3101

ICAS 2002 CONGRESS

3101.1Copyright 2002 by Fassi Kafyeke Published by the International Council of Aeronautical Sciences, with permission

DEVELOPMENT OF HIGH-LIFT SYSTEMSFOR THE BOMBARDIER CRJ-700

Fassi Kafyeke, François Pépin and Cedric KhoAdvanced Aerodynamics, Bombardier Aerospace

Montréal, Canada

Keywords: flaps, slats, high-lift, wind tunnel testing

Abstract

This paper describes the development process ofthe low-speed configuration of the CRJ-700,with an emphasis on half-model high Reynoldsnumber wind tunnel testing. The processincluded CFD design and analysis, wind tunneltests and flight tests. The experimentaldevelopment of the flaps and slats wasperformed using a 7% scale half-model, in theIAR 5ft x5ft High Reynolds Number wind tunnel.Results from these tests are in very goodagreement with flight test data and decisionsmade in the wind tunnel lead to a short andsuccessful flight test program.

1 Introduction

Over the last decade, Bombardier Aerospacedesigned and put into production a number ofregional and business aircraft. The CRJ-700 is a70-seat version of the highly successfulCanadair Regional Jet CRJ-200. The CRJ-700wing was essentially a blank sheet design with acompletely new planform and with the additionof leading edge devices. During thedevelopment of the CRJ-700 aerodynamicconfiguration, the low-speed characteristics ofthe airplane were carefully studied. Threeconsiderations guided the design: safety,simplicity and performance. In practice, safetyrequired clearly identifiable stall characteristicswith a nose down pitching moment and little orno roll excursion. For simplicity, double-slottedhinged flaps similar to those of the CRJ-200were used. Performance meant as much usablelift as possible and good second-segment-climbcapability.

2 Aircraft Design Characteristics

The CRJ-700 was designed to carry seventypassengers and three crew members over arange of 1685 Nautical Miles (1985 NM for theextended range version) at Mach number 0.78.Maximum operating Mach number is 0.83 andmaximum cruising altitude 41,000 feet. Theaircraft, designed with a Maximum Take-OffWeight (MTOW) of 73,000 lbs (75,250 lbs forthe extended range version), is powered by twofuselage mounted General Electric CF-34-8C1turbofans, each developing 12,670 lbs thrust.The take-off distance at MTOW is 5,130 ft andthe landing distance, at the Maximum LandingWeight of 67,000 lbs is 4850 ft. Figure 1 givesthe general arrangement of the aircraft.

Figure 1: CRJ-700 aircraft general arrangement

26.60 ft(8.11 m)

13.51 ft(4.12 m)

76.25 ft(23.16 m)

24.00 ft(7.32 m)

106.34 ft (32.41 m)

Fassi Kafyeke, François Pépin and Cedric Kho

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3 CRJ-700 High Lift Wing Design

3.1 Planform DesignThe CRJ-700 initial wing planform W30 wasobtained from the CRJ-200 wing by adding a72” root plug and extending the leading edge by9% chord (Figure 2). The intention was tomaintain the existing wing box structure on theouter wing panel and limit airfoil modificationsto the part of the wing forward of the front spar.This planform was abandoned after it wasrealized that the resulting wing sections couldnot provide adequate high-lift characteristics.The leading edge crank was also an impedimentto the installation of a simple leading edgedevice. A new planform, W33, was designed,with the 72” root plug but this time with astraight leading edge. The constraint ofpreserving the CRJ-200 front spar was alsoremoved and new advanced supercritical wingsections were designed. The wing retained thesimple double-slotted hinged flaps in twosegments of the CRJ-200 and was fitted withnew leading edge slats. After the first series ofhigh speed and low speed tests, the W33 wingwas replaced by wing W34, with the sameplanform and improved wing sections. Flapsand slats were redesigned to accommodate thenew wing geometry.

