Navier-Stokes Analysis of a High Wing Transport High-Lift ... · PDF fileNAVIER-STOKES ANALYSIS OF A HIGH WING TRANSPORT HIGH-LIFT CONFIGURATION WITH EXTERNALLY BLOWN FLAPS ... developed

AIAA-2000-4219

NAVIER-STOKES ANALYSIS OF A HIGH WING TRANSPORTHIGH-LIFT CONFIGURATION WITH EXTERNALLY BLOWN FLAPS

Jeffrey P. Slotnick*, Michael Y. An, Stephen J. Mysko, and David T. YehThe Boeing Company-Long Beach

Stuart E. Rogers and Karlin Roth*

NASA Ames Research Center

M. David Baker and Steven M. NashMCAT, Inc.

* Senior Member, AIAA.

Copyright 2000 by American Institute of Aeronautics andAstronautics, Inc. All rights reserved.

Abstract

Insights and lessons learned from the aerodynamicanalysis of a High Wing Transport high-liftconfiguration are presented. Three-dimensional Navier-Stokes computational fluid dynamics simulations usingthe OVERFLOW flow solver are compared with highReynolds-number test data obtained in the NASAAmes 12-Foot Pressure Wind Tunnel facility.Computational analysis of the baseline High WingTransport high-lift configuration with and withoutexternally blown flap jet effects is presented. Using thedeveloped computational fluid dynamics analysiscapability, several aerodynamic investigations,including an assessment of nacelle strake effectivenessand wake vortex prediction, are presented.

Introduction

To reduce the costs associated with the design,development, and manufacturing of modern transportaircraft, major improvements in both aircraftdevelopment cycle (time-to-market) and affordability(flyaway cost) must be achieved in order to make newconcepts commercially viable. An important driver inreducing aircraft development cycle time and totalsystem cost is the reduction of aerodynamic designcycle time.

To enable these improvements, significant effortsto utilize computational fluid dynamics (CFD) forhigh-lift system design are underway. Effective use ofCFD could greatly reduce the amount of wind tunneltesting required in the design process, and therebyreduce development cycle time and cost. EmergingNavier-Stokes methods have the ability to model theviscous-dominated flow physics associated withadvanced transport aircraft aerodynamics. However, to

be successful, CFD methods must be easy to use, fast,and accurate.1

High-lift analysis is particularly challenging, andis a pacing item in reducing aerodynamic cycle time.Computational analysis of realistic high-lift systems iscurrently limited by configuration geometriccomplexity, and the wide range of flow phenomenatypically present in high-lift flow-fields. High-liftaerodynamics is extremely complex, and is dominatedby viscous flow phenomena, such as separation, shock-boundary layer interactions, and multiple elementmerging confluent wakes.2 A significant body of workhas been performed in two dimensions (2D) tocharacterize high-lift flow physics.3-6 An excellentsummary of 2D high-lift CFD is given by Ying.7

Although 2D Navier-Stokes analysis is beginning tomake inroads into the high-lift system design process,high-lift flows are inherently three-dimensional (3D).Testing and simulation of complete transportconfigurations are required to accurately predict highlift system performance.

In recent years, much progress has been shown inthe ability to simulate 3D high-lift flow-fields usingNavier-Stokes methods. Building block CFDcomputations on wings with full-span and part-spanflaps,8-10 and on more complicated wing/bodygeometries11 have illustrated the range of applicabilityof present tools, and highlighted the shortcomings stillpresent in Navier-Stokes methods. However,computing the three-dimensional subsonic flow over acomplete transport high-lift configuration remains asone of the most difficult challenges of modernCFD.12,13 Moreover, the simulation of powered, high-lift systems, such as the inclusion of engine jetimpingement effects on an externally blown flap(EBF), has not been explored.

