density in which not only the coupled lines are arranged symmetrically, but also the initial and final states belonging to the same lines are arranged symmetrically as well.
    Keywords: AERODYNAMICS
    Type: NASA-TM-111075 , NAS 1.15:111075
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  • 13
    Publication Date: 2019-06-28
    Description: To identify planform characteristics which have promise for a highly maneuverable vehicle, an investigation was conducted in the Langley Subsonic Basic Research Tunnel to determine the low-speed longitudinal aerodynamics of 21 planform geometries. Concepts studied included twin bodies, double wings, cutout wings, and serrated forebodies. The planform models tested were all 1/4-in.-thick flat plates with beveled edges on the lower surface to ensure uniform flow separation at angle of attack. A 1.0-in.-diameter cylindrical metric body with a hemispherical nose was used to house the six-component strain gauge balance for each configuration. Aerodynamic force and moment data were obtained across an angle-of-attack range of 0 to 70 deg with zero sideslip at a free-stream dynamic pressure of 30 psf. Surface flow visualization studies were also conducted on selected configurations using fluorescent minitufts. Results from the investigation indicate that a cutout wing planform can improve lift characteristics; however, cutout size, shape, and position and wing leading-edge sweep will all influence the effectiveness of the cutout configuration. Tests of serrated forebodies identified this concept as an extremely effective means of improving configuration lift characteristics; increases of up to 25 percent in the value of maximum lift coefficient were obtained.
    Keywords: AERODYNAMICS
    Type: NASA-TP-3503 , L-17301 , NAS 1.60:3503
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  • 14
    Publication Date: 2019-06-28
    Description: The computational fluid dynamics code, PARC3D, is tested to see if its use of non-physical artificial dissipation affects the accuracy of its results. This is accomplished by simulating a shock-laminar boundary layer interaction and several hypersonic flight conditions of the Pegasus(TM) launch vehicle using full artificial dissipation, low artificial dissipation, and the Engquist filter. Before the filter is applied to the PARC3D code, it is validated in one-dimensional and two-dimensional form in a MacCormack scheme against the Riemann and convergent duct problem. For this explicit scheme, the filter shows great improvements in accuracy and computational time as opposed to the nonfiltered solutions. However, for the implicit PARC3D code it is found that the best estimate of the Pegasus experimental heat fluxes and surface pressures is the simulation utilizing low artificial dissipation and no filter. The filter does improve accuracy over the artificially dissipative case but at a computational expense greater than that achieved by the low artificial dissipation case which has no computational time penalty and shows better results. For the shock-boundary layer simulation, the filter does well in terms of accuracy for a strong impingement shock but not as well for weaker shock strengths. Furthermore, for the latter problem the filter reduces the required computational time to convergence by 18.7 percent.
    Keywords: AERODYNAMICS
    Type: NASA-CR-186033 , H-2071 , NAS 1.26:186033
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  • 15
    Publication Date: 2019-06-28
    Description: An optimization procedure is developed for the simultaneous improvement of the aerodynamic and sonic boom characteristics of high speed aircraft. From a sonic boom perspective, it is desirable to minimize the first peak in the overpressure signal at a specified distance away from the aircraft. From aerodynamic point of view, the aerodynamic drag coefficient ratio must be minimized while maintaining the lift coefficient at desired level. The optimization procedure is applied to wing-body configurations related to high speed aircraft. The objectives of this current research are: (1) development of a multiobjective optimization procedure for aerospace vehicles with the integration of sonic boom and aerodynamic performance criteria; and (2) development of semi-analytical approach for calculating sonic boom design sensitivities.
    Keywords: AERODYNAMICS
    Type: NASA-CR-199083 , NAS 1.26:199083
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  • 16
    Publication Date: 2019-06-28
    Description: During the Higher Harmonic Control Aeroacoustic Rotor Test, extensive measurements of the rotor aerodynamics, the far-field acoustics, the wake geometry, and the blade motion for powered, descent, flight conditions were made. These measurements have been used to validate and improve the prediction of blade-vortex interaction (BVI) noise. The improvements made to the BVI modeling after the evaluation of the test data are discussed. The effects of these improvements on the acoustic-pressure predictions are shown. These improvements include restructuring the wake, modifying the core size, incorporating the measured blade motion into the calculations, and attempting to improve the dynamic blade response. A comparison of four different implementations of the Ffowcs Williams and Hawkings equation is presented. A common set of aerodynamic input has been used for this comparison.
    Keywords: AERODYNAMICS
    Type: NASA-TM-110825 , NAS 1.15:110825 , AD-A294477
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  • 17
    Publication Date: 2019-06-28
    Description: An extensive quantity of airload measurements was obtained for a pressure-instrumented model of the BO-105 main rotor for a large number of higher-harmonic control (HHC) settings at Duits-Nederlandse Wind Tunnel (DNW). The wake geometry, vortex strength, and vortex core size were also measured through a laser light sheet technique and LDV. These results are used to verify the BVI airload prediction methodologies developed by AFDD, DLR, NASA Langley, and ONERA. The comparisons show that an accurate prediction of the blade motion and the wake geometry is the most important aspect of the BVI airload predictions.
    Keywords: AERODYNAMICS
    Type: NASA-TM-110824 , NAS 1.15:110824 , AD-A294468
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  • 18
    Publication Date: 2019-06-28
    Description: This is a guide for the use of the pressure disk rotor model that has been placed in the incompressible Navier-Stokes code INS3D-UP. The pressure disk rotor model approximates a helicopter rotor or propeller in a time averaged manner and is intended to simulate the effect of a rotor in forward flight on the fuselage or the effect of a propeller on other aerodynamic components. The model uses a modified actuator disk that allows the pressure jump across the disk to vary with radius and azimuth. The cyclic and collective blade pitch angles needed to achieve a specified thrust coefficient and zero moment about the hub are predicted. The method has been validated with experimentally measured mean induced inflow velocities as well as surface pressures on a generic fuselage. Overset grids, sometimes referred to as Chimera grids, are used to simplify the grid generation process. The pressure disk model is applied to a cylindrical grid which is embedded in the grid or grids used for the rest of the configuration. This document will outline the development of the method, and present input and results for a sample case.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4692 , NAS 1.26:4692
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  • 19
    Publication Date: 2019-06-28
    Description: Future hypersonic vehicles are going to be designed largely with computational fluid dynamic methods based on appropriate physical models. The question on how much of this design process can be completed with the present state of computational aerothermodynamics is addressed. Some limitations of current models are discussed. It is shown that much more research is required before it will be possible to accurately design a hypersonic vehicle for all of its flight conditions. The quantities that must be computed accurately so that a minimum weight hypersonic vehicle can be designed are discussed. The use of computational fluid dynamics methods coupled with current thermochemical models in order to compute the quantities under specific flow conditions is considered.
    Keywords: AERODYNAMICS
    Type: ESA, Proceedings of the 2nd European Symposium on Aerothermodynamics for Space Vehicles; p 365-37
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  • 20
    Publication Date: 2019-06-28
    Description: An experimental investigation was conducted to determine the aerodynamic characteristics of a store as it was separated from the lee side of a flat plate inclined at 15 deg to the free-stream flow at Mach 6. Two store models were tested: a cone cylinder and a roof delta. Force and moment data were obtained for both stores as they were moved in 0.5-in. increments away from the flat plate lee-side separated flow region into the free-stream flow while the store angle of attack was held constant at either 0 deg or 15 deg. The results indicate that both stores had adverse separation characteristics (i.e., negative normal force and pitching moment) at an angle of attack of 0 deg, and the cone cylinder had favorable separation characteristics (i.e., positive normal force and pitching moment) at an angle of attack of 15 deg. At an angle of attack of 15 deg, the separation characteristics of the roof delta are indeterminate at small separation distances and favorable at greater separation distances. These characteristics are the result of the local flow inclination relative to the stores as they traversed through the flat plate lee-side flow field. In addition to plotted data, force and moment data are tabulated and schlieren photographs of the stores and flat plate are presented.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4652 , L-17384 , NAS 1.15:4652
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  • 21
    Publication Date: 2019-06-28
    Description: A high Reynolds number investigation of a commercial transport model was conducted in the National Transonic Facility (NTF) at Langley Research Center. This investigation was part of a cooperative effort to test a 0.03-scale model of a Boeing 767 airplane in the NTF over a Mach number range of 0.70 to 0.86 and a Reynolds number range of 2.38 to 40.0 x 10(exp 6) based on the mean aerodynamic chord. One of several specific objectives of the current investigation was to evaluate the level of data repeatability attainable in the NTF. Data repeatability studies were performed at a Mach number of 0.80 with Reynolds numbers of 2.38, 4.45, and 40.0 x 10(exp 6) and also at a Mach number of 0.70 with a Reynolds number of 40.0 x 10(exp 6). Many test procedures and data corrections are addressed in this report, but the data presented do not include corrections for wall interference, model support interference, or model aeroelastic effects. Application of corrections for these three effects would not affect the results of this study because the corrections are systematic in nature and are more appropriately classified as sources of bias error. The repeatability of the longitudinal stability-axis force and moment data has been accessed. Coefficients of lift, drag, and pitching moment are shown to repeat well within the pretest goals of plus or minus 0.005, plus or minus 0.0001, and plus or minus 0.001, respectively, at a 95-percent confidence level over both short- and near-term periods.
