AIAA-98-0701REYNOLDS-AVERAGED NAVIER-STOKES SIMULATIONS OF
TWO PARTIAL-SPAN FLAP WING EXPERIMENTS
M.A. Takallu*Lockheed Martin Engineering & Sciences Company
Hampton, VA
Kelly R. Laflin*National Research Council
Hampton, VA
Abstract Introduction
Structured Reynolds Averaged Navier-Stokessimulations of two partial-span flap wing experimentswere performed. The high-lift aerodynamic andaeroacoustic wind-tunnel experiments were conductedat both the NASA Ames 7- by 10-Foot Wind Tunnel andat the NASA Langley Quiet Flow Facility. The purposeof these tests was to accurately document the acousticand aerodynamic characteristics associated with theprinciple airframe noise sources, including flap side-edge noise. Specific measurements were taken that canbe used to validate analytic and computational modelsof the noise sources and associated aerodynamics forconfigurations and conditions approximating flight fortransport aircraft. The numerical results are used to bothcalibrate a widely used CFD code, CFL3D, and toobtain details of flap side-edge flow features notdiscernible from experimental observations. Bothexperimental set-ups were numerically modeled byusing multiple block structured grids. Variousturbulence models, grid block-interface interactionmethods and grid topologies were implemented.Numerical results of both simulations are in excellentagreement with experimental measurements and flow-visualization observations. The flow field in the flap-edge region was adequately resolved to discern somecrucial information about the flow physics and tosubstantiate the merger of the two vortical structures. Asa result of these investigations, airframe noise modelershave proposed various simplified models which use theresults obtained from the steady-state computations asinput.
*Supervisor, Aerodynamics SectionNASA Langley Program Office. Associate Fellow AIAA.m.a.takallu @ larc.nasa.gov^Resident Research Associate, Aero and Gas DynamicsDivision, NASA Langley Research Center. [email protected]
An aerodynamic challenge related to the design ofsubsonic transport aircraft is the determination andreduction of radiated airframe noise. Research1
indicates that, as aircraft engines become quieter in thenear future, a significant contributor to airframe noise ofan aircraft on approach will be the high-lift system.Within the complex three-dimensional flow field of ahigh-lift system, the side-edge region of a partial-spanflap is suspected to produce the dominant noise source.2The noise associated with this region is believed to bedue to the formation and shedding of a flap side-edgevortex and its interaction with the free shear layer.Although the general mechanisms of flap side-edgevortex formation are well known, details of the flowfield physics and associated unsteadiness required fornoise source generation in the flap side-edge region arenot completely understood. Knowledge of these details,as well as other details of high-lift flow fields, arecrucial to the design of improved high-lift systems withreduced airframe noise levels. Required flow fielddetails, that are not discernible from experimentalmeasurements or observations, will likely be determinedthrough the use of computational fluid dynamics (CFD).
Noise generation is an unsteady phenomenon, butunsteady viscous three-dimensional flow fieldcomputations are both time consuming and expensive.For this reason, airframe noise modelers have proposedvarious simplified models which use the results obtainedfrom steady-state Reynolds Averaged Navier-Stokes(RANS) computations as input for noise sourcemodels.3"4 However, before the RANS results can beconfidently used in this capacity, they must be calibratedagainst reliable test data.
In this paper, experimental results of two partial-span flap wing experiments are used to calibrate the useof a RANS flow solver for predicting high-lift flowfields and providing the details required by airframenoise models. Although, Jones et al have demonstrated
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the ability of such a flow solver to simulate flow fieldsabout two-dimensional and three-dimensional multi-element high-lift configurations, it is not known ifRANS flow solvers are capable of providing the small-scale aerodynamic details associated with airframenoise sources.6 In the present work, structured RANScomputations were used as a modern analysis tool tostudy two high-lift aerodynamic and aeroacoustic wind-tunnel experiments in order to both calibrate a widelyused CFD code and to simulate some of the flowfeatures not captured by the experiments. Therefore, abrief discussion of the two NASA experiments isfollowed by a description of the numerical strategy.Computed results, that concentrate on the flap side-edgeregion, will be presented in detail and compared withexperimental data.
