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  • (SYA) 34-1

    Experimental and Numerical Investigation of Vortex Shedding of aRepresentative UCAV Configuration for Vortex Flow Control

    Terence A. Ghee *, Doug R. Hall

    NAVAIR, Code 4321Bldg. 2187, Suite 1320B

    Patuxent River, MD 20670-1906USA

    ABSTRACTA 4% Uninhabited Combat Air Vehicle (UCAV) has been extensively tested at low speeds in a wind tunnel toinvestigate using vortex flow control to control vehicle attitude. The program is the initial step to utilizeexperimental and computational techniques to understand the flowfield environment on a representative low-observable air vehicle and use that understanding to apply an efficient vortex flow control apparatus. Gross flowfield characteristics were identified using flow visualization and the approximate vortex location was determinedfor a number of angles-of-attack for a tunnel dynamic pressure of 26.74 psf. From this study, the model wasinstrumented with pressure transducers at appropriate locations on the wing and unsteady data was acquired for anumber of angles-of-attack and tunnel dynamic pressures. A six-component internal balance was then installedto measure aerodynamic forces and moments. Limited steady electronically scanned pressure data wereacquired. Computational fluid dynamic (CFD) analysis was conducted on the model geometry to compare withthe results from the wind tunnel study. The results show two vortex structures: a weak apex vortex and astronger wing vortex. Wing vortex frequency exhibits a broad-banded dominant frequency of approximately 10Strouhal number. Maximum suction pressure was seen to move forward on the wing leading edge as the wingvortex moved inboard with increasing angle-of-attack. The CFD results adequately predicted the force andmoment data. However, the CFD comparison to the unsteady pressure data was not stellar: CFD frequentlyfailed to predict the mean pressure coefficient and the frequency content of the signal.

    NOMENCLATURE

    * Aerospace Engineer, ph: 301 342 8536 FAX: 301 342 8588, gheeta@navair.navy.mil Aerospace Engineer, ph: 301 342 8534, FAX: 301 342 8588, halldr@navair.navy.mil

    b span, 2.160 ftC coherenceCP pressure coefficient, p-p/qCR root chord, 1.280 ftf frequency, HzG amplifier gainIraw raw integer value (digital)MAC mean aerodynamic chord, 0.765 ftM freestream Mach number, V/ a, 0.13P pressure, psiPRef reference pressure, psiPSD power spectral densityq tunnel dynamic pressure, 26.74 psf

    Re Reynolds number, ( VMAC)/, 0.73E6RMS root mean squarestd standard deviationSe transducer sensitivity, volt/psiS wing area, 1.210 ft2

    St Strouhal number, (b f)/ Vt time, secV excitation voltage, 10 voltV tunnel freestream velocity, 150 ft/s geometric angle of attack, degree viscosity, 3.7373E-7, slug/(ft sec) non-dimensional time, (t a)/MAC

    INTRODUCTIONOn air vehicles with swept wings, leading edge vortices are created at off-design conditions, Ref. 1. The leadingedge vortex generally has a beneficial effect in the form of increased lift at higher angles-of-attack. By

    Paper presented at the RTO AVT Symposium on Advanced Flow Management: Part A Vortex Flows andHigh Angle of Attack for Military Vehicles, held in Loen, Norway, 7-11 May 2001, and published in RTO-MP-069(I).

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    4. TITLE AND SUBTITLE Experimental and Numerical Investigation of Vortex Shedding of aRepresentative UCAV Configuration for Vortex Flow Control

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    7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) NATO Research and Technology Organisation BP 25, 7 Rue Ancelle,F-92201 Neuilly-Sue-Seine Cedex, France

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    13. SUPPLEMENTARY NOTES Also see: ADM001490, Presented at RTO Applied Vehicle Technology Panel (AVT) Symposium heldinLeon, Norway on 7-11 May 2001, The original document contains color images.