    Keywords: AERODYNAMICS
    Type: NASA-TP-3522 , L-17412 , NAS 1.60:3522
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  • 22
    Publication Date: 2019-06-28
    Description: An experimental investigation was conducted to determine the effect of diverter wedge half-angle and nacelle lip height on the drag characteristics of an assembly consisting of a nacelle fore cowl from a typical high-speed civil transport (HSCT) and a diverter mounted on a flat plate. Data were obtained for diverter wedge half-angles of 4.0 deg, 6.0 deg, and 8.0 deg and ratios of the nacelle lip height above a flat plate to the boundary-layer thickness (h(sub n)/delta) of approximately 0.87 to 2.45. Limited drag data were also obtained on a complete nacelle/diverter configuration that included fore and aft cowls. Although the nacelle/diverter drag data were not corrected for base pressures or internal flow drag, the data are useful for comparing the relative drag of the configuration tested. The tests were conducted in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.50, 1.80, 2.10, and 2.40 and Reynolds numbers ranging from 2.00 x 10(exp 6) to 5.00 x 10(exp 6) per foot. The results of this investigation showed that the nacelle/diverter drag essentially increased linearly with increasing h(sub n)/delta except near 1.0 where the data showed a nonlinear behavior. This nonlinear behavior was probably caused by the interaction of the shock waves from the nacelle/diverter configuration with the flat-plate boundary layer. At the lowest h(sub n)/delta tested, the diverter wedge half-angle had virtually no effect on the nacelle/diverter drag. However, as h(sub n)/delta increased, the nacelle/diverter drag increased as diverter wedge half-angle increased.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4660 , L-17416 , NAS 1.15:4660
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  • 23
    Publication Date: 2019-06-28
    Description: Water droplet trajectories within the NASA Lewis Research Center's Icing Research Tunnel (IRT) were studied through computer analysis. Of interest was the influence of the wind tunnel contraction and wind tunnel model blockage on the water droplet trajectories. The computer analysis was carried out with a program package consisting of a three-dimensional potential panel code and a three-dimensional droplet trajectory code. The wind tunnel contraction was found to influence the droplet size distribution and liquid water content distribution across the test section from that at the inlet. The wind tunnel walls were found to have negligible influence upon the impingement of water droplets upon a wing model.
    Keywords: AERODYNAMICS
    Type: NASA-TM-107023 , E-9828 , NAS 1.15:107023
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  • 24
    facet.materialart.
    Unknown
    In:  CASI
    Publication Date: 2019-06-28
    Description: This document outlines the tests performed to make aerodynamic force and torque measurements on the SOFIA wind tunnel model telescope. These tests were performed during the SOFIA 2 wind tunnel test in the 14 ft wind tunnel during the months of June through August 1994. The test was designed to measure the dynamic cross elevation moment acting on the SOFIA model telescope due to aerodynamic loading. The measurements were taken with the telescope mounted in an open cavity in the tail section of the SOFIA model 747. The purpose of the test was to obtain an estimate of the full scale aerodynamic disturbance spectrum, by scaling up the wind tunnel results (taking into account differences in sail area, air density, cavity dimension, etc.). An estimate of the full scale cross elevation moment spectrum was needed to help determine the impact this disturbance would have on the telescope positioning system requirements. A model of the telescope structure, made of a light weight composite material, was mounted in the open cavity of the SOFIA wind tunnel model. This model was mounted via a force balance to the cavity bulkhead. Despite efforts to use a 'stiff' balance, and a lightweight model, the balance/telescope system had a very low resonant frequency (37 Hz) compared to the desired measurement bandwidth (1000 Hz). Due to this mechanical resonance of the balance/telescope system, the balance alone could not provide an accurate measure of applied aerodynamic force at the high frequencies desired. A method of measurement was developed that incorporated accelerometers in addition to the balance signal, to calculate the aerodynamic force.
    Keywords: AERODYNAMICS
    Type: NASA-TM-110668 , SER-PK-001 , NAS 1.15:110668
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  • 25
    Publication Date: 2019-06-28
    Description: The work reported here pertains only to the first year of research for a three year proposal period. As a prelude to this two dimensional interface element, the one dimensional element was tested and errors were discovered in the code for built-up structures and curved interfaces. These errors were corrected and the benchmark Boeing composite crown panel was analyzed successfully. A study of various splines led to the conclusion that cubic B-splines best suit this interface element application. A least squares approach combined with cubic B-splines was constructed to make a smooth function from the noisy data obtained with random error in the coordinate data points of the Boeing crown panel analysis. Preliminary investigations for the formulation of discontinuous 2-D shell and 3-D solid elements were conducted.
    Keywords: AERODYNAMICS
    Type: NASA-CR-199951 , NAS 1.26:199951 , NIPS-96-07072
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  • 26
    Publication Date: 2019-06-28
    Description: Test flights were conducted to evaluate the capability of Differential Global Positioning System (DGPS) to provide the accuracy and integrity required for International Civil Aviation Organization (ICAO) Category (CAT) III precision approach and landings. These test flights were part of a Federal Aviation Administration (FAA) program to evaluate the technical feasibility of using DGPS based technology for CAT III precision approach and landing applications. An IAI Westwind 1124 aircraft (N24RH) was equipped with DGPS receiving equipment and additional computing capability provided by E-Systems. The test flights were conducted at NASA Ames Research Center's Crows Landing Flight Facility, Crows Landing, California. The flight test evaluation was based on completing 100 approaches and landings. The navigation sensor error accuracy requirements were based on ICAO requirements for the Microwave Landing System (MLS). All of the approaches and landings were evaluated against ground truth reference data provided by a laser tracker. Analysis of these approaches and landings shows that the E-Systems DGPS system met the navigation sensor error requirements for a successful approach and landing 98 out of 100 approaches and landings, based on the requirements specified in the FAA CAT III Level 2 Flight Test Plan. In addition, the E-Systems DGPS system met the integrity requirements for a successful approach and landing or stationary trial for all 100 approaches and landings and all ten stationary trials, based on the requirements specified in the FAA CAT III Level 2 Flight Test Plan.
    Keywords: AERODYNAMICS
    Type: NASA-TM-110368 , NAS 1.15:110368 , A-950096
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  • 27
    Publication Date: 2019-06-28
    Description: This monograph provides an extensive list of formulas for airfoil polynomials. These polynomials provide convenient expansion functions for the description of the downwash and pressure distributions of linear theory for airfoils in both steady and unsteady subsonic flow.
    Keywords: AERODYNAMICS
    Type: NASA-RP-1343 , L-17420 , NAS 1.61:1343
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  • 28
    Publication Date: 2019-06-28
    Description: Low-speed wind-tunnel tests were conducted in the Langley 12-Foot Low-Speed Tunnel on a model of the Boeing Multirole Fighter (BMRF) aircraft. This single-seat, single-engine configuration was intended to be an F-16 replacement that would incorporate many of the design goals and advanced technologies of the F-22. Its mission requirements included supersonic cruise without afterburner, reduced observability, and the ability to attack both air-to-air and air-to-ground targets. So that it would be effective in all phases of air combat, the ability to maneuver at angles of attack up to and beyond maximum lift was also desired. Traditional aerodynamic yaw controls, such as rudders, are typically ineffective at these higher angles of attack because they are usually located in the wake from the wings and fuselage. For this reason, this study focused on investigating forebody-mounted controls that produces yawing moments by modifying the strong vortex flowfield being shed from the forebody at high angles of attack. Two forebody strakes were tested that varied in planform and chordwise location. Various patterns of porosity in the forebody skin were also tested that differed in their radial coverage and chordwise location. The tests were performed at a dynamic pressure of 4 lb/ft(exp 2) over an angle-of-attack range of -4 deg to 72 deg and a sideslip range of -10 deg to 10 deg. Static force data, static pressures on the surface of the forebody, and videotapes of flow-visualization using laser-illuminated smoke were obtained.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4685 , NAS 1.26:4685
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  • 29
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    In:  CASI
    Publication Date: 2019-06-28
    Description: This paper presents an overview of the complex unsteady vortical flows that comprise the wakes of rotary-wing aircraft; of the effects these tangled vortical structures have on the performance, noise, and vibration; and of some of the recent attempts to measure, predict, and control the phenomena. The main points are illustrated with a number of examples from the recent literature and technical conferences.