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AMERICAN INSTITUTE OF AERONAUTICS AND ASTRONAUTICS PAPER 2000-4219

Under the Integrated Wing Design (IWD) elementof the NASA Advanced Subsonic Technology (AST)program, significant effort was concentrated ondeveloping advanced design methodologies to enableboth the reduction of design cycle time, and thedevelopment of advanced aerodynamic concepts.Within the AST-IWD high-lift subelement, highReynolds number test data was obtained and used togain insight into high-lift flow physics, and to calibratehigh-fidelity Navier-Stokes numerical simulations. Aspart of this activity, advanced computational tools weredeveloped to enable production level CFD analysis ofthree-dimensional high-lift systems on completetransport configurations.14 As a result of these efforts,the analysis time associated with generating a CFDsimulation from geometry to final solution on acomplete high-lift transport configuration was reducedby an order of magnitude. This technology wasextensively used in the current work to computationallyanalyze the aerodynamic performance of a High WingTransport (HWT) high-lift configuration with EBF jeteffects.

The AST-IWD program designed, built, and testedan experimental semi-span model of an HWT aircraftin high-lift mode. The main objectives of the HWTwind-tunnel test program included the development ofsemi-span, powered testing methods, and theevaluation of advanced flap concepts. Perhaps the mostimportant goal, and the focus of the current work, wasto provide quality data for CFD code calibration .

This paper gives an overview of the numericalsimulation of the complete 3D HWT landingconfiguration using Reynolds-Averaged Navier-Stokes(RANS) methods. Rather than concentrating on thenumerical details of the CFD simulations themselves,an assessment of the predictive capability of thecomputational method relative to test data isemphasized. Several technical issues, including testfacility wall interference effects and jet modelingeffects, are presented and discussed in detail. Theapplication of the computational method to studynacelle strake effectiveness and wake vortex predictionis also highlighted.

HWT Configuration

Both the experimental and computational modelsconsist of a fuselage, a wing with high-lift system, awinglet, and two under-wing-mounted nacelles at 30%and 56% span, respectively. The high-lift system iscomposed of three slat segments on the wing leading-edge, and a double-slotted (vane/flap) flap aft of themain wing. The flap is attached to the main wing withfour hinged brackets, which are covered by fairings.Inboard and outboard nacelle strakes are located on

each of the nacelles. Neither the experimental nor thecomputational model included an empennage.

Experimental Data

Under the AST-IWD program, extensive highReynolds number powered test data was obtained forthe semi-span HWT model in the NASA AmesResearch Center 12-Foot Pressure Wind Tunnel (PWT)facility. Turbine Powered Simulator (TPS) units wereused to simulate the engine exhaust flow. Datacollected included standard force and moments, and avast array of surface pressure transducer measurementson the fuselage, the main wing, the high-lift elements(slat, vane, and flap), and winglet. Also, flow-fieldpressure and temperature measurements were obtainedusing the Boeing Quantitative Wake Survey System(QWSS).15 An additional test entry, designed for TPScalibration, was used to collect jet flow-field data usingQWSS for a single TPS unit mounted on a stingsupport.

All experimental data and computational resultsare subject to the Limited Exclusive Rights provisionsof the NASA AST contract, and are consideredsensitive. Because of this, all axis labels have beenremoved from the data plots.

Computational Analysis

Grid Generation. The Chimera overset gridmethod was employed to discretize the HWT surfaceand volume domain.16 The Chimera Grid Toolspackage17, coupled with an advanced grid scriptingsystem,11,14 was used to generate the majority of thegrid system. A complete summary of the gridgeneration tools and processes used to build the HWThigh-lift configuration grid system is reported byRogers et al.18 For the work presented here, thecomputational model of the HWT high-lift landingconfiguration installed in a computational model of theNASA Ames 12-Foot PWT facility is shown in Fig. 1.

Figure 1. HWT computational model geometry.

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The computational geometry of the wind tunnelincludes an inviscid, simplified model of the testsection, and neglects the effects of viscosity of the testsection splitter plate and semi-span mountingapparatus, as well as the divergence of the wind tunnelwalls downstream of the test section. The simplifiedinviscid-tunnel model has been shown to be as accurateat simulating wind-tunnel wall interference as thecomplete viscous-tunnel approach. 19 The HWT oversetgrid system is composed of 153 grids and 35.2 milliongrid points, and contains all of the geometrycomponents on the experimental model, except the slatand vane brackets. Fig. 2 shows the HWT high-liftsurface grid system, with only every other point in eachdirection shown. Grid topologies and densities for thehigh-lift system components are based on results fromearlier 2D studies.3 Box grids were used to discretizethe volume to the far-field boundaries. Inter-gridcommunication was established using the PEGSUScode.20

Figure 2. HWT surface grid system.