Figure 2: CRJ-700 wing planform development

The reference area of wing W34 is 738 ft2 andits aspect ratio 7.38. The quarter chord sweep ofthe outboard wing panel is 26.9 degrees. Thewing is fitted with winglets canted 20 degrees.

3.2 High Lift System DesignFigure 3 shows the planform definition of theCRJ-700 high-lift system. The trailing edgeflaps are double-slotted hinged flaps. Theinboard flap has a constant 33-inch chord and isfitted with a spring-loaded constant chord flapvane. The outboard flap is physically the sameas the outboard flap of the CRJ-200. The flapchord is a constant percentage of the main wingchord (24% chord). It is fitted with a fixedtapered flap vane. The inboard flap deflectionangles are 10 and 20 degrees for take-off, 30degrees for approach and 45 degrees forlanding. The outboard flap is geared with a ratio8/9 to the inboard flap deflections, deployingtherefore to 40 degrees in landing. The gearingimproves the spanwise load and preservesexcellent aileron performance in allconfigurations. The leading edge slats cover thefull span except for an inboard segment of thewing, which was left unprotected to improve thestall characteristics. The slats are tapered,covering 15% of the wing chord at the breakand 17% chord at the wing tip. The slatdeflections are 20 degrees for 0 and 10 degreesflap settings and 25 degrees for all other flapsettings.

Figure 3: CRJ-700 flaps and slats planform definition

Two pairs of multifunction spoilers supplementthe ailerons and two pairs of ground spoilerscomplete the set of wing movable surfaces. Thetrailing edge flaps of the CRJ-700 retain thesimplicity of the CRJ-200 design. Figure 4shows various settings of the inboard flap. The

Inboard Flap

Slat #3

Multi Function SpoilersOutboard FlapAileron

Slat #2

Slat #1

Ground Spoilers

DEVELOPMENT OF HIGH-LIFT SYSTEMS FOR THE BOMBARDIER CRJ-700

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flap rotates around a fixed hinge axis to 45degrees. A small vane is nested against the flapleading edge in the flap-retracted position andsprings into position as the flap is deployed. Thevane on the outboard flap (Figure 5) is fixed tothe flap. This outboard flap rotates to 40degrees, in a geared ratio to the inboard flapsetting. A bent up trailing edge (BUTE) doorimproves the flow in the cove region on theoutboard flap. At flap deflections below 20degrees, the BUTE door effectively seals thepassage to the upper surface of the vane,allowing the flap to operate as a single-slottedflap. This improves the lift to drag ratio of thetake-off configuration.

Figure 6 illustrates the slat settings. Forsimplicity, the slats are deployed along acircular arc track with three positions: retracted,

20 degrees for normal take-off and 25 degreesfor short take-off, approach and landing.

4 Theoretical Methods

At the early stages, the development of the high-lift systems was conducted using ComputationalFluid Dynamics (CFD) methods: three codeswere used for the design of the multi-elementairfoils. CEBECI, a viscous panel method withstrong boundary layer coupling developed by T.Cebeci and his team at the University ofCalifornia in Long Beach [1], was used for rapidevaluation of configurations. The MSES viscousEuler code of Drela [2] and the NSU2D 2Dunstructured grid Navier-Stokes code developedby D. Mavriplis [3] were used for the designand final verification of flaps and slats.

Figure 7: NSU2D unstructured Navier-Stokes grid for aCRJ-700 multi-element airfoil.