    Keywords: AERODYNAMICS
    Type: NASA-TM-110822 , NAS 1.15:110822 , AD-A294465
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  • 30
    Publication Date: 2019-06-28
    Description: An unswept, semispan wing model incorporating a NACA 0012 airfoil section was tested in the Langley 14- by 22-Foot Subsonic Tunnel. This report contains pressure data which document effects of wing configuration and free-stream conditions on wing pressure distributions. The untwisted wing incorporated a full-span, leading-edge Krueger flap and a full-span, single-slotted trailing-edge flap. The trailing-edge flap was tested at a deflection angle of 40 degrees and the Krueger flap at a deflection of 55 degrees. Three wing configurations were tested: cruise, trailing-edge flap only, and Knueger flap and trailing-edge flap deployed. Tests were conducted at free-stream dynamic pressures of 15, 30 and 60 psf, with corresponding chord Reynolds numbers of 1.22 to 2.11 million and Mach numbers of 0.12 to 0.20. Angles of attack presented range from 0 to 20 degrees, depending on wing configuration. The data are presented without analysis.
    Keywords: AERODYNAMICS
    Type: NASA-TM-110148 , NAS 1.15:110148
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  • 31
    Publication Date: 2019-06-28
    Description: This report addresses the development of a multidisciplinary optimization procedure using an efficient semi-analytical sensitivity analysis technique and multilevel decomposition for the design of aerospace vehicles. A semi-analytical sensitivity analysis procedure is developed for calculating computational grid sensitivities and aerodynamic design sensitivities. Accuracy and efficiency of the sensitivity analysis procedure is established through comparison of the results with those obtained using a finite difference technique. The developed sensitivity analysis technique are then used within a multidisciplinary optimization procedure for designing aerospace vehicles. The optimization problem, with the integration of aerodynamics and structures, is decomposed into two levels. Optimization is performed for improved aerodynamic performance at the first level and improved structural performance at the second level. Aerodynamic analysis is performed by solving the three-dimensional parabolized Navier Stokes equations. A nonlinear programming technique and an approximate analysis procedure are used for optimization. The proceduredeveloped is applied to design the wing of a high speed aircraft. Results obtained show significant improvements in the aircraft aerodynamic and structural performance when compared to a reference or baseline configuration. The use of the semi-analytical sensitivity technique provides significant computational savings.
    Keywords: AERODYNAMICS
    Type: NASA-CR-199290 , NAS 1.26:199290
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  • 32
    Publication Date: 2019-06-28
    Description: The effectiveness of steady and pulsed blowing as a method of controlling delta wing vortices during ramp pitching has been investigated in flow visualization experiments conducted in a water tunnel. The recessed angled spanwise blowing technique was utilized for vortex manipulation. This technique was implemented on a beveled 60 delta wing using a pair of blowing ports located beneath the vortex core at 40% chord. The flow was injected primarily in the spanwise direction but was also composed of a component normal to the wing surface. The location of vortex burst was measured as a function of blowing intensity and pulsing frequency under static conditions, and the optimum blowing case was applied at three different wing pitching rates. Experimental results have shown that, when the burst location is upstream of the blowing port, pulsed blowing delays vortex breakdown in static and dynamic cases. Dynamic tests verified the existence of a hysteresis effect and demonstrated the improvements offered by pulsed blowing over both steady blowing and no-blowing scenarios. The application of blowing, at the optimum pulsing frequency, made the vortex breakdown location comparable in static and ramp pitch-up conditions.
    Keywords: AERODYNAMICS
    Type: NIPS-95-05494 , NASA-CR-199624 , NAS 1.26:199624 , AIAA PAPER 95-1817-CP , United States
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  • 33
    Publication Date: 2019-06-28
    Description: Primarily an experimental effort, this study focuses on the velocity and vorticity fields in the near wake of a hovering rotor. Drag terminology is reviewed, and the theory for separately determining the profile-and-induced-drag components from wake quantities is introduced. Instantaneous visualizations of the flow field are used to center the laser velocimeter (LV) measurements on the vortex core and to assess the extent of the positional mandering of the trailing vortex. Velocity profiles obtained at different rotor speeds and distances behind the rotor blade clearly indicate the position, size, and rate of movement of the wake sheet and the core of the trailing vortex. The results also show the distribution of vorticity along the wake sheet and within the trailing vortex.
    Keywords: AERODYNAMICS
    Type: NASA-TP-3577 , A-950078 , NAS 1.60:3577 , ATCOM-TR-95-A-006
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  • 34
    Publication Date: 2019-06-28
    Description: This report summarizes the research performed by North Carolina State University and NASA Ames Research Center under Cooperative Agreement NCA2-719, 'Numerical Simulation of Supersonic and Hypersonic Inlet Flow Fields". Four distinct rotated upwind schemes were developed and investigated to determine accuracy and practicality. The scheme found to have the best combination of attributes, including reduction to grid alignment with no rotation, was the cell centered non-orthogonal (CCNO) scheme. In 2D, the CCNO scheme improved rotation when flux interpolation was extended to second order. In 3D, improvements were less dramatic in all cases, with second order flux interpolation showing the least improvement over grid aligned upwinding. The reduction in improvement is attributed to uncertainty in determining optimum rotation angle and difficulty in performing accurate and efficient interpolation of the angle in 3D. The CCNO rotational technique will prove very useful for increasing accuracy when second order interpolation is not appropriate and will materially improve inlet flow solutions.
    Keywords: AERODYNAMICS
    Type: NASA-CR-199428 , NAS 1.26:199428
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  • 35
    Publication Date: 2019-06-28
    Description: Aeroelastic stability analyses have been performed for the MOD-5A blade/aileron system. Various configurations having different aileron torsional stiffness, mass unbalance, and control system damping have been investigated. The analysis was conducted using a code recently developed by the General Electric Company - AILSTAB. The code extracts eigenvalues for a three degree of freedom system, consisting of: (1) a blade flapwise mode; (2) a blade torsional mode; and (3) an aileron torsional mode. Mode shapes are supplied as input and the aileron can be specified over an arbitrary length of the blade span. Quasi-steady aerodynamic strip theory is used to compute aerodynamic derivatives of the wing-aileron combination as a function of spanwise position. Equations of motion are summarized herein. The program provides rotating blade stability boundaries for torsional divergence, classical flutter (bending/torsion) and wing/aileron flutter. It has been checked out against fixed-wing results published by Theodorsen and Garrick. The MOD-5A system is stable with respect to divergence and classical flutter for all practical rotor speeds. Aileron torsional stiffness must exceed a minimum critical value to prevent aileron flutter. The nominal control system stiffness greatly exceeds this minimum during normal operation. The basic system, however, is unstable for the case of a free (or floating) aileron. The instability can be removed either by the addition of torsional damping or mass-balancing the ailerons. The MOD-5A design was performed by the General Electric Company, Advanced Energy Program Department under Contract DEN3-153 with NASA Lewis Research Center and sponsored by the Department of Energy.