Flow Solver. The OVERFLOW finite-differenceNavier-Stokes flow solver,21 with multigridconvergence acceleration,22 was used to solve theviscous HWT high-lift flow-field. The ARC3Ddiagonal scheme was used on the left-hand side of theflow equations. Third-order Roe upwind differencingwas used for the right-hand side. Thin-layer viscousterms in all three directions were used in all grids. LowMach number preconditioning was used in all boxgrids. Component forces and moments are computedwithin OVERFLOW using “zipper” grid informationsupplied by the FOMOCO23 utility MIXSUR.

Turbulence and Transition Modeling. Based onresults from previous 2D high-lift studies,3-6 theSpalart-Allmaras (S-A) one-equation turbulencemodel24 is currently accepted as the most appropriateturbulence model for multi-element high-lift flow-field

analysis. In conjunction with upwind differencing usedin the HWT flow-field simulations, the S-A modeldemonstrates good overall convergence and robustness.However, as will be discussed later, the use of the S-Amodel is not successful in predicting the spreading andmixing of the EBF jet flow.

A significant body of research in boundary-layertransition and relaminarization for high-lift flow-fieldsexists in the literature.6,7,25 Although it is concludedthat modeling transition is critical in predicting thevelocity deficits in the confluent wakes accurately for2D multi-element airfoils, the ability to modeltransition for complete 3D configurations currentlydoes not exist. For this reason, coupled with a lack ofadequate HWT transition location data, allcomputations are performed assuming fully turbulentflow everywhere.

Solution Details. All HWT high-lift configurationOVERFLOW simulations were performed on either a16 processor Cray C90 or a 128 node Silicon Graphics(SGI) ORIGIN 2000 (O2K) computing platform. Anew, multi-level parallel version of OVERFLOW,known as OVERMLP,26 was used on the O2K. A totalof 47 different HWT cases have been computed fordifferent geometry configurations and flow conditionswithin the AST program. Each of these cases requiredbetween 200 to 400 Cray C90 CPU hours, or 1000 to2000 O2K CPU hours, depending on the angle-of-attack and the power setting. For most cases,convergence to a steady-state was achieved withapproximately 2000 multi-grid cycles. These caseswere considered converged when the change in totalconfiguration lift coefficient over the last 100 cycleswas less than 0.001. A detailed summary of typicalOVERFLOW convergence characteristics for large-scale transport high-lift configurations is reported byCao et al.12

Results and Comparisons

Baseline Configuration. Both experimental dataand computational simulations for several HWTconfigurations at various flow conditions and flapdeflections were obtained during the AST/IWDprogram. In this paper, the baseline configurationmodeled with CFD is defined as the nominal landingHWT high-lift geometry with slats and a single double-slotted (vane/flap) flap deflected. To evaluate thecapability of the OVERFLOW code to accuratelypredict the aerodynamic performance of the baselineHWT high-lift configuration, several simulations withand without EBF power effects at a range of angle-of-attack were computed. The freestream Mach number(M∞) was 0.175, and the Reynolds number based onmean aerodynamic chord (MAC) was 13 million.

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In Fig. 3, total configuration lift coefficient versusangle of attack is plotted. In Fig 3a, test data notcorrected for wall interference (uncorrected) iscompared with CFD simulations with wind tunnelwalls modeled in both unpowered (cµ=0.03) andpowered (cµ=0.5) modes, where cµ is the engine thrustcoefficient. Fig 3b shows corrected test data comparedwith free-air CFD results. For illustration purposes, thelinear portions of each experimental lift curve in Fig.3a have been extrapolated and are plotted with black,dashed lines. Numbers are used to denote CFDdatapoints at specific angles-of-attack. Wallinterference corrections to the experimental data havebeen determined using the Two Variable Method,27

which is based on surface pressure integration usinginterpolated and extrapolated wall-pressuremeasurements.

Test Data (Cmu=0.03)Test Data (Cmu=0.5)OVERFLOW (Cmu=0.03)OVERFLOW (Cmu=0.5)

1

2

Test Data (Cmu=0.03)OVERFLOW (Cmu=0.03)

1

2

3 4

5

CL

CL

Angle of Attack

(a) Uncorrected test data versus CFD with wind tunnel walls modeled.