Experimental Investigations
A series of partial-span flap wing wind-tunnelexperiments were conducted at both the NASA AmesResearch Center 7- by 10-Foot Wind Tunnel7 (7x10)and at the NASA Langley Quiet Flow Facility8 (QFF).The purpose of these tests was to accurately documentthe acoustic and aerodynamic characteristics associatedwith the principle airframe noise sources, including flapside-edge noise. Specific measurements were taken thatcan be used to validate analytic and computationalmodels of the noise sources and associatedaerodynamics for configurations and conditionsapproximating flight for transport aircraft.
7x10 Experiment
The NASA Ames 7x10 test provided a data base ofaerodynamic forces and moments with accompanyingwing pressure distributions (750 pressure taps) for apartial-span flap wing model . A sketch of the modelset-up is shown in Figure 1. The model was designed sothat a constant-chord rectangular wing with a NACA632-215 Mod B airfoil cross-section resulted when theflap was retracted; thereby providing a reference "two-dimensional" flow to which the three-dimensional flowsof deployed-flap configurations could be compared. Thewing chord and span are cw=30 inches and 2.5 cw,respectively. The half-span flap has a chord of 0.3cw.Measurements were taken at angles of attack rangingfrom 0 to 15 degrees with flap deflections of 29 and 39degrees. Microphone array and 7-hole probe near wakesurveys where also conducted as part of the 7x10 test.
OFF ExperimentIn an effort complementary to the 7x10 test,
aerodynamic and noise issues were investigated at theQFF using a 16 inch chord model of similar geometry tothat of the 7x10 test model but with an aspect ratio of2.25. In the QFF (Figure 2), jet flow enters verticallythrough the floor of an anechoic chamber via aconverging inlet nozzle which accelerates the jet flow tothe desired entrance conditions. The partial-span flapmodel (Figure 3) is mounted between two rectangular-shaped end plates and positioned in the vertical jet(Figure 4). The jet is constrained in the spanwisedirection of the wing by the two end plates but isotherwise allowed to expand. Flow exits through aconverging exhaust nozzle positioned at the ceiling ofthe large anechoic chamber. Because of the expansion ofthe jet flow between the two end plates, wing loadingmeasurements were lower than those of the 7x10 test forsimilar angles of attack and flap-deflection angles.Therefore, the angles of attack of the QFF experimentwere adjusted so that the wing loadings more closelymatched those obtained in the Ames test. Microphonearrays, directional microphones, 5-hole probe wakesurveys, pressure sensitive paint (PSP) applications, andlaser light sheet flow visualization were employed aspart of the QFF test.
Numerical Approach
All computations were performed on the Cray C-90and Cray Y-MP super-computers of the NationalAerodynamic Simulation (NAS) facility located at theNASA Ames Research Center.
Flow SolverThe RANS flow solver CFL3D9 was used to
simulate both the 7x10 and the QFF experiments. Thiscode uses the unsteady, three-dimensional,compressible, thin-layer Navier-Stokes equations tomodel the flow on body-fitted structured grids withmultiple-block topologies. It is a semi-discrete cell-centered finite-volume scheme, that can make use ofgrid sequencing, local time-stepping, and multigriddingto accelerate convergence to a steady state; all threetechniques were utilized for all computations presentedherein. CFL3D allows both CO-continuous (one-to-onematching at block interfaces) and patched blockinterfaces; both of these were used in this study.
