NASA LogoNTRSNTRS - NASA Technical Reports ServerSearchmore_vertCollectionsAboutNewsHelpLogin Back to ResultsAn Experimental Investigation of a Wing-Fuselage Junction Model in the NASA Langley 14- by 22-Foot Subsonic TunnelCurrent turbulence models, such as those employed in Reynolds-averaged Navier-Stokes CFD, are unable to reliably predict the onset and extent of the three-dimensional separated flow that typically occurs in wing-fuselage junctions. To critically assess, as well as to improve upon, existing turbulence models, experimental validation-quality flow-field data in the junction region is needed. In this report, we present an overview of experimental measurements on a wing-fuselage junction model that addresses this need. The experimental measurements were performed in the NASA Langley 14- by 22-Foot Subsonic Tunnel. The model was a full-span wing-fuselage body that was configured with truncated DLR-F6 wings, both with and without leading-edge extensions at the wing root. The model was tested at a fixed chord Reynolds number of 2.4 million, and angles-of-attack ranging from -10 degrees to +10 degrees were considered. Flow-field measurements were performed with a pair of miniature laser Doppler velocimetry (LDV) probes that were housed inside the model and attached to three-axis traverse systems. One LDV probe was used to measure the separated flow field in the trailing-edge junction region. The other LDV probe was alternately used to measure the flow field in the leading-edge region of the wing and to measure the incoming fuselage boundary layer well upstream of the leading edge. Both LDV probes provided measurements from which all three mean velocity components, all six independent components of the Reynolds-stress tensor, and all ten independent components of the velocity triple products were calculated. In addition to the flow-field measurements, static and dynamic pressures were measured at selected locations on the wings and fuselage of the model, infrared imaging was used to characterize boundary-layer transition, oil-flow visualization was used to visualize the separated flow in the leading- and trailing-edge regions of the wing, and unsteady shear stress was measured at limited locations using capacitive shear-stress sensors. Sample results from the measurement techniques employed during the test are presented and discussed.Document ID20190027403Acquisition SourceLangley Research CenterDocument TypeTechnical Memorandum (TM)AuthorsKegerise, Michael A.*(NASA Langley Research Center Hampton, VA, United States)*Neuhart, Dan H.*(NASA Langley Research Center Hampton, VA, United States)*Date AcquiredJuly 17, 2019Publication DateJune 1, 2019Subject CategoryFluid Mechanics And ThermodynamicsReport/Patent NumberNASA/TM-2019-220286NF1676L-33264Funding Number(s)WBS: 109492.02.07.01.01Distribution LimitsPublicCopyrightWork of the US Gov. Public Use Permitted.Available DownloadsNameType20190027403.pdfSTIcloud_downloadcontent_copyvisibilityRelated RecordsThere are no records associated with this record.visibility_offNo Preview Available
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