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  • AERODYNAMICS
  • Aerodynamics
  • Chemical Engineering
  • GENERAL
  • PROPELLANTS
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  • 2005-2009  (11)
  • 1955-1959
  • 2006  (11)
  • 1
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    Reynolds-Averaged Navier-Stokes Analysis of Zero Efflux Flow Control over a Hump Model (2006)
    In:  CASI
    Details
    Publication Date: 2018-06-11
    Description: The unsteady flow over a hump model with zero efflux oscillatory flow control is modeled computationally using the unsteady Reynolds-averaged Navier-Stokes equations. Three different turbulence models produce similar results, and do a reasonably good job predicting the general character of the unsteady surface pressure coefficients during the forced cycle. However, the turbulent shear stresses are underpredicted in magnitude inside the separation bubble, and the computed results predict too large a (mean) separation bubble compared with experiment. These missed predictions are consistent with earlier steady-state results using no-flow-control and steady suction, from a 2004 CFD validation workshop for synthetic jets.
    Keywords: Aerodynamics
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  • 2
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    Characteristic Lifelength of Coherent Structure in the Turbulent Boundary Layer (2006)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A characteristic lifelength is defined by which a Gaussian distribution is fit to data correlated over a 3 sensor array sampling streamwise sidewall pressure. The data were acquired at subsonic, transonic and supersonic speeds aboard a Tu-144. Lifelengths are estimated using the cross spectrum and are shown to compare favorably with Efimtsov's prediction of correlation space scales. Lifelength distributions are computed in the time/frequency domain using an interval correlation technique on the continuous wavelet transform of the original time data. The median values of the lifelength distributions are found to be very close to the frequency averaged result. The interval correlation technique is shown to allow the retrieval and inspection of the original time data of each event in the lifelength distribution, thus providing a means to locate and study the nature of the coherent structure in the turbulent boundary layer. The lifelength data can be converted to lifetimes using the convection velocity. The lifetime of events in the time/frequency domain are displayed in Lifetime Maps. The primary purpose of the paper is to validate these new analysis techniques so that they can be used with confidence to further characterize coherent structure in the turbulent boundary layer.
    Keywords: Aerodynamics
    Type: AIAA Paper 2006-2410 , 12th AIAA/CEAS Aeroacoustics Conference; May 08, 2006 - May 10, 2006; Cambridge, MA; United States
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  • 3
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    An Assessment of NASA Glenn's Aeroacoustic Experimental and Predictive Capabilities for Installed Cooling Fans (2006)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Quiet, high performance electronics cooling fans are needed for both commercial applications and future manned space exploration missions. Researchers at NASA Glenn focusing on aircraft engine noise, have long been familiar with the challenge of reducing fan noise without sacrificing aerodynamic performance. Is it possible to capitalize on the lessons-learned in aircraft engine noise reduction to identify inexpensive ways to improve the aerodynamic and acoustic performance of electronics cooling fans? Recent tests at NASA Glenn have begun to look for answers to this question. The overall aerodynamic and acoustic performance of a commercially available, spaceflight qualified 80 mm diameter axial flow fan has been measured using an automated plenum in accordance with ISO 10302 in the hemi-anechoic chamber of NASA Glenn s Acoustical Testing Laboratory. These measurements are complemented by detailed aerodynamic measurements of the inlet, exhaust, and rotor wake regions of the fan using Particle Image Velocimetry and hot-wire probes. A study of preliminary results yielded recommendations for system designers, fan manufacturers, and researchers.
    Keywords: Aerodynamics
    Type: NASA/TM-2006-214450 , E-15713 , 35th International Congress and Exposition on Noise Control Engineering (INTER-NOISE 2006); Dec 03, 2006 - Dec 06, 2006; Honolulu, HI; United States
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  • 4
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    Designing and Testing a Blended Wing Body with Boundary Layer Ingestion Nacelles (2006)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A knowledge-based aerodynamic design method coupled with an unstructured grid Navier-Stokes flow solver was used to improve the propulsion/airframe integration for a Blended Wing Body with boundary-layer ingestion nacelles. A new zonal design capability was used that significantly reduced the time required to achieve a successful design for each nacelle and the elevon between them. A wind tunnel model was built with interchangeable parts reflecting the baseline and redesigned configurations and tested in the National Transonic Facility (NTF). Most of the testing was done at the cruise design conditions (Mach number = 0.85, Reynolds number = 75 million). In general, the predicted improvements in forces and moments as well as the changes in wing pressures between the baseline and redesign were confirmed by the wind tunnel results. The effectiveness of elevons between the nacelles was also predicted surprisingly well considering the crudeness in the modeling of the control surfaces in the flow code.