    Keywords: AERODYNAMICS
    Type: DASCON Engineering, Collected Papers on Wind Turbine Technology; p 99-114
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  • 36
    Publication Date: 2019-06-28
    Description: A method has been developed to accurately compute the viscous flow in three-dimensional (3-D) enclosures. This method is the 3-D extension of a two-dimensional (2-D) method developed for the calculation of flow over airfoils. The 2-D method has been tested extensively and has been shown to accurately reproduce experimental results. As in the 2-D method, the 3-D method provides for the non-iterative solution of the incompressible Navier-Stokes equations by means of a fully coupled implicit technique. The solution is calculated on a body fitted computational mesh incorporating a staggered grid methodology. In the staggered grid method, the three components of vorticity are defined at the centers of the computational cell sides, while the velocity components are defined as normal vectors at the centers of the computational cell faces. The staggered grid orientation provides for the accurate definition of the vorticity components at the vorticity locations, the divergence of vorticity at the mesh cell nodes and the conservation of mass at the mesh cell centers. The solution is obtained by utilizing a fractional step solution technique in the three coordinate directions. The boundary conditions for the vorticity and velocity are calculated implicitly as part of the solution. The method provides for the non-iterative solution of the flow field and satisfies the conservation of mass and divergence of vorticity to machine zero at each time step. To test the method, the calculation of simple driven cavity flows have been computed. The driven cavity flow is defined as the flow in an enclosure driven by a moving upper plate at the top of the enclosure. To demonstrate the ability of the method to predict the flow in arbitrary cavities, results will he shown for both cubic and curved cavities.
    Keywords: AERODYNAMICS
    Type: NASA-TM-110352 , A-950063 , NAS 1.15:110352
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  • 37
    Publication Date: 2019-06-28
    Description: An effort to understand and control the unsteady separated flow associated with the dynamic stall of airfoils was funded for three years through the NASA cooperative agreement program. As part of this effort a substantial data base was compiled detailing the effects various parameters have on the development of the dynamic stall flow field. Parameters studied include Mach number, pitch rate, and pitch history, as well as Reynolds number (through two different model chord lengths) and the condition of the boundary layer at the leading edge of the airfoil (through application of surface roughness). It was found for free stream Mach numbers as low as 0.4 that a region of supersonic flow forms on the leading edge of the suction surface of the airfoil at moderate angles of attack. The shocks which form in this supersonic region induce boundary-layer separation and advance the dynamic stall process. Under such conditions a supercritical airfoil profile is called for to produce a flow field having a weaker leading-edge pressure gradient and no leading-edge shocks. An airfoil having an adaptive-geometry, or dynamically deformable leading edge (DDLE), is under development as a unique active flow-control device. The DDLE, formed of carbon-fiber composite and fiberglass, can be flexed between a NACA 0012 profile and a supercritical profile in a controllable fashion while the airfoil is executing an angle-of-attack pitch-up maneuver. The dynamic stall data were recorded using point diffraction interferometry (PDI), a noninvasive measurement technique. A new high-speed cinematography system was developed for recording interferometric images. The system is capable of phase-locking with the pitching airfoil motion for real-time documentation of the development of the dynamic stall flow field. Computer-aided image analysis algorithms were developed for fast and accurate reduction of the images, improving interpretation of the results.
    Keywords: AERODYNAMICS
    Type: NASA-CR-198972 , NAS 1.26:198972 , MCAT-95-09
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  • 38
    Publication Date: 2019-06-28
    Description: A numerical scheme utilizing a chimera zonal grid approach for solving the full potential equation in two spatial dimensions is described. Within each grid zone a fully-implicit approximate factorization scheme is used to advance the solution one interaction. This is followed by the explicit advance of all common zonal grid boundaries using a bilinear interpolation of the velocity potential. The presentation is highlighted with numerical results simulating the flow about a two-dimensional, nonlifting, circular cylinder. For this problem, the flow domain is divided into two parts: an inner portion covered by a polar grid and an outer portion covered by a Cartesian grid. Both incompressible and compressible (transonic) flow solutions are included. Comparisons made with an analytic solution as well as single grid results indicate that the chimera zonal grid approach is a viable technique for solving the full potential equation.
    Keywords: AERODYNAMICS
    Type: NASA-TM-110360 , A-950082 , NAS 1.15:110360
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  • 39
    Publication Date: 2019-06-28
    Description: Distributions of static pressure coefficient over the afterbody and axisymmetric nozzles of a generic, twin-tail twin-engine fighter were obtained in the Langley 16-Foot Transonic Tunnel. The longitudinal positions of the vertical and horizontal tails were varied for a total of six aft-end configurations. Static pressure coefficients were obtained at Mach numbers between 0.6 and 1.2, angles of attack between 0 deg and 8 deg, and nozzle pressure ratios ranging from jet-off to 8. The results of this investigation indicate that the influence of the vertical and horizontal tails extends beyond the vicinity of the tail-afterbody juncture. The pressure distribution affecting the aft-end drag is influenced more by the position of the vertical tails than by the position of the horizontal tails. Transonic tail-interference effects are seen at lower free-stream Mach numbers at positive angles of attack than at an angle of attack of 0 deg.
    Keywords: AERODYNAMICS
    Type: NASA-TP-3509 , L-17438 , NAS 1.60:3509
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  • 40
    Publication Date: 2019-06-28
    Description: The National Aeronautics and Space Administration and the Defense Research Agency (United Kingdom) have ongoing experimental research programs in rotary-flow aerodynamics. A cooperative effort between the two agencies is currently underway to collect an extensive database for the development of high angle of attack computational methods to predict the effects of Reynolds number on the forebody flowfield at dynamic conditions, as well as to study the use of low Reynolds number data for the evaluation of high Reynolds number characteristics. Rotary balance experiments, including force and moment and surface pressure measurements, were conducted on circular and rectangular aftbodies with hemispherical and ogive noses at the Bedford and Farnborough wind tunnel facilities in the United Kingdom. The bodies were tested at 60 and 90 deg angle of attack for a wide range of Reynolds numbers in order to observe the effects of laminar, transitional, and turbulent flow separation on the forebody characteristics when rolling about the velocity vector.
    Keywords: AERODYNAMICS
    Type: NASA-CR-195033 , REPT-94-4 , NAS 1.26:195033
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  • 41
    Publication Date: 2019-06-28
    Description: A wind-tunnel investigation of the effectiveness of an aerodynamic yaw controller mounted on the lower surface of a shuttle orbiter model body flap was conducted in the Langley 31-Inch Mach 10 Tunnel. The controller consisted of a 60 deg delta fin mounted perpendicular to the body flap lower surface and yawed 30 deg to the free stream direction. The control was tested at angles of attack from 20 deg to 40 deg at zero sideslip for a Reynolds number based on wing mean aerodynamic chord of 0.66 x 10(exp 6). The aerodynamic and control effectiveness characteristics are presented along with an analysis of the effectiveness of the controller in making a bank maneuver for Mach 18 flight conditions. The controller was effective in yaw and produced a favorable rolling moment. The analysis showed that the controller was as effective as the reaction control system in making the bank maneuver. These results warrant further studies of the aerodynamic/aerothermodynamic characteristics of the control concept for application to future transportation vehicles.
    Keywords: AERODYNAMICS
    Type: NASA-TM-109179 , NAS 1.15:109179
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  • 42
    Publication Date: 2019-06-28
    Description: The approach to carrying out multi-discipline aerospace design studies in the future, especially in massively parallel computing environments, comprises of choosing (1) suitable solvers to compute solutions to equations characterizing a discipline, and (2) efficient optimization methods. In addition, for aerodynamic optimization problems, (3) smart methodologies must be selected to modify the surface shape. In this research effort, a 'direct' optimization method is implemented on the Cray C-90 to improve aerodynamic design. It is coupled with an existing implicit Navier-Stokes solver, OVERFLOW, to compute flow solutions. The optimization method is chosen such that it can accomodate multi-discipline optimization in future computations. In the work , however, only single discipline aerodynamic optimization will be included.
    Keywords: AERODYNAMICS
    Type: NASA-CR-198045 , NAS 1.26:198045 , OMI-02-93
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  • 43
    Publication Date: 2019-06-28
    Description: Modeling enhancements made to a radial-inflow turbine conceptual design code are documented in this report. A stator-endwall clearance-flow model was added for use with pivoting vanes. The rotor calculations were modified to account for swept blades and splitter blades. Stator and rotor trailing-edge losses and a vaneless-space loss were added to the loss model. Changes were made to the disk-friction and rotor-clearance loss calculations. The loss model was then calibrated based on experimental turbine performance. A complete description of code input and output along with sample cases are included in the report.