The accuracy of these codes for high-liftcomputations was verified in a two-dimensionalhigh Reynolds number high-lift test that wascarried out on a Bombardier generic multi-element airfoil in the IAR 2D test facility inOttawa [4]. In Figure 8, pressure distributionsobtained with the NSU2D Navier-Stokes codeand the CEBECI2D viscous panel method arecompared. Some of the differences may be dueto different transition settings between the twocodes. The Navier-Stokes solution is a fullyturbulent one, whereas natural transitionlocations are predicted, using semi-empiricalcriteria, in the viscous panel code. The 3D high-lift characteristics were predicted using theAnalytical Methods Inc.’s VSAERO 3D panelmethod and Bombardier’s semi-empirical stallprediction methods developed and validated onthe Challenger CL-604, the CRJ-200 and the

Figure 4: inboard flap motion

Figure 5: outboard flap motion

Figure 6: slat motion


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Global Express. This method is based on thecriterion proposed by Valarezo [5]. Figure 9shows a typical VSAERO solution. Figure 10shows the predicted CLmax for the cleanconfiguration in wind tunnel conditions,compared to the experimental data.

Figure 8: Pressure distributions computed with theCEBECI and NSU2D codes on a CRJ-700 section; slat 28degrees, flap 40 degrees; M = 0.20, α=10 degrees, Re=12Million.

Figure 9: VSAERO solution for a CRJ-700 early landingconfiguration, flap setting: 45 degrees, slat setting: 28degrees Mach 0.2, Alpha=10 degrees

5 Wind Tunnel Testing

The CFD design was further optimized in thewind tunnel. The low speed wind tunnel testswere conducted in three phases:− Initial development wind tunnel tests to vali-

date CFD designs and to obtain initial

performance figures. These tests used theearly wing configuration W33.

− Detailed design and development wind tun-nel tests leading to the aerodynamic freezeof the external lines. These tests used thefinal wing configuration W34.

− Production wind tunnel tests to verify in de-tail many aircraft possible configurationsand to obtain the data needed before theaircraft first flight. These tests used theproduction configuration of the aircraft.

Two low-speed tests were conducted in eachphase: a high-Reynolds number half-model testand a low Reynolds number full model test.

Figure 10: Comparison of wind tunnel data withtheoretical prediction of CLmax for the CRJ-700 cleanconfiguration at wind tunnel conditions: Mach 0.20, chordReynolds number 6.5 million.

The first high-lift development test wasconducted in January 1996 using a 7% scalehalf-model in the 5ft x 5ft wind tunnel of theCanadian Institute for Aerospace Research(IAR) in Ottawa. The objectives in this testwere:- to evaluate various flap and slat designs and

establish optimal angles, gaps and overlaps;- to determine the optimal location and

geometry of the inboard slat inboard end;- to measure CLmax and determine longitudinal

stall characteristics;

x/c

Cp

-0.1 0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 1.1

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- to investigate Mach number and Reynoldsnumber effects;

- to measure the effectiveness of controlsurfaces in high-lift configurations at highReynolds number.

Comparison with data obtained at lowerReynolds number in full model tests at the IAR6ftx9ft atmospheric tunnel and at theMicroCraft 7ftx7ft high-speed tunnel in ElSegundo, California, were used to estimate thehalf-model corrections.

The second high-lift development test wascarried out in January 1997, using aconfiguration with the final wing W34. Theobjectives of the test were the same as those ofthe first phase, but with a view to freezing thehigh-lift systems. The test was also used todetermine the final take-off and landing flap andslat settings, including alternate configurationsand to establish the effect of various ice shapeson the wing and tailplane. Slat loads weremeasured using specially designed strain-gauged components.

5.1 Wind Tunnel DescriptionThe IAR 5ft x 5ft tunnel is a blow-downfacility, shown schematically in Figure 11,which can be used to test in the subsonic,transonic and supersonic flow regimes.

Figure 11: NRC/IAR High Reynolds number blow-downtunnel

The reflection plane model was installed in the5-ft transonic test section with perforated wallsas shown in Figure 12. The wall porosity wasset at 4.0% open during the entire investigation.

The test section was configured for half-modeltesting with a solid 135.5-inch long and 0.25-inch thick reflection plate, with an ellipticleading edge, installed on stand-off from thenorth perforated wall.