    Keywords: Aerodynamics
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  • 5
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    A Summary of the Experimental Results for a Generic Tractor-Trailer in the Ames Research Center 7- by 10-Foot and 12-Foot Wind Tunnels (2006)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: Experimental measurements of a generic tractor-trailer were obtained in two wind tunnels at Ames Research Center. After a preliminary study at atmospheric conditions in the 7- by 10-Foot Wind Tunnel, additional testing was conducted at Reynolds numbers corresponding to full-scale highway speeds in the 12-Foot Pressure Wind Tunnel. To facilitate computational modeling, the 1:8-scale geometry, designated the Generic Conventional Model, included a simplified underbody and omitted many small-scale details. The measurements included overall and component forces and moments, static and dynamic surface pressures, and three-component particle image velocimetry. This summary report highlights the effects of numerous drag reduction concepts and provides details of the model installation in both wind tunnels. To provide a basis for comparison, the wind-averaged drag coefficient was tabulated for all configurations tested. Relative to the baseline configuration representative of a modern class-8 tractor-trailer, the most effective concepts were the trailer base flaps and trailer belly box providing a drag-coefficient reduction of 0.0855 and 0.0494, respectively. Trailer side skirts were less effective yielding a drag reduction of 0.0260. The database of this experimental effort is publicly available for further analysis.
    Keywords: Aerodynamics
    Type: NASA/TM-2006-213489 , A-060009
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  • 6
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    Low Reynolds Number Aerodynamic Characteristics of Several Airplane Configurations Designed to Fly in the Mars Atmosphere at Subsonic Speeds (2006)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A 1/4-scale wind tunnel model of an airplane configuration developed for short duration flight at subsonic speeds in the Martian atmosphere has been tested in the Langley Research Center Transonic Dynamics Tunnel. The tunnel was pumped down to extremely low pressures to represent Martian Mach/Reynolds number conditions. Aerodynamic data were obtained and upper and lower surface wind pressures were measured at one spanwise station on some configurations. Three unswept wings of the same planform but different airfoil sections were tested. Horizontal tail incidence was varied as was the deflection of plain and split trailing-edge flaps. One unswept wing configuration was tested with the lower part of the fuselage removed and the vertical/horizontal tail assembly inverted and mounted from beneath the fuselage. A sweptback wing was also tested. Tests were conducted at Mach numbers from 0.50 to 0.90. Wing chord Reynolds number was varied from 40,000 to 100,000 and angles of attack and sideslip were varied from -10deg to 20deg and -10deg to 10deg, respectively.
    Keywords: Aerodynamics
    Type: NASA/TM-2006-214312 , L-19252
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  • 7
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    Convergence Acceleration for Multistage Time-Stepping Schemes (2006)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The convergence of a Runge-Kutta (RK) scheme with multigrid is accelerated by preconditioning with a fully implicit operator. With the extended stability of the Runge-Kutta scheme, CFL numbers as high as 1000 could be used. The implicit preconditioner addresses the stiffness in the discrete equations associated with stretched meshes. Numerical dissipation operators (based on the Roe scheme, a matrix formulation, and the CUSP scheme) as well as the number of RK stages are considered in evaluating the RK/implicit scheme. Both the numerical and computational efficiency of the scheme with the different dissipation operators are discussed. The RK/implicit scheme is used to solve the two-dimensional (2-D) and three-dimensional (3-D) compressible, Reynolds-averaged Navier-Stokes equations. In two dimensions, turbulent flows over an airfoil at subsonic and transonic conditions are computed. The effects of mesh cell aspect ratio on convergence are investigated for Reynolds numbers between 5.7 x 10(exp 6) and 100.0 x 10(exp 6). Results are also obtained for a transonic wing flow. For both 2-D and 3-D problems, the computational time of a well-tuned standard RK scheme is reduced at least a factor of four.