    Keywords: AERODYNAMICS
    Type: NASA-CR-195454 , E-9538 , NAS 1.26:195454
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  • 44
    Publication Date: 2019-06-28
    Description: Internal fluid flows are subject not only to self-sustained oscillations of the purely hydrodynamic type but also to the coupling of the instability with the acoustic mode of the surrounding cavity. This situation is common to turbomachinery, since flow instabilities are confined within a flow path where the acoustic wavelength is typically smaller than the dimensions of the cavity and flow speeds are low enough to allow resonances. When acoustic coupling occurs, the fluctuations can become so severe in amplitude that it may induce structural failure of engine components. The potential for catastrophic failure makes identifying flow-induced noise and vibration sources a priority. In view of the complexity of these types of flows, this report was written with the purpose of presenting many of the methods used to compute frequencies for self-sustained oscillations. The report also presents the engineering formulae needed to calculate the acoustic resonant modes for ducts and cavities. Although the report is not a replacement for more complex numerical or experimental modeling techniques, it is intended to be used on general types of flow configurations that are known to produce self-sustained oscillations. This report provides a complete collection of these models under one cover.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4671 , M-778 , NAS 1.26:4671
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  • 45
    Publication Date: 2019-06-28
    Description: This study was conducted to experimentally characterize the flow field created by the interaction of a single-expansion ramp-nozzle (SERN) flow with a hypersonic external stream. Data were obtained from a generic nozzle/afterbody model in the 3.5 Foot Hypersonic Wind Tunnel at the NASA Ames Research Center, in a cooperative experimental program involving Ames and McDonnell Douglas Aerospace. The model design and test planning were performed in close cooperation with members of the Ames computational fluid dynamics (CFD) team for the National Aerospace Plane (NASP) program. This paper presents experimental results consisting of oil-flow and shadow graph flow-visualization photographs, afterbody surface-pressure distributions, rake boundary-layer measurements, Preston-tube skin-friction measurements, and flow field surveys with five-hole and thermocouple probes. The probe data consist of impact pressure, flow direction, and total temperature profiles in the interaction flow field.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4638 , A-94119 , NAS 1.15:4638
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  • 46
    Publication Date: 2019-06-28
    Description: A methodology is described for computing viscous flows of air and sulfur hexafluoride (SF6). The basis is an existing flow solver that calculates turbulent flows in two dimensions on unstructured triangular meshes. The solver has been modified to incorporate the thermodynamic model for SF6 and used to calculate the viscous flow over two multielement airfoils that have been tested in a wind tunnel with air as the test medium. Flows of both air and SF6 at a free-stream Mach number of 0.2 and a Reynolds number of 9 x 10(exp 6) are computed for a range of angles of attack corresponding to the wind-tunnel test. The computations are used to investigate the suitability of SF6 as a test medium in wind tunnels and are a follow-on to previous computations for single-element airfoils. Surface-pressure, lift, and drag coefficients are compared with experimental data. The effects of heavy gas on the details of the flow are investigated based on computed boundary-layer and skin-friction data. In general, the predictions in SF6 vary little from those in air. Within the limitations of the computational method, the results presented are sufficiently encouraging to warrant further experiments.
    Keywords: AERODYNAMICS
    Type: NASA-TP-3496 , L-17401 , NAS 1.60:3496
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  • 47
    Publication Date: 2019-06-28
    Description: It is well known in the aerodynamic field that pressure distribution measurement over the surface of an aircraft model is a problem in experimental aerodynamics. For one thing, a continuous pressure map can not be obtained with the current experimental methods since they are discrete. Therefore, interpolation or CFD methods must be used for a more complete picture of the phenomenon under study. For this study, a new technique was investigated which would provide a continuous pressure distribution over the surface under consideration. The new method is pressure sensitive paint. When pressure sensitive paint is applied to an aerodynamic surface and placed in an operating wind-tunnel under appropriate lighting, the molecules luminesce as a function of the local pressure of oxygen over the surface of interest during aerodynamic flow. The resulting image will be brightest in the areas of low pressure (low oxygen concentration), and less intense in the areas of high pressure (where oxygen is most abundant on the surface). The objective of this investigation was to use pressure sensitive paint samples from McDonnell Douglas (MDD) for calibration purpose in order to assess the response of the paint under appropriate lighting and to use the samples over a flat plate/conical fin mounted at 75 degrees from the center of the plate in order to study the shock/boundary layer interaction at Mach 6 in the Von Karman wind-tunnel. From the result obtained it was concluded that temperature significantly affects the response of the paint and should be given the uppermost attention in the case of hypersonic flows. Also, it was found that past a certain temperature threshold, the paint intensity degradation became irreversible. The comparison between the pressure tap measurement and the pressure sensitive paint showed the right trend. However, there exists a shift when it comes to the actual value. Therefore, further investigation is under way to find the cause of the shift.
    Keywords: AERODYNAMICS
    Type: NASA-TM-106824 , E-9373 , NAS 1.15:106824
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  • 48
    Publication Date: 2019-06-28
    Description: Longitudinal characteristics and wing-section pressure distributions are compared for the EA-6B airplane with and without airfoil modifications. The airfoil modifications were designed to increase low-speed maximum lift for maneuvering, while having a minimal effect on transonic performance. Section contour changes were confined to the leading-edge slat and trailing-edge flap regions of the wing. Experimental data are analyzed from tests in the Langley 16-Foot Transonic Tunnel on the baseline and two modified wing-fuselage configurations with the slats and flaps in their retracted positions. Wing modification effects on subsonic and transonic performance are seen in wing-section pressure distributions of the various configurations at similar lift coefficients. The modified-wing configurations produced maximum lift coefficients which exceeded those of the baseline configuration at low-speed Mach numbers (0.300 and 0.400). This benefit was related to the behavior of the wing upper surface leading-edge suction peak and the behavior of the trailing-edge pressure. At transonic Mach numbers (0.725 to 0.900), the wing modifications produced a somewhat stronger nose-down pitching moment, a slightly higher drag at low-lift levels, and a lower drag at higher lift levels.
    Keywords: AERODYNAMICS
    Type: NASA-TP-3516 , L-17360 , NAS 1.60:3516
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  • 49
    Publication Date: 2019-06-28
    Description: This report describes the aerodynamics model used in a simulation model of an advanced short takeoff and vertical landing (ASTOVL) lift-fan fighter aircraft. The simulation model was developed for use in piloted evaluations of transition and hover flight regimes, so that only low speed (M approximately 0.2) aerodynamics are included in the mathematical model. The aerodynamic model includes the power-off aerodynamic forces and moments and the propulsion system induced aerodynamic effects, including ground effects. The power-off aerodynamics data were generated using the U.S. Air Force Stability and Control Digital DATCOM program and a NASA Ames in-house graphics program called VORVIEW which allows the user to easily analyze arbitrary conceptual aircraft configurations using the VORLAX program. The jet-induced data were generated using the prediction methods of R. E. Kuhn et al., as referenced in this report.
    Keywords: AERODYNAMICS
    Type: NASA-TM-110347 , A-950051 , NAS 1.15:110347
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  • 50
    Publication Date: 2019-06-28
    Description: A flow visualization study of several configurations of a jet-powered vertical takeoff and landing (VTOL) aircraft model during hover in ground effect was conducted. A surface oil flow technique was used to observe the flow patterns on the lower surfaces of the model. There were significant configuration effects. Wing height with respect to fuselage, the presence of an engine inlet duct beside the fuselage, and nozzle pressure ratio are seen to have strong effects on the surface flow angles on the lower surface of the wing. This test was part of a program to improve the methods for predicting the hot gas ingestion (HGI) for jet-powered vertical/short takeoff and landing (V/STOL) aircraft. The tests were performed at the Jet Calibration and Hover Test (JCAHT) Facility at Ames Research Center.
    Keywords: AERODYNAMICS
    Type: NASA-TM-108860 , A-95025 , NAS 1.15:108860
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  • 51
    Publication Date: 2019-06-28
    Description: Comprehensive experimental and analytical studies have been conducted to assess the potential aerodynamic benefits from spanwise blowing at the tip of a moderate-aspect-ratio swept wing. Previous studies on low-aspect-ratio wings indicated that blowing from the wingtip can diffuse the tip vortex and displace it outward. The diffused and displaced vortex will induce a smaller downwash at the wing, and consequently the wing will have increased lift and decreased induced drag at a given angle of attack. Results from the present investigation indicated that blowing from jets with a short chord had little effect on lift or drag, but blowing from jets with a longer chord increased lift near the tip and reduced drag at low Mach numbers. A Navier-Stokes solver with modified boundary conditions at the tip was used to extrapolate the results to a Mach number of 0.72. Calculations indicated that lift and drag increase with increasing jet momentum coefficient. Because the momentum of the jet is typically greater than the reduction in the wing drag and the increase in the wing lift due to spanwise blowing is small, spanwise blowing at the wingtip does not appear to be a practical means of improving the aerodynamic efficiency of moderate-aspectratio swept wings at high subsonic Mach numbers.