Figure 12: Wind tunnel half-model test section

A non-metric boundary layer plate was installedbetween the model and the reflection plate. Theoutboard of the plate was attached to a turntableand slaved to the balance rotation. The loadswere measured with the IAR 5-componentsidewall balance. The incidence of the modelwas measured using an accelerometer mountedin the fuselage center section. A potentiometermounted on the half-model support systemprovided a secondary reading of the modelincidence, which was used to control modelpitching. A check calibration of both incidencemeasurements was made prior to theinvestigation and the incidence zero for bothmeasurements established as part of the process.The instrumentation to establish the tunnelconditions was as follows: three 200-psiabsolute pressure transducers were used tomeasure the settling chamber stagnationpressure, the tunnel static pressure (on thetunnel ceiling) and a “slugged volume”reference pressure. A 45-psi absolute pressuretransducer was used to measure the atmosphericpressure. A resistance thermometer in thesettling chamber measured the stagnationtemperature. Model pressure measurementswere performed using an Electronic ScanningPressure (ESP) system using modules mounted


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in the fuselage cavity. All pressures weremeasured relative to the “slug volume”reference pressure. Other modules connected tothe ESP system were used to obtain the tunnelwall pressure signatures from six wall statictubes installed in the test section.

5.2 Model DescriptionThe model was a 7% scale representation of thestarboard half of the CRJ-700 aircraft, as shownin Figure 13. The reflection plane model wasmounted to the tunnel sidewall balance throughthe use of an L-shaped bracket. A 1-inch thickboundary layer plate, identical in plan view tothe perimeter of the aircraft plane of symmetry,isolated the model from the wall. There was a0.1-inch spacing between the model plate andthe boundary layer plate. A triple labyrinth sealon the outboard face of the boundary layer plateprevented dynamic pressure build up betweenthe model plate and the boundary layer plate. Asingle electric contact was installed to indicateany labyrinth fouling under load.

Figure 13: 7 percent scale half-model of the CRJ-700installed in the IAR 5-ft high Reynolds number windtunnel

The gap between the reflection plate and theboundary layer plate was pneumatically sealedwith Teflon strips. In addition, a rubber skirtwas attached to the outer perimeter of theboundary layer plate as shown in Figure 14. Thewing was made out of high tempered stainlesssteel. It was bolted to the L-shaped bracket.Conventional lap joints machined at theperiphery of the wing bar accepted detachable

clean wing leading and trailing edges or high-lift devices.

Figure 14: Schematic of the model mounting arrangement

A detachable winglet, BUTE doors and anassortment of control surfaces (ailerons, groundspoilers and multifunction spoilers) completedthe list of wing components. The wing wasinstrumented with three rows of pressure tapsoriented streamwise: the first one at the inboardflap mid-span, the second one at the outboardflap mid-span and the third one at the aileronmid-span. The tubes were all inside the wingand ran into conventional chordwise drilling andspanwise troughs towards the pressure scannerslocated inside the forward fuselage. The cleanwing configuration was fitted with 111 pressuretaps and the high-lift configuration with 192pressure taps. The part-span steel leading edgeslat was built in three sections representative ofthe aircraft arrangement. Three brackets persegment, representative of the real aircraftinstallation in their orientation and location, ranon fixed circular tracks inside the fixed wingunder-slat surface (WUSS). Each slat segmentwas instrumented with two rows of tapsoriented hingewise (normal to the front spar) forslat load measurements. Due to their largenumber, the rigid pressure tubes ran inside adummy bracket replacing the middle slatbracket on each segment. An additionalslat/WUSS assembly with three stain-gaugedbalances located inside the wing was


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manufactured. These components were used tomeasure slat loads in selected runs.