CFL3D advances the solution using an implicitapproximate factorization (AF) scheme. A blocktridiagonal inversion results for each AF sweep by using
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first-order accurate upwind-biased differencing toapproximate the implicit spatial derivatives. However,because steady state solutions are considered, flux-difference splitting (FDS) can be used in conjunctionwith a diagonal scheme to require only a scalartridiagonal inversion for each AF sweep. Thissimplification reduces both the required memory andCPU time without decreasing the accuracy of thecomputations. CFL3D can also employ flux-vectorsplitting, but this option was not used to perform thecomputations presented here. Explicit spatial derivativesare approximated using third-order accurate upwind-biased differencing for the inviscid terms and second-order accurate centered differencing for the viscousterms. Fully turbulent solutions were obtained using theSpalart-Allmaras10 one-equation turbulence models.The Spalart-Allmaras turbulence model is oftenemployed when high-lift configurations are consideredbecause it has been performing adequately for multiplesolid surfaces and wake/boundary layer interactions. Inaddition, it seems to produce more accurate solutions inreverse flow regions, as compared with other turbulencemodels.
For patched zones, data transfer between zones isaccomplished by linear interpolation in thecomputational coordinate system. The interpolationcoefficients required for CFL3D are obtained as a pre-processing step using the RONNIE preprocessor.
Gridding StrategySince these computed high-lift flow fields are to be
directly compared with experimental measurements andobservations, it is logical, as demonstrated by Cao etal11 that the wind tunnel walls need to be modeled aspart of the numerical simulation. The multiplecomponents of a high-lift system require multiple gridblocks. As it was mentioned earlier, CFL3D allows bothCO-continuous (one-to-one matching at blockinterfaces) and patched block interfaces; both of thesewere used in this study. The addition of the rectangularwalls adds some difficulty to a CO-continuous gridconstruction, especially when the test facility geometryis complex, such as the geometry of the QFF. Analternative to CO-continuous grid construction is to usepatched block interfaces. By using patched grids, thegrid construction procedure is simplified and therequired number of grid points can generally bereduced. In order to substantiate the use of patched gridinterfaces in future studies, computations whereperformed on both CO-continuous and patched grids andthe resulting solutions where compared to theexperimental data.
Numerical Simulation of 7x10 ExperimentThe partial-span flap wing tested in the NASA
Ames 7x10 was modeled using both a 16-block gridwith CO-continuous interfaces (Figure 5) and a 19-blockgrid with mixed CO-continuous and patched interfaces(Figure 6). The 16-block and 19-block grids have a totalof approximately 2.3 and 2.6 million points,respectively. In each grid, the constant-area rectangulartest section was extended 15 chord lengths upstream anddownstream from the wing. Wing angles of attack of 4,8, and 10 degrees with a flap deflection of 29 degreeswere considered for the numerical simulations. TheReynolds number and Mach number used for thesimulations were 3.6x10 and 0.2, respectively;matching the 7x10 test conditions. The computationalgrids used viscous grid spacing suitable for turbulentboundary-layer computations at all wing surfaces. Inaddition, a high concentration of grid-points was used toresolve the flow field about the flap side-edge. Inviscidgrid-points spacing was used at all solid surfaces, apartfrom those of the wing components, to limit the totalnumber of grid points and, thus, the required memoryand CPU time of the simulations. This methodology wassuccessfully used by other researchers to compute theabove set-up both with structured6 and unstructuredgrids.12
Numerical Simulation of OFF Experiment
When a constant area rectangular-shaped testsection is considered, the grid generation, boundary-condition specification and computations for anumerical simulation are relatively simple; this is thecase for the numerical simulation of the 7x10experiment. However, numerical simulations of the QFFexperiment present several challenges that have not beenpreviously addressed for the calculation of three-dimensional, viscous, high-lift flow simulations. Thesechallenges include the numerical simulation of a high-lift device in a semi-constrained jet flow, the presence oflarge-scale circulatory flow throughout a large testchamber, and the existence of large regions of near-stagnant flow which is known to cause convergence andaccuracy difficulties for compressible codes. 13A seriesof grid alterations, such as grid refinements around theend-plates and inside the area representing the anechoicroom, were performed to meet these computationalchallenges.
The QFF experimental set-up was modeled using a21-blocked grid with CO-continuous interfaces havingapproximately 3.2 million grid points (Figure 7). Asillustrated in Figure 7a and 7b, the first 12 grid blocks
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used in the inner-blocks of the 7x10 grid were utilizedalso as the inner-blocks of the QFF grid. These blockssimulate the wing/flap system and the space between thesupporting end-plates. Figure 7c shows the 21-blockQFF grid. For ease of illustration, each block is coloreddifferently.
