SCIENTIFIC AND TECHNICAL AEROSPACE REPORTS
A Biweekly Publication of the National Aeronautics and Space Administration
VOLUME 43, ISSUE 19 - SEPTEMBER 23, 2005
02 AERODYNAMICS
Includes aerodynamics of flight vehicles, test bodies, airframe components and combinations, wings, and control surfaces.
Also includes aerodynamics of rotors, stators, fans, and other elements of turbomachinery.
For related information see also 34 Fluid Mechanics and Thermodynamics.
20050210134 NASA Langley Research Center, Hampton, VA, USA
Real-Time Aerodynamic Flow and Data Visualization in an Interactive Virtual Environment
Schwartz, Richard J.; Fleming, Gary A.; [2005]; 6 pp.; In English; 2005 IEEE Instrumentation and Measurement Technology Conference, 17-19 May 2005, Ottawa, Ontario, Canada Contract(s)/Grant(s): 23-090-80-40 Report No.(s): Paper IMTC-5246; No Copyright; Avail: CASI; A02, Hardcopy
Significant advances have been made to non-intrusive flow field diagnostics in the past decade. Camera based techniques are now capable of determining physical qualities such as surface deformation, surface pressure and temperature, flow velocities, and molecular species concentration. In each case, extracting the pertinent information from the large volume of acquired data requires powerful and efficient data visualization tools. The additional requirement for real time visualization is fueled by an increased emphasis on minimizing test time in expensive facilities. This paper will address a capability titled LiveView3D, which is the first step in the development phase of an in depth, real time data visualization and analysis tool for use in aerospace testing facilities. Author
Aerodynamic Characteristics; Real Time Operation; Flow Visualization; Metrology; Deformation; Flow Velocity; Molecular Gases; Scientific Visualization
20050210229 NASA Glenn Research Center, Cleveland, OH, USA
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Optimal disturbances in boundary layers subject to streamwise pressure gradient
Ashpis, David E.; Tumin, Anatoli; [2003]; 1 pp.; In English; American Physical Society, Div. of Fluid Dynamics, 55th Annual Meeting, 24-26 Nov. 2002, Dallas, TX, USA Contract(s)/Grant(s): NCC3-991; Copyright; Avail: Other Sources; Abstract Only
An analysis of the optimal non-modal growth of perturbations in a boundary layer in the presence of a streamwise pressure gradient is presented. The analysis is based on PSE equations for an incompressible fluid. Examples with Falkner-Scan profiles indicate that a favorable pressure gradient decreases the non-modal growth, while an unfavorable pressure gradient leads to an increase of the amplification. It is suggested that the transient growth mechanism be utilized to choose optimal parameters of tripping elements on a low-pressure turbine (LPT) airfoil. As an example, a boundary layer flow with a streamwise pressure gradient corresponding to the pressure distribution over a LPT airfoil is considered. It is shown that there is an optimal spacing of the tripping elements and that the transient growth effect depends on the starting point. Author
Boundary Layer Flow; Boundary Layers; Pressure Gradients; Pressure Distribution; Airfoils; Perturbation; Incompressible Fluids
20050212113 NASA Langley Research Center, Hampton, VA, USA
Shock Interaction Control for Scramjet Cowl Leading Edges
Albertson, Cindy W.; Venkat, Venki, S.; [2005]; 19 pp.; In English; 13th AIAA/CIRA International Space Planes and Hypersonic Systems Technologies Conference, 16-20 May 2005, Capua, Italy Contract(s)/Grant(s): 23-745-30-20 Report No.(s): AIAA Paper 2005-3289; Copyright; Avail: CASI; A03, Hardcopy
An experimental study was conducted to qualitatively determine the effectiveness of stagnation-region gas injection in protecting a scramjet cowl leading edge from the intense heating produced by Type III and Type IV shock interactions. The model consisted of a two-dimensional leading edge, representative of that of a scramjet cowl. Tests were conducted at a nominal freestream Mach number of 6. Gaseous nitrogen was supersonically injected through the leading-edge nozzles at various mass flux ratios and with the model pitched at angles of 0deg and -20deg relative to the freestream flow. Qualitative data, in the form of focusing and conventional schlieren images, were obtained of the shock interaction patterns. Results indicate that large shock displacements can be achieved and both the Type III and IV interactions can be altered such that the interaction does not impinge on the leading edge surface. Author
Shock Wave Interaction; Leading Edges; Stagnation Point; Gas Injection; Supersonic Combustion Ramjet Engines
20050212277 Air Force Inst. of Tech., Wright-Patterson AFB, OH USA
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Experimental Study of the Subsonic Aerodynamics of a Blended Wing Body Air Vehicle with a Focus on Rapid Technology Assessment
Gebbie, David A.; Mar. 2005; 157 pp.; In English; Original contains color illustrations Report No.(s): AD-A436457; AFIT/GAE/ENY/05-M09; No Copyright; Avail: CASI; A08, Hardcopy