    Keywords: AERODYNAMICS
    Type: NASA-TP-3515 , L-17368 , NAS 1.60:3515
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  • 52
    Publication Date: 2019-06-28
    Description: This report summarizes accomplishments and progress for the period ending April 1995. Much of the work during this period has concentrated on preparation for an analysis of data produced by an extensive wind tunnel test. Time has also been spent further developing an empirical theory to account for the effects of blade-vortex interaction upon the circulation distribution of the vortex and on preliminary measurements aimed at controlling the vortex core size.
    Keywords: AERODYNAMICS
    Type: NASA-CR-198590 , NAS 1.26:198590
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  • 53
    Publication Date: 2019-06-28
    Description: Blade-element/momentum performance prediction codes are routinely used for wind turbine design and analysis. A weakness of these codes is their inability to consistently predict peak power upon which the machine structural design and cost are strongly dependent. The purpose of this study was to compare post-stall airfoil characteristics synthesization theory to a systematically acquired wind tunnel data set in which the effects of aspect ratio, airfoil thickness, and Reynolds number were investigated. The results of this comparison identified discrepancies between current theory and the wind tunnel data which could not be resolved. Other factors not previously investigated may account for these discrepancies and have a significant effect on peak power prediction.
    Keywords: AERODYNAMICS
    Type: DASCON Engineering, Collected Papers on Wind Turbine Technology; p 35-39
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  • 54
    Publication Date: 2019-06-28
    Description: Test flights were conducted to evaluate the capability of Differential Global Positioning System (DGPS) to provide the accuracy and integrity required for International Civil Aviation Organization (ICAO) Category (CAT) 3 precision approach and landings. These test flights were part of a Federal Aviation Administration (FAA) program to evaluate the technical feasibility of using DGPS based technology for CAT 3 precision approach and landing applications. A United Airlines Boeing 737-300 (N304UA) was equipped with DGPS receiving equipment and additional computing capability provided by Stanford University. The test flights were conducted at NASA Ames Research Center's Crows Landing Flight Facility, Crows Landing, California. The flight test evaluation was based on completing 100 approaches and autolandings; 90 touch and go, and 10 terminating with a full stop. Two types of accuracy requirements were evaluated: 1) Total system error, based on the Required Navigation Performance (RNP), and 2) Navigation sensor error, based on ICAO requirements for the Microwave Landing System (MLS). All of the approaches and autolandings were evaluated against ground truth reference data provided by a laser tracker. Analysis of these approaches and autolandings shows that the Stanford University/United Airlines system met the requirements for a successful approach and autolanding 98 out of 100 approaches and autolandings, based on the total system error requirements as specified in the FAA CAT 3 Level 2 Flight Test Plan.
    Keywords: AERODYNAMICS
    Type: NASA-TM-110354 , A-950066 , NAS 1.15:110354
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  • 55
    Publication Date: 2019-06-28
    Description: A proposed wing-body reusable launch vehicle was tested in the NASA Marshall Space Flight Center's 14 x 14-inch trisonic wind tunnel during the winter of 1994. This test resulted in the vehicle's subsonic and transonic, Mach 0.3 to 1.96, longitudinal and lateral aerodynamic characteristics. The effects of control surface deflections on the basic vehicle's aerodynamics, including a body flap, elevons, ailerons, and tip fins, are presented.
    Keywords: AERODYNAMICS
    Type: NASA-TM-108489 , NAS 1.15:108489
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  • 56
    Publication Date: 2019-06-28
    Description: A linearized unsteady aerodynamic analysis for axial-flow turbomachinery blading is described in this report. The linearization is based on the Euler equations of fluid motion and is motivated by the need for an efficient aerodynamic analysis that can be used in predicting the aeroelastic and aeroacoustic responses of blade rows. The field equations and surface conditions required for inviscid, nonlinear and linearized, unsteady aerodynamic analyses of three-dimensional flow through a single, blade row operating within a cylindrical duct, are derived. An existing numerical algorithm for determining time-accurate solutions of the nonlinear unsteady flow problem is described, and a numerical model, based upon this nonlinear flow solver, is formulated for the first-harmonic linear unsteady problem. The linearized aerodynamic and numerical models have been implemented into a first-harmonic unsteady flow code, called LINFLUX. At present this code applies only to two-dimensional flows, but an extension to three-dimensions is planned as future work. The three-dimensional aerodynamic and numerical formulations are described in this report. Numerical results for two-dimensional unsteady cascade flows, excited by prescribed blade motions and prescribed aerodynamic disturbances at inlet and exit, are also provided to illustrate the present capabilities of the LINFLUX analysis.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4677 , E-9575 , NAS 1.26:4677 , R95-970293
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  • 57
    Publication Date: 2019-06-28
    Description: Prediction of the detailed dynamic behavior in rocket propellant feed systems and engines and other such high-energy fluid systems requires precise analysis to assure structural performance. Designs sometimes require placement of bluff bodies in a flow passage. Additionally, there are flexibilities in ducts, liners, and piping systems. A design handbook and interactive data base have been developed for assessing flow/structural interactions to be used as a tool in design and development, to evaluate applicable geometries before problems develop, or to eliminate or minimize problems with existing hardware. This is a compilation of analytical/empirical data and techniques to evaluate detailed dynamic characteristics of both the fluid and structures. These techniques have direct applicability to rocket engine internal flow passages, hot gas drive systems, and vehicle propellant feed systems. Organization of the handbook is by basic geometries for estimating Strouhal numbers, added mass effects, mode shapes for various end constraints, critical onset flow conditions, and possible structural response amplitudes. Emphasis is on dense fluids and high structural loading potential for fatigue at low subsonic flow speeds where high-frequency excitations are possible. Avoidance and corrective measure illustrations are presented together with analytical curve fits for predictions compiled from a comprehensive data base.
    Keywords: AERODYNAMICS
    Type: NASA-CR-4652 , M-773 , NAS 1.26:4652
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  • 58
    Publication Date: 2019-06-28
    Description: An experimental force and moment study was conducted in the Langley 8-Foot Transonic Pressure Tunnel for a generic store in and near rectangular box cavities contained in a flat-plate configuration at subsonic and transonic speeds. Surface pressures were measured inside the cavities and on the flat plate. The length-to-height ratios were 5.42, 6.25, 10.83, and 12.50. The corresponding width-to-height ratios were 2.00, 2.00, 4.00, and 4.00. The free-stream Mach number range was from 0.20 to 0.95. Surface pressure measurements inside the cavities indicated that the flow fields for the shallow cavities were either closed or transitional near the transitional/closed boundary. For the deep cavities, the flow fields were either open or near the open/transitional boundary. The presence of the store did not change the type of flow field and had only small effects on the pressure distributions. For transitional or open transitional flow fields, increasing the free-stream Mach number resulted in large reductions in pitching-moment coefficient. Values of pitching-moment coefficient were always much greater for closed flow fields than for open flow fields.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4611 , L-17388 , NAS 1.15:4611
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  • 59
    Publication Date: 2019-06-28
    Description: This research has been aimed at validating numerical methods for computing the flow about the complete F-18 HARV at alpha = 30 deg and alpha = 45 deg. At 30 deg angle of attack, the flow about the F-18 is dominated by the formation, and subsequent breakdown, of strong vortices over the wing leading-edge extensions (LEX). As the angle of attack is increased to alpha = 45 deg, the fuselage forebody of the F-18 contains significant laminar and transitional regions which are not present at alpha = 30 deg. Further, the flow over the LEX at alpha = 45 deg is dominated by an unsteady shedding in time, rather than strong coherent vortices. This complex physics, combined with the complex geometry of a full-aircraft configuration, provides a challenge for current computational fluid dynamics (CFD) techniques. The following sections present the numerical method and grid generation scheme that was used, a review of prior research done to numerically model the F-18 HARV, and a discussion of the current research. The current research is broken into three main topics; the effect of engine-inlet mass-flow rate on the F-18 vortex breakdown position, the results using a refined F-18 computational model to compute the flow at alpha = 30 deg and alpha = 45 deg, and research done using the simplified geometry of an ogive-cylinder configuration to investigate the physics of unsteady shear-layer shedding. The last section briefly summarizes the discussion.