(b) Corrected test data versus CFD without wind tunnel walls modeled (free−air).

Figure 3. Comparisons of total configuration lift.

In general, the unpowered CFD results, with andwithout wind tunnel walls modeled, compare very wellwith respective test data up to maximum lift. For thefree-air OVERFLOW simulations, an additional

solution beyond maximum lift was obtained. Wingsurface pressure comparisons for the unpowered casesare also excellent, demonstrating that both theintegrated forces and local flow details are captured.Figure 4 shows slat, main wing, vane, and flap pressurecomparisons at 42% span between free-airOVERFLOW results and corrected test data at the fiveangles-of-attack indicated in Fig 3b.

Point 1

Point 2

Point 3

Point 4

Point 5

X Location

−Cp

−Cp

−Cp

X Location

95%

87%

78%69%

60%

53%

42%

34%

26%

17%

56%

30%

OutboardNacelle

InboardNacelle

12

34

Figure 4. Surface pressure comparisons at 42%span, cµ=0.03 ( o Corrected test data,² free-air OVERFLOW solutions)

Also shown is the wing pressure transducer layout,with the high-lift elements in exploded view form. Arow of pressure transducers was located at 42% spanspecifically to capture the flow in the region of the

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high-lift system where EBF effects are minimal,halfway between the two nacelles.

As can be seen, experimental pressures agree wellon all the high-lift elements. The slat, andcorresponding wing leading edge, peak pressureincrease with increasing angle-of-attack is wellcaptured by CFD. Vane and flap pressures comparewell between CFD and test data up to maximum lift.Beyond maximum lift, however, OVERFLOW resultsshow a sharper decrease in lift on the vane and flapelements relative to experimental data. We believe,based on these results, that OVERFLOW doesaccurately predict the unpowered HWT high-lift flow-field up to maximum lift. However, further studies arerequired to properly determine if current Navier-Stokesmethods can simulate the complex flow breakdownmechanism encountered at stall.

Unlike the unpowered comparisons, the poweredcomputational results differ markedly from thecorresponding experimental results. The test datashows a distinct non-linear increase in the lift-curveslope at high, pre-stall angles of attack. Although theexact mechanism is not well understood, it is believedthat the non-linearity in the lift-curve slope of thepowered test data is principally due to wind-tunnel wallinterference effects.

Wind Tunnel Wall Interference Effects. Toillustrate the relative proximity of the HWT high-liftsystem with respect to the tunnel test section, Machnumber contours on a spanwise cutting plane throughthe inboard nacelle, pylon, and wing components areshown in Fig. 5.

Tunnel Side Wall

Tunnel Side Wall

FlapVane

Figure 5. Mach number contours in inboardnacelle/pylon pitch plane.

The outboard nacelle/pylon and flap hinge fairingsare not drawn for clarity. The surfaces of the windtunnel side walls are depicted with heavy black lines.

As seen in this tunnel top view, the EBF-deflected jet isforced to bend downstream (upward) prematurely atthe wall. Because of this, a reactive force effectivelypushes the jet vertically upward. The jet impingeshigher on the flap, the flap normal force is increased,and overall wing circulation levels are elevated due tomore jet flow through the vane and flap gaps. As aresult, the total configuration lift increases.

Unfortunately, no independent test data exists toquantify the wall-interference effects. However,significant insights into wind-tunnel-wall interferencehave been obtained using CFD, even though thepowered OVERFLOW results show only a slight non-linearity in lift-curve slope at the higher angles ofattack. In Fig. 6, the slat, main wing, vane, and flapsurface pressure comparisons between OVERFLOWsolutions with and without wind tunnel walls modeled,at the second alpha point indicated in Fig. 3a, areplotted.

-Cp

-Cp

-Cp

X LocationX Location

30%

42%

56%

34%

60%

OVERFLOW, Wind Tunnel Walls

OVERFLOW, Free−air

Figure 6. Surface pressure comparisons, cµ=0.5.

As expected, the presence of the wind tunnel wallsprovides more suction on the upper-wing surface.However, at 30% and 56% span, the flap lower surfacepressures are significantly higher in the computationwith walls modeled, relative to the free-air simulation.Moreover, at 30% span (the location of maximum jet

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impingement), the peak pressure on the flap occursfurther forward towards the flap leading edge in thewind-tunnel wall solution. This is indicative of the jetbeing pushed vertically upward due to the presence ofthe walls.