The subsonic aerodynamic performance of a blended wing body aircraft constructed using selective laser sintering was assessed in the AFIT low-speed wind tunnel. The scaled-down model of a strike tanker aircraft consisted of a shaped fuselage and sweptback wings. The Reynolds number, based on mean wing chord, during testing was on the order of 105 while the Mach number ranged from 0.10 to 0.20. The model evaluation and analysis process included force and moment measurements acquired from a wind tunnel balance, pressure data measured with 8 taps located on the model’s upper surface, a comparison to computational fluid dynamics (CFD) solutions acquired in a parallel study conducted by AFRL/ VAAC, and global pressure sensitive paint (PSP) measurements. Paint measurements were compared to pressure tap data to ensure their accuracy while lift and drag coefficients, as well as pitching and rolling moments were examined to determine performance characteristics, including stability attributes and aircraft stall. One of the most interesting results was the striking difference in the force and moment measurements before and after the paint was applied to the surface. The average surface roughness, Ra, was measured with a profilometer and was found to have increased from approximately 0.3µm to 0.7µm when the paint was applied. When traditional 2-D boundary layer approaches to assessing the effect of roughness, the 0.7µm value falls well below the threshold at which one would anticipate roughness to have any effect. There is support in archival literature for the notion that roughness effects are more pronounced in a 3-D boundary layer, and the pitching moment data and the PSP data indicate that the for the painted model, there is a gradual onset of wing stall marching inward from the wingtips toward the body. DTIC
Aerodynamic Characteristics; Aerodynamics; Blended-Wing-Body Configurations; Technology Assessment
20050212345 Consolidated Analysis Centers, Inc., San Diego, CA USA
System Dynamics Aviation Readiness Modeling Demonstration
McDevitt, Michael E.; Aug. 2005; 64 pp.; In English; Original contains color illustrations Contract(s)/Grant(s): N00014-04-F-0265 Report No.(s): AD-A436605; No Copyright; Avail: CASI; A04, Hardcopy
A proof-of-concept demonstration of System Dynamics modeling was developed to determine the relative merit of the approach and then an experiment compared the results of the System Dynamics model with a traditional linear regression readiness model. This report documents the proof of concept model development and the experiment’s results. An aviation readiness production model was formulated using system dynamics. The model was developed over a period of six months with participation from subject matter experts at Commander Naval Air Force and Commander Strike Fighter Wing Pacific. The model incorporates monthly flight hour execution taking into consideration aircraft mission capable rates and the location of a squadron relative to the Fleet Response Plan (FRP) training cycle to generate Primary Mission Area points in strike warfare that are accumulated to generate an Attack Index. Concurrently, a renewed Strike Warfare Proficiency (Strike PRO) analysis was conducted for comparison. Twenty-four Carrier Air Wing (CVW) events at NSAWC Fallon, NV were used to benchmark performance. The results of the Strike PRO algorithm and the results of the System Dynamics Model were each compared against the performance metric. Source code for the System Dynamics Model is included as an appendix to the report. DTIC
Aeronautics; Maintainability; Warfare
20050214045 NASA Glenn Research Center, Cleveland, OH, USA
Efficient Low Noise Fan Overview and Preliminary Aerodynamic Design
NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts; [2001], pp. 25; In English; See also 20050214031; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document
With support from the Ultra-Efficient En,aine Technology (UEET) and Quiet Aircraft Technology (QAT) programs, an efficient low noise fan will be designed and tested to investigate aerodynamic and acoustic performance effects using trailing edge blowing (TEB). The fan uses TEB to fill the momentum deficit behind the fan rotor. This presentation will include an overview of the plan to design, build, and test the f a n in the NASA Glenn Research Center 9- by 15-Foot Low-Speed Wind Tunnel. In addition, the preliminary aerodynamic design will be presented. Author
Aerodynamic Characteristics; Trailing Edges; Blowing; Aerodynamics; Aeroacoustics; Low Speed; Low Noise
20050214050 NASA Langley Research Center, Hampton, VA, USA
Variable Radius Nacelle Studies
McGowan, David M.; NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts; [2001], pp. 39; In English; See also 20050214031; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document
An overview of the active shape control for a variable radius nacelle leading edge program is presented. The current technical plan and schedule will be discussed. Results from the structural shape change of curved plates demonstration will be presented, as well as the NASA LaRC concept for a variable radius nacelle leading edge. Results of a Boeing systems integration study of this concept will be discussed briefly. The status of the sensors, actuators, and computational design tools tasks will also be presented. Author
Nacelles; Radii; Systems Integration; Active Control; Actuators
20050214061 NASA Langley Research Center, Hampton, VA, USA
CFD Design for Bypass Ratio 15 Nacelle Integration
Milholen, William E., II; NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts; [2001], pp. 41; In English; See also 20050214031; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document
A computational study is being conducted to evaluate the installation effects of ultra-high bypass ratio (BPR) nacelles on conventional twin-engine transonic transport aircraft. An unstructured Navier-Stokes flow solver, USM3D, is being utilized for the study. The results have been compared to wind tunnel data obtained in the NASA LaRC 16-Foot Transonic Tunnel, for nacelle BPRs of nine and twelve. The USM3D flow solver was found to adequately predict the flows of interest, and has subsequently been used to analyze the installation effects of a theoretical BPR-15 nacelle. In addition, a design code is being used in conjunction with USM3D to redesign the wing in the presence of the BPR-15 nacelle. The preliminary design results will be presented. Author