The trailing edge flaps consisted of two double-slotted hinged flaps. As on the real airplane, theinboard flap was supported by two flap trackfairings and the outboard flap by three flap trackfairings. The two outer fairings of each flapincorporated a compass mechanism that wasused to set all the required flap angles. A vaneattached to each flap could be repositioned forgap and overlap investigation. The fuselage wasbolted to the L-shaped bracket. It integrated thevertical stabilizer in order to attach thehorizontal stabilizer. A pylon and a flow-through nacelle were attached to the fuselage.The nacelle was designed to represent theengine inlet mass flow at one idle operatingcondition. Electronic pressure scannersconnected to the wing pressure ports and themodel plate inboard taps were fitted inside theforward fuselage. The cantilevered horizontaltailplane could be indexed at all requiredincidences. The elevator, mounted on hinges,was deflected into position using brackets.

5.3 Test ProceduresBombardier used this high Reynolds numberfacility for high-lift systems development forthe first time during the Global-Expressbusiness jet project. Procedures leading to validand repeatable forces, moments and pressuremeasurements were established then. The mainchallenge was to obtain stabilized flow qualitiesduring a period sufficiently long to traverse thelarge range of angles of incidence required forhigh-lift experiments. Most of the CRJ-700 runswere performed at a nominal Mach number of0.20 and a Reynolds number of 8.40 million perfoot or 6.5 million based on the model referencechord. Additional runs were made at Mach 0.15,0.30 and 0.40 to investigate Mach numbereffects. At Mach 0.20, runs were also made atthree additional Reynolds numbers: 4.7, 7.3 and10 million per foot to investigate scale effects.The model was pitched through a range ofangles of incidence between –6 degrees and 25degrees for non-slatted configurations andbetween –6 degrees and 30 degrees for

configurations with deployed slats. A few runswere made during which the model was pitchedto 33.5 degrees. After examination of therepeatability of force and pressure data obtainedwith fixed pitch, pitch pause and continuoussweep runs, subsequent runs were made withcontinuous pitch motion at the slowest sweeprate compatible with available stable flowconditions. This time varied between 15 and 30seconds, depending on the Mach number andReynolds number selected [6]. Theelectronically collected data was filtered at 3Hertz then digitized at 100 samples per second.The aerodynamic data was corrected for windtunnel wall interference using the wall pressuresignature. The correction was based on aprocedure developed for half-models by Mokry[7]. The procedure uses the model geometry, thecross-sectional area distribution and the tunnelparameters. It yields corrections ∆M to thefreestream Mach number, ∆α to the model angleof incidence and ∆CDbuyancy to the drag.Corrections for the tunnel freestream flowangularity were applied to the data using pre-calibrated correction algorithms. All highReynolds number high-lift runs were performedtransition-free on the wing. Transition was fixedon all other surfaces using polyester-resincylinders. A selected number of flowvisualization runs were performed using thefluorescent mini-tuft technique. Repeatability onmaximum lift coefficient was ∆CLmax = 0.01 to0.02 and the scatter on αstall was negligible.Repeatability on the drag of high liftconfigurations was ∆CD = 20 to 30 drag counts.Measurements of pitching moments wererepeatable within ∆CM = 0.01. With thesevalues, the IAR 5-ft tunnel test was consideredadequate for the validation of flaps and slatsgeometry and kinematics and the prediction ofCLmax and global longitudinal stallcharacteristics. To predict tailplane angles totrim and the drag polars of take-off and landingconfigurations, it was necessary to resort to datafrom a 7% scale full model test performed at theIAR 6ftx9ft wind tunnel (Figure 15)


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Figure 15: 7% scale full model of the CRJ-700 in IAR 6ftx 9ft atmospheric wind tunnel

5.4 Selected Wind Tunnel Test ResultsIn Figure 16, results obtained in the 5-ft tunnelon a tail-off half-model are compared with thoseobtained on a sting-mounted full model. Bothresults were obtained for the same Machnumber (0.25), chord Reynolds number (2.8Million) and the same transition tripconfiguration (wing tripped at 0.5% chordbelow the leading edge).