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  • (SYA) 34-2

    controlling the location of the shed vortex or vortices, vehicle roll and pitch control may be possible.Uninhabited Combat Air Vehicles (UCAVs) that utilize stealth to avoid detection suffer a radar signatureincrease when control surfaces are deflected. Thus, there is an advantage to be gained by limiting flap deflectionby utilizing vortex flow control to change vehicle attitude. However, a thorough understanding of the vehicleflowfield is needed to best mate a flow control device to the vehicle. In view of this, an experimental andnumerical investigation was conducted on a representative UCAV configuration to define the flowfield andinvestigate methods to control vortex location and, ultimately, vehicle attitude. The test program was developedin three phases: 1) vortex location identification using laser light sheet flow visualization and fluorescent oilapplied to the model surface, 2) vortex quantification through surface and off-body measurements, and 3) vortexmanipulation utilizing a flow control device. This paper reports the results of the first two phases of theprogram.

    EXPERIMENTAL SETUPThe tests were conducted in the Naval Aerodynamic Test Facility (NATF) as part of the Naval Air WarfareCenter In-House Laboratory Independent Research (ILIR) program. The NATF is a four-foot by four-footclosed test section, open-return wind tunnel. The facility incorporates a 200 horsepower motor that drives avariable pitch fan and delivers a maximum velocity of 205 ft/s. In addition, the facility has honeycomb and threesets of flow conditioning screens that minimize freestream turbulence intensity to approximately 0.15% andfreestream velocity differences of 1%. Fig. 1 shows the 4% UCAV model in the NATF.

    Static Pressure probe

    Fig. 1 4% UCAV in NATF

    Sharp Leading EdgeRounded Leading Edge

    Transducers

    Fig. 2 - 4% UCAV Leading Edge

    A 4% UCAV model with 47-degree leading edge was used as a representative configuration. The model wasfabricated of stainless steel for the US Air Force Research Laboratory and tested previously to assessaerodynamic performance at high subsonic and transonic speeds, Ref. 2. The leading edge of the cranked-delta-wing vehicle has a sharp chine at the nose that transitions to round in the vicinity of the wing/body juncture, seeFig. 2. All data reported herein was for the boundary layer transition free. Also visible in Fig. 2 and detailed inFig. 3, are the 13 Kulite fast-response, 5-psi sealed-gauge, pressure transducers installed on the model. An inletnose plug was installed for the majority of the testing to create a more simplified geometry without the addedcomplexity of a flow-through duct. In addition, transducer signal/power cables could be run through the duct,thus avoiding expensive machining of the model. The nose/propulsion inlet plug also allows the testing of novelvortex flow control devices to be easily fabricated and evaluated when affixed to the plug.

    A sting assembly attached to the facility pitch strut supported the model and the support system was constrainedvertically; the model moved off-centerline with changing angle-of-attack. Angle-of-attack was measured at themodel support system using an Allied Signal QA-2000 accelerometer. The model was set to 0.108 degrees in yaw tocorrect cross flow angularity.

    The unsteady pressure data was acquired using a 32-channel, 14-bit DSP Technology, Inc. IMPAX unsteady dataacquisition system connected to a personal computer. This system acquired the data simultaneously and thus thedata was coincident.

  • (SYA) 34-3

    V

    # X/c Y/b

    1 0.047 0.008

    2 0.131 0.036

    3 0.183 0.038

    4 0.161 -0.044

    5 0.263 0.078

    6 (lower surf.) 0.352 0.108

    7 0.377 0.116

    8 0.476 0.148

    9 0.581 0.183

    10 0.665 0.211

    11 0.609 -0.192

    12 1.154 0.132

    13 1.115 0.257

    1

    4

    2,38

    6,7

    5

    910

    13

    12

    11

    Fig. 3 Transducer Location

    WingVortex

    ApexVortex

    a) Video data, view looking upstream. b) Reference grid video, used

    to quantify vortex location .Fig. 4 Vortex Location Using Laser Light Sheet, x/MAC = 0.735

    PROCEDURE AND DATA REDU

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