Although the CFD solutions contain a discerniblewind-tunnel wall interference effect, there are severalother plausible reasons that may account for thedifferences between CFD and test data. For thecomplicated HWT model, the use of a simplified windtunnel computational model, for example, may notadequately represent the true flow in the tunnel. Theinclusion of the semi-span splitter plate and relatedhardware may be required to compute the correcttunnel dynamic pressure. Also, there may beinadequate grid density or resolution in critical regionsof the flow-field for the powered simulation. However,the most likely explanation for the discrepancybetween CFD and experimental data is inadequateturbulence modeling in the EBF jets.

Jet Flow Simulation Capability. Even thoughdeficiencies in the turbulence modeling of the EBF jetflow affect the computational prediction of HWTaerodynamic performance at all angles-of-attack, theshortcoming is clearly illustrated at a relatively lowangle-of-attack. Fig. 7 shows the slat, main wing, vane,and flap element surface pressure comparisons betweenCFD and test data at five spanwise rows of pressuretaps for both unpowered (a) and powered (b)simulations, at the first alpha point indicated in Fig. 3a.For the unpowered case, surface pressure comparisonsbetween CFD and experimental data for the baselineHWT configuration show excellent agreement. For thepowered case, OVERFLOW fails to predict the largepressure rise due to the jet exhaust impingement on theflap at 34% and 60% span. As a result, the total liftpredicted by OVERFLOW is lower than the test data.

To calibrate the TPS thrust levels in the NASAAmes 12 Foot PWT in support of the HWT modelexperiment, an isolated TPS test was conducted. Aspart of this test, QWSS flow-field measurements wereobtained at three streamwise locations in the jetexhaust to aid in OVERFLOW code calibrationactivities for HWT EBF jet-flow simulation.

Figure 8a shows the overset computational modelof the TPS calibration setup. The sting-mounted modelcontains 2.6 million points in 32 grids. All simulationswere obtained at a Mach number of 0.2, angle of attackof zero degrees., a Reynolds number of 6.2 million, anda fan rotor speed of 53,000 RPM. Fig 8b showsrepresentative Mach contours in the pitch planepredicted using OVERFLOW for the TPS calibrationconfiguration at these flow conditions. Also depicted

are the three streamwise plane stations where QWSSflow-field data was collected.

26%

34%

42%

53%

60%

26%

34%

42%

53%

60%

X Location X Location

−Cp

−Cp

−Cp

−Cp

−Cp

(a) unpowered, cµ=0.03.

(b) power−on, cµ=0.5.

Figure 7. Surface pressure comparisons.( o Uncorrected test data, ² OVERFLOW with tunnel

walls modeled)

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(a) TPS calibration model grid system

1 2 3

(b) Mach numbers contours, cµ=0.5.

Figure 8. TPS calibration model CFD analysis.

In Fig. 9, Mach contours at these three stations fortest data (left) and OVERFLOW using the S-Aturbulence model (right) are shown. The fan and coremold line traces are superimposed to better show thecomparison of the jet exhaust expansion between theCFD and test data. The experimental data clearlyshows an asymmetric distortion of the jet at all threeQWSS stations, whereas the computed results wereconstrained to be symmetric. The experimentalasymmetry is likely caused by flow swirl in the TPSunit, and is perhaps influenced by the presence of total-pressure rakes placed asymmetrically inside the fan andcore ducts. Also, the test data shows an increasingamount of flow mixing as the jet exhausts downstream.This level of mixing may influenced by the presence ofa fine-grain flow straightening mesh placed at the headof the fan duct.

By comparison, the OVERFLOW simulationsshow markedly lower levels of jet spreading andmixing. Because the CFD results are obtained on ahalf-model of the TPS calibration geometry withoutinternal duct pressure rakes modeled, no asymmetricdistortion is simulated. Also, for the OVERFLOWruns, the engine boundary conditions are set atupstream faces within the fan and core ducts,respectively, based only on average available flow dataat these stations. Because of this, the OVERFLOWupstream jet boundary conditions are defined asaverage, radially-constant flow without swirl.

Therefore, there is no mechanism within the numericalsimulation to introduce swirl into the solution.