The chord Reynold's number and average inlet-throat Mach number are l.SxlO6 and 0.2, respectively.To achieve the correct throat Mach number, the pressureratios at the inlet nozzle and exhaust exit boundaries areadjusted iteratively.
Even though the numerical simulation of the QFFexperiments included computations of a three elementhigh-lift configuration composed of a wing, partial-spanned flap and a full-span leading edge slat, theresults presented in the following will be related to thetwo-element configuration only14. In addition toCFL3D, another widely used Langley RANS code,TLNS3D-MB15 was also utilized to simulate the QFFexperiment.14 Both grid methodologies, grid blockswith CO-continuous block boundaries and patched blockinterfaces, were utilized. Also, various angles-of-attackand flap deflections were computed. However, forbrevity only results of the computations using CFL3Dcode and the 21 block grid with CO-continuous gridblock boundaries for the angle-of-attack of 16 degreesand 29 degree flap-deflection are presented.
ResultsAll numerical simulations of the experimental test
cases have been completed and show excellentagreement with the experimental data. Various grid sizesand topologies along with various turbulence modelswere computed. A wide range of wing angle of attackand flap deflections were also analyzed. Some sampleresults are presented. These results (Figures 8-18) willfocus on the validation of the converged numericalsolutions through comparisons with experimental dataand on the analysis of the three-dimensionality of thehigh-lift flow field. The analysis will concentrate on theflow field in the flap side-edge region.
7x10 SimulationFigure 8 gives the convergence history for the 10
degree angle-of-attack. Solution convergence of thisturbulent-viscous simulation is reached within 3000iterations, with a four-order drop in the residualmagnitude. Both multigridding and mesh sequencingwere used to increase the solution convergence rate.
Figure 9 compares the computational andexperimental surface pressure distributions at fourspanwise locations. Two of the spanwise stations are inthe flapped region of the wing and the other two are inthe non-flapped segment of the wing. As is shown, thecomputations are in good agreement with theexperimental data. The saddle type sectional pressureprofile near the flap side-edge is an indication of theexistence of some vorticity generation in this region. Infact a detailed analysis of the flow field at the flap sideedge reveals the (see Figures 10 and 11) merging of astrong vortex at the lower (pressure) side of the flapedge with a weaker vortex from the top (suction) side.Figure 10 shows the surface pressure contour and theextent of the flow disturbance due to the flap on bothflapped region and non-flapped segment of the wing.
Figure 12 presents flap side-edge total pressurecontours and grid topologies for both the 16-block gridwith CO-continuous grid boundaries and the 19-blockgrid with patched block boundaries. Despite thedifferences in grid topologies and grid-pointconcentrations, both numerical solutions produceagreeable results. This fact indicates some grid-insensitivity of the present numerical solutions in theflap side-edge region.
OFF SimulationAs it is shown in Figure 13a, the Mach contour on a
cutting plane through the mid-span of the tunnelillustrates the turning of the flow towards the flaptrailing edge. The turning phenomenon is much strongerfor the QFF set-up than those observed for the 7x10simulation and could be explained by the lack ofconfining walls on the top and bottom of the wing set-up. This can affect the down stream location andshedding of the merged flap side-edge vortex. Figure13b shows a portion of the QFF experimentconfiguration and grid geometry downstream of thenozzle looking onto the suction side of the wing.Streamlines in the figure illustrate the extent of thedownwash that results from the semi-constrained jet.
Figures 14 and 15 show pressure contours on theflap side-edge and in several planar cuts normal to thestream direction. Note that the flap side-edge surfacepressure patterns are similar to those of the 7x10simulation (Figure 10). This is another indication thatthe QFF simulation can adequately capture some of theflow phenomena of particular interest to the acousticscommunity. As mentioned earlier, the appearance andmerging of the vortices at the flap side-edge has beensuspected to be one of the major sources of airframe
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Figure 15 gives planar cuts normal to the flap side-edge at various chord locations. The positions of thesecuts are indicated in Figurel5f. The pressure contourlines show details of developing vortical structures andindicates that the flap side-edge flow of the QFFsimulation is well resolved.