Nacelles; Computational Fluid Dynamics; Bypass Ratio; Supersonic Aircraft; Wings
20050214078 Army Research Lab., Cleveland, OH, USA
UEET Compressor
Larosiliere, Louis M.; NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts; [2001], pp. 26; In English; See also 20050214031; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document
Turbocompression technology has been advanced continuously by higher work capacity per stage owing to increases in rotor speed, aerodynamic loading, and throughflow Mach numbers. Using sophisticated diagnostic tools involving CFD and measurement techniques, more suitable blade shapes, with relatively low losses at higher diffusion and Mach number levels, have been deployed. The quest for further aerodynamic performance advancements is becoming progressively more difficult due to a dwindling residue of losses. A vital issue is whether to advance by refining and extending well-proven concepts, perhaps in the face of diminishing returns, or by changing to something more unpredictable yet inviting. A program is being conducted by the Glenn Research center to study and develop advanced design concepts that will enable compact, high-efficiency, wide-operability compressors. This is a direct response to the need for further performance gains from current turbomachinery systems. A combination of evolutionary and revolutionary approaches to technology development has been adopted. The evolutionary approach employs advancements in simulation techniques to refine traditional design concepts in a bid for higher efficiencies at increased aerodynamic loading levels, whereas the revolutionary approach attempts to explore unconventional concepts and paradigms for increased pressure ratio, higher efficiencies, and wider operability. This presentation starts by describing the level of technical advancement being sought and quantifying the advancement in terms of basic aerodynamic technology elements of current design systems. An outline of the preliminary aerodynamic design of a 4-stage core compressor configuration having the potential to integrate and demonstrate these advancements is given. The front two stages of the 4-stage configuration were further developed using the best available CFD tools. A three-dimensional viscous inverse design method was used to define transonic rotor blading, and the design considerations are discussed. This serves to illustrate achievable performance using an evolutionary approach buttressed by advanced turbomachinery CFD. Author
Compressors; Computational Fluid Dynamics; Turbomachinery; Aerodynamic Characteristics; Aerodynamic Loads; Aerodynamics; Pressure Ratio
20050214082 NASA Langley Research Center, Hampton, VA, USA
CFD Modeling for Active Flow Control
Buning, Pieter G.; NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts; [2001], pp. 37; In English; See also 20050214031; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document
This presentation describes current work under UEET Active Flow Control CFD Research Tool Development. The goal of this work is to develop computational tools for inlet active flow control design. This year s objectives were to perform CFD simulations of fully gridded vane vortex generators, micro-vortex genera- tors, and synthetic jets, and to compare flowfield results with wind tunnel tests of simple geometries with flow control devices. Comparisons are shown for a single micro-vortex generator on a flat plate, and for flow over an expansion ramp with sidewall effects. Vortex core location, pressure gradient and oil flow patterns are compared between experiment and computation. This work lays the groundwork for evaluating simplified modeling of arrays of devices, and provides the opportunity to test simple flow control device/sensor/control loop interaction. Author
Active Control; Computational Fluid Dynamics; Flow Distribution; Flow Visualization; Inlet Flow; Vortex Generators; Pressure Gradients
20050214083 NASA Langley Research Center, Hampton, VA, USA
Active Flow Control For Inlets
Gorton, Susan A.; NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) Program: Agenda and Abstracts; [2001], pp. 35; In English; See also 20050214031; No Copyright; Abstract Only; Available from CASI only as part of the entire parent document
This presentation describes the progress to date of the Small-Scale Demonstration for the Active Flow Control element of the Propulsion Airframe Integration Project. The goal of this work package is to demonstrate at small scale the ability to improve pressure recovery and distortion in an S-inlet with boundary layer ingestion representative of a Blended Wing Body (BWB) configuration. The effectiveness of several active and passive devices to control flow in an adverse pressure gradient with secondary flows present was evaluated in the Langley 15-Inch Low-Turbulence Tunnel. In this study, passive microvanes, microbumps, and piezoelectric synthetic jets were evaluated for their flow control characteristics using surface static pressures, flow visualization, and 3D Stereo Digital Particle Image Velocimetry. The microvanes imparted a higher level of vorticity to the flow than any of the other devices tested. Alternative actuator concepts are being pursued to support the Small-Scale Demonstration Level 1 milestone in FY03. Author
Active Control; Flow Distribution; Flow Visualization; Aerodynamic Configurations; Ingestion (Engines); Low Turbulence; Engine Airframe Integration; Blended-Wing-Body Configurations.
Source: NASA.
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