Consider also that the S-A one-equation turbulencemodel is used in the simulations. Although the S-Amodel has been shown to be a good general-purposeturbulence model appropriate for high-lift flowanalysis, the model was not designed for the predictionof jet flows.28

Test Data OVERFLOW

(a) QWSS Plane 1.

Test Data OVERFLOW

(b) QWSS Plane 2.

Test Data OVERFLOW

(c) QWSS Plane 3.

Figure 9. QWSS Mach number comparisons.

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Based on these results, OVERFLOW does notaccurately predict the proper mixing and spreading ofthe jet-engine exhaust. Specifically, for the completeHWT model, the jet impingement on the flap isconfined to a smaller area than in the experiment. Thisresults in lower total-configuration lift relative to thetest data. It is clear that without the TPS geometricasymmetries and complex jet inflow flow conditionsmodeled appropriately in OVERFLOW, exactmatching of jet flow features is impossible. However,even with these deficiencies, the matching of gross jetflow magnitudes has been achieved. Furtherinvestigation of TPS jet flow-fields with more complexturbulence models is warranted.

Additional Studies

As demonstrated, complete HWT high-liftcomputational simulations do not predict exact vehicleperformance at either high angle-of-attack conditionsand/or high power settings. However, valuable insightsinto important aerodynamic issues can be gleaned fromHWT Navier-Stokes analysis. Two significant issues,nacelle strake effectiveness and wake vortexprediction, were investigated under the AST-IWDhigh-lift program. A summary of the computationalanalysis and lessons learned are presented below.

Nacelle Strake Effectiveness. Nacelle strakes (orchines) are commonly used to alleviate the adverseaerodynamic affects of wing-mounted high-bypass-ratio engine installations. The close coupling of thelarge engines with the wing results in increased flow-field interaction at the wing/pylon juncture at highangles-of-attack. Without nacelle strakes, this flow-field interaction results in reduced aircraftperformance, especially at maximum lift.29 Increasedunderstanding of nacelle strake effectiveness is criticalto properly optimize strake size and location. If thisassessment is to be done numerically, determining thecomputational requirements to predict nacelle-strakeeffects is equally as important. For these reasons, theability of Navier-Stokes CFD methods to capturenacelle-strake effects was evaluated.

OVERFLOW solutions without power (cµ=0.03) ata moderately high angle-of-attack were obtained forHWT high-lift configurations with all (inboard nacelleand outboard nacelle) strakes on and with all strakesoff. As with the previous baseline HWT simulations,the freestream Mach number was 0.175, and theReynolds number was 13 million based on MAC.Other than the fine grid boxes surrounding the nacelleand main wing structures, no special grid resolutionwas built into the baseline HWT grid system tospecifically capture the nacelle-strake vortices.

Nacelle strakes are designed to increase maximumlift through two flow mechanisms. First, nacelle strakesreduce upwash over the unprotected portion of themain wing near the wing/pylon intersection. As aresult, the flow is straightened as it reaches the mainwing, thereby relieving some of the adverse effects ofthe engine installation. Figure 10 shows the surfacestreamline patterns on the outboard nacelle with strakesoff (a) and with strakes on (b). These images clearlyshow the reduction of flow upwash into the wing/pylonjuncture with strakes on. As a result, cross-flowseparation on the nacelle near the pylon juncture isreduced (Fig. 11).

(a) Strakes−off

(b) Strakes−on

Figure 10. Nacelle surface oilflow.

Strakes On

Strakes Off

Cleaner flow at the interruptedslat region of the wing−pylon

intersection

Strakes eliminatespanwise separated

flow across pylon

Figure 11. Nacelle strake particle traces.

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Second, the presence of the strake vorticesprovides downwash on the wing upper surface, andserves to energize the boundary layer. Fig. 12 depictstotal pressure contours above the wing upper surface at20% chord aft of the inboard wing/pylon juncture.

Slat EdgeVortices

(a) Nacelle strakes off.

Nacelle StrakeVortices

(b) Nacelle strakes on.

Figure 12. Total pressure contours on wing at 20%chord near the inboard nacelle-pylon installation.