Both the NASA Ames 7x10 and the NASA LangleyQFF experiment simulations show the development andmerger of a weak and a strong vortex in the flap side-edge region. In each case, a strong vortex develops as aresult of flow separation from the lower surface at theflap edge (Figure 15f). Flow reattachment associatedwith this vortex occurs on the flap edge. This attachedflow separates at the upper surface of the flap and formsa weaker vortex along the flap upper surface (Figure15b). Eventually, the reattachment position of the strongvortex exceeds the confines of the flap edge (Figure 15c)and the two vortices merge into a single vortex (Figure15d). The vortex detatches at the flap trailing edge(Figure 15e).
These computed results are consistent with laser-light sheet, surface pressure measurements16 andpressure-sensitive paint (PSP) flow visualizationobservations17 conducted during the QFF experiment.Comparisons of the computed results with experimental(PSP) observations on the top and flap side-edge areshown in figure 16 and figure 17, respectively.
Figure 18 compares the computational andexperimental surface pressure distributions at sixspanwise locations. Four of the spanwise stations are inthe flapped region of the wing and the other two are inthe non-flapped region of the wing. As is shown, thecomputations are in good agreement with theexperimental data. Figures 18c and 18d show thesignature of the flap side-edge vortex on these twosection pressure profiles. An in-depth look at the areasenclosed by the pressure distributions indicates theextent of the span loading for the set-up discussedabove.
Summary
Numerical simulations of two partial-span flapwing experiments have been successfully completed.These three-dimensional steady-state Reynolds-averaged Navier-Stokes simulations involved the
modeling of two partial-span high-lift wings, includingthe wind tunnel walls, mounted in two different testfacilities: the NASA Ames 7- by 10-Foot Wind Tunneland the NASA Langley Quiet Flow Facility. Numericalsimulation of the QFF experiment which consisted ofmodeling of the high-lift wing, end-plates, inlet nozzle,anechoic room and exhaust nozzle, presented severalchallenges that were successfully addressed for the firsttime in this work. Various turbulence models, gridblock-interface interaction methods and grid topologieswere implemented. The flow field in the flap side-edgeregion was adequately resolved to discern some crucialinformation about the flow physics and to substantiatethe merger of the two vortical structures. Numericalresults of both simulations are in excellent agreementwith experimental measurements and flow-visualizationobservations.
Acknowledgments
The work reported in this paper was performedunder NASA Contract NAS1-96014. The authors wouldlike to thank the many researchers who contributedideas and assistance to this work, particularly membersof the Subsonic Aerodynamics Branch and members ofthe Airframe Noise Team at the NASA LangleyResearch Center. The authors would also like to thankMr. Ken Jones of Subsonic Aerodynamics Branch, Dr.Michele Macaraeg and Dr. Robert Biedron of FluidMechanics and Acoustics Branch and Mr. Mike Wieseof Computer Sciences Corporation.
References
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2Revel, J.D., "The Calculation of Aerodynamic NoiseGenerated by Large Aircraft at Landing Approach,"Paper JJ9, 87th Meeting, Acoustical Society ofAmerica, April 1974.
3Khorrami, M.R., Singer, B.A. and Takallu, M.A.,"Analysis of Flap Side-Edge Flow Field forIdentification and Modeling of Possible Noise Sources,"SAE Paper No. 971917, presented at the 1997 SAENoise and Vibration Conference and Exposition, GrandTraverse, MI, May 1997.
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4Steett, C.L. "DNS of Flap-Edge FlowfieldInstabilities," proceedings of the Airframe NoiseWorkshop, Long Beach, California, 11-12 June 1997.
5Jones, K.M., Biedron, R.T., and Whitlock, M.,"Application of a Navier-Stokes Solver to the Analysisof Multielement Airfoils and Wings Using MultizonalGrid Techniques," AIAA Paper 95-1855,1995.