    Keywords: Aerodynamics
    Type: AIAA Paper 2006-3523 , 36th AIAA Fluid Dynamics Conference and Exhibit; Jun 05, 2006 - Jun 08, 2006; San Francisco, CA; United States
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  • 8
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    Apparent Transition Behavior of Widely-Used Turbulence Models (2006)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The Spalart-Allmaras and the Menter SST kappa-omega turbulence models are shown to have the undesirable characteristic that, for fully turbulent computations, a transition region can occur whose extent varies with grid density. Extremely fine two-dimensional grids over the front portion of an airfoil are used to demonstrate the effect. As the grid density is increased, the laminar region near the nose becomes larger. In the Spalart-Allmaras model this behavior is due to convergence to a laminar-behavior fixed point that occurs in practice when freestream turbulence is below some threshold. It is the result of a feature purposefully added to the original model in conjunction with a special trip function. This degenerate fixed point can also cause nonuniqueness regarding where transition initiates on a given grid. Consistent fully turbulent results can easily be achieved by either using a higher freestream turbulence level or by making a simple change to one of the model constants. Two-equation kappa-omega models, including the SST model, exhibit strong sensitivity to numerical resolution near the area where turbulence initiates. Thus, inconsistent apparent transition behavior with grid refinement in this case does not appear to stem from the presence of a degenerate fixed point. Rather, it is a fundamental property of the kappa-omega model itself, and is not easily remedied.
    Keywords: Aerodynamics
    Type: AIAA Paper 2006-3906 , 36th AIAA Fluid Dynamics Conference and Exhibit; Jun 05, 2006 - Jun 08, 2006; San Francisco, CA; United States
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  • 9
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    Pressure Gradient Effects on Hypersonic Cavity Flow Heating (2006)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The effect of a pressure gradient on the local heating disturbance of rectangular cavities tested at hypersonic freestream conditions has been globally assessed using the two-color phosphor thermography method. These experiments were conducted in the Langley 31-Inch Mach 10 Tunnel and were initiated in support of the Space Shuttle Return-To-Flight Program. Two blunted-nose test surface geometries were developed, including an expansion plate test surface with nearly constant negative pressure gradient and a flat plate surface with nearly zero pressure gradient. The test surface designs and flow characterizations were performed using two-dimensional laminar computational methods, while the experimental boundary layer state conditions were inferred using the measured heating distributions. Three-dimensional computational predictions of the entire model geometry were used as a check on the design process. Both open-flow and closed-flow cavities were tested on each test surface. The cavity design parameters and the test condition matrix were established using the computational predictions. Preliminary conclusions based on an analysis of only the cavity centerline data indicate that the presence of the pressure gradient did not alter the open cavity heating for laminar-entry/laminar-exit flows, but did raise the average floor heating for closed cavities. The results of these risk-reduction studies will be used to formulate a heating assessment of potential damage scenarios occurring during future Space Shuttle flights.
    Keywords: Aerodynamics
    Type: AIAA Paper 2006-0185 , 44th AIAA Aerospace Sciences Meeting and Exhibit; Jan 09, 2006 - Jan 12, 2006; Reno, NV; United States
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  • 10
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    Cascaded Perforates as One-Dimensional, Bulk Absorbers (2006)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Porous cell honeycomb liners for aircraft engine nacelles offer the possibility of exploiting extended reaction effects to improve liner attenuation bandwidth as generally attributed to the performance of bulk absorbers. This paper describes an analytical procedure, starting with an impedance prediction model for a single perforated plate, to estimate the bulk-absorber parameters for a cascade of such perforates - a first step to modeling a porous wall honeycomb structure. The objective is to build confidence in a lumped element impedance model, when applied to a uniformly-spaced set of porous plates to predict its .bulk. absorber properties. The model is based upon a modified version of the two-parameter flow resistance model of the form A + BV(sub inc), where A and B are physics-based, semi-empirical parameters that are adjusted to provide an optimum fit to a composite dataset from three plate porosities of 2.5, 5 and 10%. The composite dataset is achieved by reformulating the two-parameter flow resistance model into a .reduced pressure drop coefficient. dependency on perforate hole Reynolds number. The resulting impedance model is employed to calculate surface impedance spectra for N and 2N-layer perforate cascades. The well-known two-thickness method for experimental determination of bulk-absorber parameters is then applied to these .synthesized. data sets to predict the characteristic impedance and propagation constant for the perforate cascades. These results are then compared with experimental results reported in a companion paper.
    Keywords: Aerodynamics
    Type: AIAA Paper 2006-2402 , 12th AIAA/CEAS Aeroacoustics Conference; May 08, 2006 - May 10, 2006; Cambridge, MA; United States
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