Figure 16: Half-Model effects; comparison of lift, dragand pitching moment coefficients on CRJ-700 tail-offconfiguration at Mach 0.25 Rec=2.8 Million. Opensymbols: full model, closed symbols: half-model

This figure shows a good correlation of lift anddrag but a clear shift in pitching moment. Testsperformed earlier on the Global Express hadshown that this is due to a larger down force onthe fuselage mounted nacelle and pylon on thehalf-model. Pitching moments measured onwing-body configurations are much closer.This, however, did not change the stall angle orthe pattern of the stall on the wing. The high-lifttests showed that the redesigned airfoil sectionson wing W34 had improved the clean

configuration CLmax by 0.04. Configurationsinvestigated included flap angles 0, 10, 20,30and 45 degrees. Slats were deflected along asimple circular arc track. The slat anglesavailable were 15 degrees (sealed slats), 17, 20,22, 25, 28 and 30 degrees (gapped slats).Decisions made on the basis of wind tunnel testdata included: selection of optimum slat anglesfor take-off and landing; selection of a trimlocation for the inboard slat; extension of theslat chord at the wing tip from 15% chord to17% chord. Investigation of Reynolds numbereffects on lift and drag was done to supportextrapolation of wind tunnel test data to fullscale.

Figure 17 shows the lift and pitching momentcharacteristics with various trim locations of theinboard end of the inboard slat: “sl1b” is the fullspan slat, “sl2b” an intermediate trim locationand “sl3b” the most outboard trim locationconsidered. This figure shows CLmax reducing asmore of the leading edge is left unprotected. Thebenefit is improved pitching momentcharacteristics at stall.

Figure 17: Effect of varying the location of the inboardslat trim on CRJ-700 configuration; Mach 0.20 Rec=6.5Million. Flaps 45 degrees, slats 28 degrees; it=-9 degrees.

The depth of the pitching moment “bucket”occurring at stall represents the extent of flow

Run 080: sl1bRun 087: sl2bRun 094: sl3b

CL

CM

Alpha

Slat #1

Sl3bSl3

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50 Cts

IAR6166: Run 037 B206 W34w14F100 N73P82 V11TWT627: Run 065 B206 W34w14F100 N73P82

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separation on the inboard wing. The width ofthe bucket indicates the angle of incidencemargin between the initial inboard flowseparation and flow separation on outer parts ofthe wing. Selection the “sl3b” trim location forthe aircraft effectively meant trading off someCLmax for better stall characteristics.

A decision was also made to extend the slatchord at the wing tip from 15% of the wingchord to 17% chord. Figure 18 shows the effectof this extension on the high-lift characteristicsof a landing configuration with the tail off.

Figure 18: Effect of extending the slat chord at the wingtip from 15%c to 17%c on CRJ-700 tail-off configuration;Mach 0.20 Rec=6.5 Million. Flaps 45 degrees, slats 25degrees

Figure 19 shows the effect on CLmax of varyingthe Reynolds number at Mach 0.2 on the finalaircraft configuration. The variation is seen tobe larger on the configuration with retractedslats and much smaller when the slats aredeployed. The results showed that the dataobtained at a reference chord Reynolds numberof 6.5 million was a good basis for predictingfull aircraft behavior.

The wind tunnel test lead to the selection of 25degrees landing slats and 20 to 22 degrees take-off slats. Given the lower accuracy in the dragaxis of the half-model test, a decision was made

to leave the final selection of the take-off slatangle to the flight test phase.

Figure 19: Effect of Reynolds number on variousconfigurations of the CRJ-700; Mach 0.20.

6 Flight Testing

The CRJ-700 aircraft first flew on May 27,1999 (Figure 20). The aircraft obtainedTransport Canada certification on December 22,2000. The aircraft obtained subsequentlyEuropean JAA certification in January 2001 andthe American FAA certified the airplane inFebruary 2001. The aircraft entered revenuepassenger service on February 2001. The high-lift configuration was ultimately validatedduring an extensive flight test program.