6Mathias, D.L., Roth, K.R., Ross, J.C., Rogers, S.E.,and Cummings, R.M., "Navier-Stokes Analysis of theFlow About a Flap Edge," AIAA Paper 95-0185,1995.
7Storms, B.L., Takahashi, T.T., and Ross, J.C.,"Aerodynamic Influence of a Finite-Span Flap on aSimple Wing," SAE paper no. 95-1977, presented atSAE Aerotech '95, Los Angeles, CA, September 1995.
8Hubbart, H.H. and Manning, J.C., "AeroacousticResearch Facilities at NASA Langley Research Center'"NASA TM 84585, March 1983.
9Krist, S.L., Biedron, R.T., and Rumsey, C.L., "CFL3DUser's Manual (Version 5.0)," Aerodynamic andAcoustic Methods Branch, NASA Langley ResearchCenter, 1996.
10Spalart, P. and Allmaras, S., "A One-EquationTurbulence Model for Aerodynamic Flows," AIAA 92-0439,1992.
nCao, H., Kusunose, K., Spalar, P., Ishimitsu, K.,Rogers, S., and McGhee, R., "Study of Wind TunnelWall Interference for Multi-Element Airfoils Using aNavier-Stokes Code," AIAA Paper 94-1993, June 1994.
12Anderson, W.K., Rausch, R.D., and Bonhaus, D.L.,"Implicit/Multigrid Algorithms for IncompressibleTurbulent Flows on Unstructured Grids," AIAA 95-1740, 1995.
13Volpe, G., "On the Use and Accuracy of CompressibleFlow Codes at Low Mach Numbers," AIAA 91-1662,1991.
14Takallu, M.A.,"Reynolds-Averaged Navier-StokesSimulations of Two Partial-Span flap WingExperiments", proceedings of the Airframe NoiseWorkshop, Long Beach, California, 11-12 June 1997.
15Vatsa, V.N., Sanetrik, M.D., Parlette, E.B., Eiseman,P., and Cheng, Z., "Multi-block Structured GridApproach for Solving Flows over Complex
Aerodynamic Configurations," AIAA 94-0655,1994.
16Meadows, K.R., Brooks, T.F., Humphreys, W.M.,Hunter, W.H., and Gerhold, C.H., "AeroacousticMeasurements of a Wing-Flap Configurations," AIAA97-1595, 1997.
17Radeztsky, Jr., R.R., "Experimental Measurements ofthe Flap-Edge Flowfield," proceedings of the AirframeNoise Workshop, Long Beach, California, 11-12 June1997.
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Figure 5: Grid blocks with CO continuous block Figure 6: Grid blocks with patched block interfaces forinterfaces for 16-block "7x10 grid". 19-block "7x10 grid".
.s" ..--'
_______ / .,..,...;/;._.... / /™
Figure 5b: Close-up view of the flap side-edge grid with Figure 6b: Close-up view of the flap side edge grid withCO continuous block interfaces. patched block interfaces.
8American Institute of Aeronautics and Astronautics
4) QBE M X-1.10C 5) one It *-l.ltC t) EMt> *e X-O.MSC-l.LK
Figure 15: Pressure contour lines on various cuttingplanes perpendicular to flap side-edge for the "QFFsimulation", M=0.2, Re=1.8xl06, a=16°, 8^29°.
Figure 13b: Close-up view of pressure contour andstream lines near the jet nozzle for the "QFFsimulation", M=0.2, Re=1.8xl06, o=16", 8f=29".
Figure 16: Comparison of PSP and CFD pressurecoefficients on the suction side of the wing for the "QFFsimulation", M=0.2, Re=1.8xl06, a=16°, 8f=29°.
Figure 14: Close-up view of pressure contours in planarcuts perpendicular to the flap side-edge for the "QFFsimulation", M=0.2, Re=1.8xl06, o=16°, 8f=29°. Figure 17: Comparison of PSP and CFD pressure
coefficients on the flap side-edge for the "QFFsimulation", M=0.2, Re=1.8xl06, o=16", 8^=29°.
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