For both strake-off and strake-on configurations,the individual slat-edge vortices and slat wakes arewell pronounced. For the strake-on case, the nacelle-strake vortices, positioned above the slat-edge vortices,are discernible as well. In the strake-off simulation, asthe inboard slat-edge vortex wraps around the uppersurface of the wing, it pulls flow away from thesurface. In contrast, in the strake-on computation, thestrake vortices provide counter-rotating downwash thateffectively alleviates this flow behavior. As a result,the health of the boundary layer on the wing uppersurface is improved, and flow separation is delayed.Also, note that the presence of the nacelle strakes tendsto diminish the strength of the middle-slat-edge vortex.

It is clear that OVERFLOW HWT nacelle strakeanalysis has yielded valuable insight into the effect ofstrake vortices on high-lift flow physics. However, useof CFD to predict the actual magnitude of the strake liftincrement is not yet appropriate. The strake-liftincrement is defined as the difference between total liftwith strakes on minus total lift with strakes off. At thismoderately high angle-of-attack, experimental datashows a strake-lift increment that is nearly twice themagnitude of that predicted by CFD. The difference inthe strake-lift increment can be attributed to twosources. First, the relatively coarse grid used to capturethe strake vortices is inadequate to properly simulatethe correct amount of vortex dissipation. Second, atthis angle-of-attack, the flow is on the verge ofseparation. The onset of separation near maximum liftis not well predicted by current Navier-Stokesmethods. Since nacelle strakes were designed toimprove maximum lift, many additional simulationsare needed to evaluate the accuracy of Navier-Stokesfor flows with incipient separation.

Wake Vortex Prediction. Reducing the adverseeffects of transport aircraft wake-vortex formation is animportant element in increasing airport operationcapacity. By reducing wake-vortex turbulence, aircraftseparation distances on approach can be decreased.With more aircraft able to serve busy airports, costlydelays could be minimized.30,31

To alleviate trailing-wake vortices, anunderstanding of the flow physics associated with wakevortices is important. Moreover, for the baseline EBF-powered HWT configuration, evaluation of the jet-blowing effects on wake vortex formation is critical.Much work on the theoretical understanding of wakevortices exists in the literature,30 but quantitativeevaluation of wake-vortex formation for completetransport configurations using Navier-Stokes methodsis not widely available.

During the HWT model test, QWSS flow-fieldsurvey data was obtained specifically to study the wakevortex field behind the HWT high-lift configuration.The constant streamwise QWSS measurement plane islocated approximately one tip-chord downstream of thewing trailing edge.

To evaluate the capability of the OVERFLOWcode to predict the near-field wake vortex structure,several CFD simulations at representative approachpower conditions for a variety of flap configurationswere obtained, and compared with QWSS data. Figure13 shows the Mach contour comparison between theQWSS test data (a), OVERFLOW modeled withoutwalls (b), and OVERFLOW modeled with wind tunnel

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walls (c), for the baseline landing configuration. Theview is looking forward from tail to nose, and validQWSS data is contained within the concentration ofcontour lines around the probe sweep perimeter.

(a) Test data.

(b) OVERFLOW, free-air.

(c) OVERFLOW, wind tunnel walls.

Figure 13. Comparison of Mach number contours inQWSS wake vortex survey plane.

In general, the CFD simulation results modeledwith walls and the experimental data agree quite well.The flap edge vortex and winglet-wake are wellpredicted by OVERFLOW. The distortion of both theinboard and outboard engine jets downstream of theflap is also well simulated by OVERFLOW. Theflattened shape of the inboard nacelle jet contour in theQWSS measurement plane gives further evidence ofthe impact of wind-tunnel wall-interference effects onthe wake-vortex flow-field. The free-air CFD resultshows some jet exhaust flattening, but not to the sameextent as the wind-tunnel wall case. Also, theexperimental Mach contours appear to be shifted upvertically relative to both CFD solutions. This indicatesthat both inboard and outboard jet exhausts are closer

to the wing, and further supports the discrepancy in thepowered lift-curve slopes between CFD and test datashown earlier.

Fig. 14 shows the vorticity and velocity vectorcomparison between QWSS-obtained test data andOVERFLOW CFD results. The x-component ofvorticity is shown. Red contours indicate clockwiseflow, and blue contours indicate counterclockwiseflow. The vortex field aft of the wing trailing edge ismodeled well using OVERFLOW. The experimentalresults and CFD simulations both clearly show thestrong flap-edge vortex, the winglet junction vortex,and the tip vortex. The vortices in the jet exhaust arewashed out due to inadequate grid resolution.