Figure 20: The CRJ-700 prototype aircraft #10001 on itsmaiden flight.

Run 176: sl5 (15% chord slat)Run 212: sl4 (15-17% chord slat)

Alpha

CL

CM

3 4 5 6 7 8

cleantake-offlanding

∆CL = 0.10

CLmax

Chord Reynolds number


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The aircraft is fitted with a stall protectionsystem including a stick shaker for stall warningand a stick pusher for stall recovery. The stickpusher activates ahead of natural stall for theclean configuration and post-natural stall for allconfigurations with the slats deployed. Handlingstall tests demonstrated compliance with all stallcharacteristics requirements. These includedstraight and turning stalls at 1 kt/secdeceleration, with power on and off, withspoilers retracted and extended and with andwithout lateral center of gravity imbalance.Dynamic stall entries at 3 kt/sec were alsodemonstrated. Figure 21 shows lift curves fromthe IAR 5-ft wind tunnel test, trimmed for themost forward centre of gravity, compared toequivalent data from performance flight tests.This figure shows good correlation for all flapangles. For take-off configurations requiringgood climb gradient, a slat angle of 20 degreeswas finally selected.

Figure 21: Comparison of trimmed lift curves from theIAR 5-ft wind tunnel test with flight test data.

7 Conclusions

The geometry and deployment schedule of theCRJ-700 flaps and slats were developed usingCFD methods. The predicted characteristics ofthe aircraft were verified in low speed windtunnel tests conducted with a full model at lowReynolds number and a half-model at highReynolds number. There was good correlationbetween the results of the two models, except

for the pitching moments. Tailplane angles totrim and drag of high-lift configurations weretherefore taken from the full-model tests withappropriate corrections for scale. Flight test dataindicated that the results of half-model tests inthe IAR 5ft x 5ft wind tunnel at 6.5-Millionchord Reynolds number provide accurateestimates of the aircraft characteristics includingCLmax. A similar conclusion was reachedpreviously with the Global Express.

Acknowledgements

The authors would like to thank several personswho provided information used for this paper:Dr Vinh Nguyen of IAR for information on the5-ft wind tunnel, Nick Perkins, CRJ-700 ProjectDirector for material on the CRJ-700 aircraftand the staff of Bombardier Experimental WindTunnel Group, in particular. Stephane Boisvertand Guy Bergeron, for information on the windtunnel model.

References

[1] Cebeci, T., Besnard, E. and Messie, S. “AStability / Transition Interactive Boundary LayerApproach to Multi-Element Wings at High Lift”.AIAA Paper 94-0292 (1994).

[2] Drela, M. “Newton Solution of CoupledViscous/Inviscid Multi-Element Airfoil Flows”.AIAA Paper 90-1470 (1970).

[3] Mavriplis, D.J. and Jameson, A. “Multigrid Solutionof the Navier-Stokes Equations on TriangularMeshes”. AIAA Journal, vol. 28, no. 8, pp. 1415-1425 (1990).

[4] Langlois, M., Kho, C., Kafyeke F., Mokhtarian,F.and Jones, D. "Investigation of a Multi-ElementAirfoil High-Lift Characteristics at High ReynoldsNumbers", Proc. 8th CASI Aerodynamics Symposium,Toronto, April 2001.

[5] Valarezo, W.O. and Chin, V.D. “Maximum LiftPrediction for Multielement Wings”. AIAA-92-0401.

[6] Brown, D. “Information for Users of the NationalResearch Council’s f’x5’ Blowndown Wind Tunnelat the National Aeronautical Establishment”. LTR-HA-6. National Research Council Canada.September 1977.

[7] Mokry, M. “Subsonic Wall Interference Correctionsfor Half-Model Test Section Using Sparse WallPressure Data”, NAE Aeronautical Report LT-616,1985.

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