    Keywords: AERODYNAMICS
    Type: NASA-CR-197755 , NAS 1.26:197755 , MCAT-95-13
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  • 60
    Publication Date: 2019-06-28
    Description: A wind tunnel investigation was conducted on a generic, high-wing transport model in the Langley 14- by 22-Foot Subsonic Tunnel. This report contains pressure data that document effects of various model configurations and free-stream conditions on wing pressure distributions. The untwisted wing incorporated a full-span, leading-edge Krueger flap and a part-span, double-slotted trailing-edge flap system. The trailing-edge flap was tested at four different deflection angles (20 deg, 30 deg, 40 deg, and 60 deg). Four wing configurations were tested: cruise, flaps only, Krueger flap only, and high lift (Krueger flap and flaps deployed). Tests were conducted at free-stream dynamic pressures of 20 psf to 60 psf with corresponding chord Reynolds numbers of 1.22 x 10(exp 6) to 2.11 x 10(exp 6) and Mach numbers of 0.12 to 0.20. The angles of attack presented range from 0 deg to 20 deg and were determined by wing configuration. The angle of sideslip ranged from minus 20 deg to 20 deg. In general, pressure distributions were relatively insensitive to free-stream speed with exceptions primarily at high angles of attack or high flap deflections. Increasing trailing-edge Krueger flap significantly reduced peak suction pressures and steep gradients on the wing at high angles of attack. Installation of the empennage had no effect on wing pressure distributions. Unpowered engine nacelles reduced suction pressures on the wing and the flaps.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4583 , L-17380 , NAS 1.15:4583
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  • 61
    Publication Date: 2019-06-28
    Description: A series of thrust-vectored axisymmetric ejector nozzles were designed and experimentally tested for internal performance and pumping characteristics at the Langley research center. This study indicated that discontinuities in the performance occurred at low primary nozzle pressure ratios and that these discontinuities were mitigated by decreasing expansion area ratio. The addition of secondary flow increased the performance of the nozzles. The mid-to-high range of secondary flow provided the most overall improvements, and the greatest improvements were seen for the largest ejector area ratio. Thrust vectoring the ejector nozzles caused a reduction in performance and discharge coefficient. With or without secondary flow, the vectored ejector nozzles produced thrust vector angles that were equivalent to or greater than the geometric turning angle. With or without secondary flow, spacing ratio (ejector passage symmetry) had little effect on performance (gross thrust ratio), discharge coefficient, or thrust vector angle. For the unvectored ejectors, a small amount of secondary flow was sufficient to reduce the pressure levels on the shroud to provide cooling, but for the vectored ejector nozzles, a larger amount of secondary air was required to reduce the pressure levels to provide cooling.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4610 , L-17386 , NAS 1.15:4610
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  • 62
    Publication Date: 2019-06-28
    Description: A three-dimensional Navier-Stokes solver was used to determine how accurately computations can predict local and average skin friction coefficients for attached and separated flows for simple experimental geometries. Algebraic and transport equation closures were used to model turbulence. To simulate anisotropic turbulence, the standard two-equation turbulence model was modified by adding nonlinear terms. The effects of both grid density and the turbulence model on the computed flow fields were also investigated and compared with available experimental data for subsonic and supersonic free-stream conditions.
    Keywords: AERODYNAMICS
    Type: NASA-TP-3480 , L-17354 , NAS 1.60:3480
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  • 63
    Publication Date: 2019-06-28
    Description: Modifications made to the axial-flow compressor conceptual design code CSPAN are documented in this report. Endwall blockage and stall margin predictions were added. The loss-coefficient model was upgraded. Default correlations for rotor and stator solidity and aspect-ratio inputs and for stator-exit tangential velocity inputs were included in the code along with defaults for aerodynamic design limits. A complete description of input and output along with sample cases are included.
    Keywords: AERODYNAMICS
    Type: NASA-TM-106833 , E-9394 , NAS 1.15:106833
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  • 64
    Publication Date: 2019-06-28
    Description: A numerical model has been developed at the NASA Lewis Research Center which can predict both the unsteady flow quantities within a wave rotor passage and the steady averaged flows in the ports. The model is based on the assumptions of one-dimensional, unsteady, perfect gas flow. The model assesses not only the dominant wave behavior, but the loss effects of finite passage opening time, leakage from the passage ends, viscosity, and heat transfer to and from the passages. The model operates in the rotor reference frame; however, until recently no account was made for the often significant effect of the rotor circumferential velocity component. The present model accounts for this by modifying the passage boundary conditions, allowing the internal computational scheme to remain the rotor reference frame, while quantities such as inlet duct stagnation properties may be specified in the fixed or absolute reference frame. Accurate modeling of this effect is critical to successful wave rotor analysis and design, particularly in off-design predictions where the flows in the inlet ducts are mismatched with the rotor passages and significant turning may take place (i.e., work is done on the gas). The relative simplicity of the model makes it useful for design and optimization, as well as analysis, of wave rotor cycles for many applications. This report, building on several earlier papers, describes the most recent modifications to the model. These include accounting for the relative/absolute transition at the passage boundaries and refinements to the viscous source term correlation which resulted from this accounting. Comparison of model predictions with measured data is then presented and discussed.
    Keywords: AERODYNAMICS
    Type: NASA-TM-106913 , E-9621 , NAS 1.15:106913
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  • 65
    Publication Date: 2019-06-28
    Description: An analytical/numerical method has been developed to predict the static thrust performance of non-axisymmetric, two-dimensional convergent-divergent exhaust nozzles. Thermodynamic nozzle performance effects due to over- and underexpansion are modeled using one-dimensional compressible flow theory. Boundary layer development and skin friction losses are calculated using an approximate integral momentum method based on the classic karman-Polhausen solution. Angularity effects are included with these two models in a computational Nozzle Performance Analysis Code, NPAC. In four different case studies, results from NPAC are compared to experimental data obtained from subscale nozzle testing to demonstrate the capabilities and limitations of the NPAC method. In several cases, the NPAC prediction matched experimental gross thrust efficiency data to within 0.1 percent at a design NPR, and to within 0.5 percent at off-design conditions.
    Keywords: AERODYNAMICS
    Type: NASA-CR-195050 , NAS 1.26:195050
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  • 66
    Publication Date: 2019-06-28
    Description: The Chimera overset grid method is reviewed and discussed in the context of a method of solution and analysis of unsteady three-dimensional viscous flows. The state of maturity of the various pieces of support software required to use the approach is discussed. A variety of recent applications of the method is presented. Current limitations of the approach are identified.
    Keywords: AERODYNAMICS
    Type: NASA-CR-198603 , NAS 1.26:198603 , OMI-01-92
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  • 67
    Publication Date: 2019-06-28
    Description: A flow visualization investigation with a laser light sheet system was conducted on a 27-percent-scale AH-64 attack helicopter model fitted with two mast-mounted sights in the langley 14- by 22-foot subsonic tunnel. The investigation was conducted to identify aerodynamic phenomena that may have contributed to adverse vibration encountered during full-scale flight of the AH-64D apache/longbow helicopter with an asymmetric mast-mounted sight. Symmetric and asymmetric mast-mounted sights oriented at several skew angles were tested at simulated forward and rearward flight speeds of 30 and 45 knots. A laser light sheet system was used to visualize the flow in planes parallel to and perpendicular to the free-stream flow. Analysis of these flow visualization data identified frequencies of flow patterns in the wake shed from the sight, the streamline angle at the sight, and the location where the shed wake crossed the rotor plane. Differences in wake structure were observed between the sight configurations and various skew angles. Analysis of lateral light sheet plane data implied significant vortex structure in the wake of the asymmetric mast-mounted sight in the configuration that produced maximum in-flight vibration. The data showed no significant vortex structure in the wake of the asymmetric and symmetric configurations that produced no increase in in-flight adverse vibration.