(a) Test data.

(b) OVERFLOW, free-air.

(c) OVERFLOW, wind tunnel walls.

Figure 14. Comparison of vorticity contours andvelocity vectors in QWSS wake vortex survey plane.

Close inspection of the differences between CFDand test data show further evidence of the impact ofwind-tunnel-wall interference on near-field wakevortex structure. Behind the trailing edge of the flap,

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the vortex behind and below the outboard nacelle iswell pronounced in the QWSS flow-field survey data.This vortex is shed principally from the flowexpanding around flap-hinge fairing #3 (where #4 isthe furthest outboard fairing). It is believed that sincethe jet exhaust is pushed up closer to the wing by thetunnel walls, the spanwise flow is stronger, and thuscreates a stronger vortex. The OVERFLOW simulationwith wind-tunnel walls predicts a weak vortex, whilethe free-air CFD solution does not capture the vortex atall.

Based on this initial evaluation, it is believed thatOVERFLOW can adequately predict the grossmagnitude and location of trailing wake vortices. Gridrefinement aft of the main wing will most likelyimprove the accuracy of these results.

Conclusions

The OVERFLOW code has been applied to theaerodynamic analysis of the complete HWT high-liftconfiguration. The comparison of unpowered CFDsimulations with experimental data up to maximum liftis excellent. Power-on CFD simulations, however,underpredict the total configuration lift relative to thetest data, which exhibits a distinctly non-linearcharacter to the lift curve slope at higher angles-of-attack. Evidence shows that wind-tunnel-wallinterference effects is the likely reason for the non-linear lift behavior seen in the test data, whileinadequate turbulence modeling for jet flows is one ofthe likely causes for differences seen between CFD andpowered test data.

Although vast improvements have been made tothe tools that enable large-scale CFD analysis of high-lift systems, there are still significant shortcomings inthe accuracy of Reynolds-Averaged-Navier-Stokes(RANS) methods. Use of current general-purposeturbulence models is not fully applicable for the widerange of flow features (confluent shear layers, vortex-viscous interactions, EBF jet effects, etc.) typicallypresent in high-lift flow-fields. What is needed is animproved turbulence model that is able to adequatelysimulate the different types of high-lift flow physicspresent within a given CFD computation. Also, currentexperience with HWT CFD analysis at the maximumlift condition suggests that RANS methods adequatelypredict the location of flow separation, but fail topredict both the maximum lift magnitude, and theangle-of-attack at stall. Improved technology toaccurately predict the onset of separation at criticaldesign conditions is needed.

However, by increasing our understanding of EBFHWT high-lift flow physics, the CFD simulations have

been successfully used to gain valuable insights intoseveral aircraft aerodynamic issues. Navier-Stokesanalysis has greatly increased our understanding of theflow physics associated with nacelle strakes and theireffect on increasing maximum lift. Also, preliminaryassessment of OVERFLOW to predict near-field wakevortex characteristics has been successful. As a result,use of OVERFLOW to numerically investigate wakevortex alleviation concepts will be further explored.

Finally, the use of complete-configuration CFDanalysis in the advanced stages of a high-lift systemdesign process shows great promise. But, due to thelarge computational costs associated with the analysisof complex geometries, and current limitations insolution accuracy at critical flow conditions,application of RANS methods for high-lift design isnot yet practical.

Acknowledgments

The authors would like to acknowledge thetechnical contributions and assistance of M. E.Whitlock and R. C. Potter, The Boeing Company,Pieter Buning, NASA Langley Research Center, and D.C. Jespersen and T. H. Pulliam, NASA Ames ResearchCenter. Also, many thanks to the AST-IWDmanagement team for their valuable leadership andinsight: F. T. Lynch, AST/IWD Program and TechnicalManager, and D. H. Leopold, High-Lift Task Lead,The Boeing Company, and E. B. Plentovich, AST-IWDProgram Manager, and K. M. Jones, High-LiftSubelement Lead, NASA Langley Research Center.This work was funded under the Advanced SubsonicTechnology program through NASA contract NAS2-20268.

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