    Keywords: AERODYNAMICS
    Type: NASA-TM-4634 , L-17409 , NAS 1.15:4634 , ATCOM-TR-95-A-001
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  • 68
    Publication Date: 2019-07-20
    Description: In recent years, active research has been conducted to study the technological feasibility of supersonic laminar flow control on the wing of the High Speed Civil Transport (HSCT). For this study, the F-16XL has been chosen due to its highly swept crank wing planform that closely resembles the HSCT configurations. During flights, it is discovered that the shock wave generated from the aircraft inlet introduces disturbances on the wing where the data acquisition is conducted. The flow field about a supersonic inlet is characterized by a complex three dimensional pattern of shock waves generated by the geometrical configuration of a deflector and a cowl lip. Hence, in this study, experimental method is employed to investigate the effects of the variation of deflector configuration on the flow field, and consequently, the possibility of diverting the incoming shock-disturbances away from the test section. In the present experiments, a model composed of a simple circular tube with a triangular deflector is designed to study the deflector length and the deflector base width variation in the flow field. Experimental results indicate that the lowest external pressure ratio is observed at the junction where the deflector lip and the inlet cowl lip merge. Also, it is noted that the external pressure ratio, the internal pressure ratio, the coefficient of spillage drag, and the shock standoff distance decrease as the deflector length increases. In addition, the Redefined Total Pressure Recovery Ratio (RTPRR) increases with an increase in the deflector length. Results from the study of the effect of the deflector's base width variation on the flow field indicate that the lowest external pressure ratio is observed at the junction between the inlet cowl lip and the deflector lip. As the base width of the deflector increases, the external pressure ratio at 0 rotation increases, whereas the external pressure ratio at 180 rotation decreases. In addition, the internal pressure ratio and the coefficient of spillage drag decrease as the base width of the deflector increases. However, RTPRR and shock standoff distance increase as the base width increases. In conclusion, as deflector dimensions vary, distinctive patterns in the pressure variation around the inlet deflector are observed. With an increase in the deflector length and base width, the magnitude of shock-disturbances are weakened due to a decrease in the external pressure ratio. Also, as the deflector length and base width increase, a smaller bow shock angle is formed. Therefore, the inlet shock wave formation would be significantly altered, and consequently, shock disturbances on the wing test section can be avoided through appropriately designing the deflector.
    Keywords: AERODYNAMICS
    Type: NASA-CR-199617 , NAS 1.26:199617
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  • 69
    Publication Date: 2019-07-20
    Description: An investigation was conducted from October 1, 1990 to May 31, 1994 on the development of methodologies to improve the designs (more specifically, the shape) of aerodynamic surfaces of coupling optimization algorithms (OA) with Computational Fluid Dynamics (CFD) algorithms via sensitivity analyses (SA). The study produced several promising methodologies and their proof-of-concept cases, which have been reported in the open literature.
    Keywords: AERODYNAMICS
    Type: NASA-CR-197419 , NAS 1.26:197419
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  • 70
    Publication Date: 2019-07-13
    Description: The NPARC code, a Reynolds-averaged full Navier-Stokes code, was validated for nozzle afterbody (boatail) flow fields at transonic speeds. The flow fields about three geometries were studied: an axisymmetric nozzle with attached flow; an axisymmetric nozzle with separated flow: and a two-dimensional (rectangular) nozzle with separated flow. Three turbulence models, Baldwin-Lomax, Baldwin-Barth, and Chien k-epsilon, were used to determine the effect of turbulence model selection on the flow field solution. Static pressure distributions on the nozzle surfaces and pitot pressure measurements in the exhaust plume were examined. Results from the NPARC code compared very well with experimental data for all cases. For attached flow fields, the effect of the turbulence models showed no discernable differences. The Baldwin-Barth model yielded better results than either the Chien k-epsilon or the Baldwin-Lomax model for separated flow fields.
    Keywords: AERODYNAMICS
    Type: NASA-TM-106971 , E-9732 , NAS 1.15:106971 , AIAA PAPER 95-2614 , AIAA, ASME, SAE, and ASEE; Jul 10, 1995 - Jul 12, 1995; US
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  • 71
    Publication Date: 2019-07-13
    Description: The effects of grid resolution and specification of turbulent inflow boundary conditions were examined using the NPARC code with the Baldwin-Lomax and Chien k-e turbulence models. Three benchmark turbulent test cases were calculated: two were wall bounded flows and the third was a compressible mixing layer. The wall bounded flows were essentially insensitive to axial grid density; however, the location of the first point off the wall had a substantial effect on flow solutions. It was determined that the first point off the wall must be in the laminar sublayer (y+ less than or equal to 5) for the entire boundary layer. For the compressible mixing layer cases, the axial grid density affected the capturing of oblique shock waves in the mixing region, but the overall mixing rate was not strongly dependent on grid resolution. In specifying the inflow turbulent boundary conditions, it was very important to match the boundary layer and momentum thicknesses of the two flows entering the mixing region; calculations obtained with smaller or no boundary layers resulted in substantially reduced mixing. The solutions were relatively insensitive to freestream turbulence level.
    Keywords: AERODYNAMICS
    Type: NASA-TM-106959 , E-9711 , NAS 1.15:106959 , AIAA PAPER 95-2613 , Joint Propulsion Conference and Exhibit; Jul 10, 1995 - Jul 12, 1995; San Diego, CA; United States
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  • 72
    Publication Date: 2019-07-13
    Description: The transonic performance of a dual-throat, single-expansion-ramp nozzle (SERN) was investigated with a PARC computational fluid dynamics (CFD) code, an external flow Navier-Stokes solver. The nozzle configuration was from a conceptual Mach 5 cruise aircraft powered by four air-breathing turboramjets. Initial test cases used the two-dimensional version of PARC in Euler mode to investigate the effect of geometric variation on transonic performance. Additional cases used the two-dimensional version in viscous mode and the three-dimensional version in both Euler and viscous modes. Results of the analysis indicate low nozzle performance and a highly three-dimensional nozzle flow at transonic conditions. In another comparative study using the PARC code, a single-throat SERN configuration for which experimental data were available at transonic conditions was used to validate the results of the over/under turboramjet nozzle.
    Keywords: AERODYNAMICS
    Type: NASA-TM-106924 , E-9638 , NAS 1.15:106924 , AIAA PAPER 95-2616 , Joint Propulsion Conference and Exhibit; Jul 10, 1995 - Jul 12, 1995; San Diego, CA; United States
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  • 73
    Publication Date: 2019-07-13
    Description: An experimental study is conducted to examine the crossplane structure and streamwise decay of vortices shed from airfoil-type vortex generators. The vortex generators are set in a counter-rotating array spanning the full circumference of a straight pipe. The span of the vortex generators above the duct surface, h, is approximately equal to the local turbulent boundary layer thickness, delta. Measurement of three-component mean flow velocity in downstream crossplanes are used to characterize the structure of the shed vortices. Measurements in adjacent crossplanes (closely spaced along the streamwise coordinate) characterize the interaction and decay of the embedded vortices. A model constructed by the superposition of Oseen vortices is compared to the data for one test case.
    Keywords: AERODYNAMICS
    Type: NASA-CR-198356 , E-9730 , NAS 1.26:198356 , AIAA PAPER 95-1797 , Applied Aerodynamics Conference; Jun 19, 1995 - Jun 22, 1995; San Diego, CA; United States
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  • 74
    Publication Date: 2019-07-13
    Description: The NASA Steering Committee for Surface Modeling and Grid Generation (SMAGG) sponsored a workshop on surface modeling, grid generation, and related issues in Computational Fluid Dynamics (CFD) solutions at Lewis Research Center, Cleveland, Ohio, May 9-11, 1995. The workshop provided a forum to identify industry needs, strengths, and weaknesses of the five grid technologies (patched structured, overset structured, Cartesian, unstructured, and hybrid), and to exchange thoughts about where each technology will be in 2 to 5 years. The workshop also provided opportunities for engineers and scientists to present new methods, approaches, and applications in SMAGG for CFD. This Conference Publication (CP) consists of papers on industry overview, NASA overview, five grid technologies, new methods/ approaches/applications, and software systems.
    Keywords: AERODYNAMICS
    Type: NASA-CP-3291 , E-9458 , NAS 1.55:3291 , May 09, 1995 - May 11, 1995; Cleveland, OH; United States
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  • 75
    Publication Date: 2019-07-13
    Description: A test program was conducted on the third Mod-2 unit at Goldendale, Washington, to systematically study the effect of vortex generators (VG's) on power performance. The subject unit was first tested without VG's to obtain baseline data. Vortex generators were then installed on the mid-blade assemblies, and the resulting 70% VG configuration was tested. Finally, vortex generators were mounted on the tip assemblies, and data was recorded for the 100% VG configuration. This test program and its results are discussed in this paper. The development of vortex generators is also presented.