Wind tunnel investigation of vortex flows on F/A-18 configuration at subsonic through transonic speed

Techn,cal

Paper

Wind Tunnel Investigation v

of Vortex Flows on F/A-18

Configuration at Subsonic

ThroughTransonic Speeds =

t Z ......

Z

NASA

Technical

Paper

Wincl Tunnel Investigation

of Vortex Flows on F/A-18

Configuration at Subsonic

Through Transonic Speeds

Gary E. Erickson Langley Research Center Hampton, Virginia National Aeronautics and Space Administration Office of Management Scientific and Technical Information Program Symbols B swept, three-dimensional laminar separation bubble BL butt line, inches full scale (0.06 scale) b 0.06-scale model reference wing span, 2.245 ft drag coefficient, D_

CD

q_D Lift

CL

lift coefficient, body-axis rolling-moment coefficient, Rolling moment

A

q_c Sb Cm pitching-moment coefficient referenced to 0.255, Pitching moment q_S_ body-axis yawing-moment coefficient, Yawing moment qocSb forebody surface static pressure coefficient, pbc_l - p_

Cp

q_¢ LEX upper surface static pressure coefficient, Pl°cabu - P_ Cp,u qo¢ pressure coefficient corresponding to local speed of sound, c; 3'+1 -I side-force coefficient, Side force

cz

CFD computational fluid dynamics 0.06-scale model wing mean aerodynamic chord, 0.691 ft DTRC David Taylor Research Center FS fuselage station, inches full scale (0.06 scale) HARV High-Angle-of-Attack Research Vehicle HATP High-Angle-of-Attack Technology Program Inbd. inboard LaRC Langley Research Center LE leading edge LEX wing leading-edge extension LP laminar cross-flow separation pattern LS primary laminar separation 0.06-scale model length measured from nose to exhaust nozzle exit plane, 3.265 ft free-stream Mach number M_ MHB maximum half-breadth NASA National Aeronautics and Space Administration Outbd. outboard local surface static pressure, lb/ft 2 Plocal local upper surface static pressure, lb/ft 2 Plocal,u .o.

PRECEDING PAGE BLANK NOT FILMED free-stream static pressure, lb/ft 2 PoG free-stream dynamic pressure, lb/ft 2 q_c R location of reattaehment Ree Reynolds number based on S 0.06-scale model reference wing area, 1.440 ft 2 SS secondary separation local semispan distance from LEX -fuselage junction to LEX leading edge, ft TE trailing edge TP turbulent cross-flow separation pattern TRP transitional cross-flow separation pattern , TS primary turbulent separation X LEX vortex breakdown location measured from nose along model eenterline, ft distance along LEX local semispan, ft Y Ol angle of attack, deg angle of sideslip, deg ratio of specific heat constants Y leading-edge flap deflection angle measured normal to hingeline, 6LE positive leading edge down, deg 6TE trailing-edge flap deflection angle measured normal to hingeline, positive trailing edge down, deg forebody cross-section angular location (0 ° is bottom dead center, positive is clockwise as seen from a front view), deg iv adverse vortex flow interactions with other airframe

Summary

components.

A wind tunnelexperiment wasconducted in the

NASA is conducting a High-Angle-of-Attack 7- by 10-Foot Transonic Tunnel at the David Taylor Technology Program (HATP) to provide design Research Center (formerly the Naval Ship Research guidelines and new concepts for vortex control on ad- and Development Center) of the wing leading-edge vanced, highly maneuverable fighter airplanes. The extension (LEX) and forebody vortex flows at sub- program consists of wind tunnel testing of subscale sonic and transonic speeds about a 0.06-scale model models of complete aircraft configurations, subscale of the F/A-18. The primary goal was to improve and full-scale models of airplane components, piloted the understanding and control of the vortical flows, simulations, development and validation of computa- including the phenomena of vortex breakdown and tional fluid dynamics (CFD) methods, and full-scale vortex interactions with the vertical tails. Laser va- flight testing. The flight experiments are performed por screen flow visualizations, LEX and forebody sur- with a highly instrumented F-18 as a High-Alpha Re- face static pressures, and six-component forces and search Vehicle (HARV) (fig. 1). The NASA HATP moments were obtained at angles of attack from 10 ° provides a unique opportunity for "closed-loop" cor- to 50 °, free-stream Mach numbers from 0.20 to 0.90, relations of the results from ground-based test facil- and Reynolds numbers based on the wing mean aero- ities, CFD methods, and flight.

dynamic chord from 0.96 x 106 to 1.75 x 106. The wind tunnel results were correlated with in-flight flow A fundamental issue in the NASA HATP is the visualization and handling qualities trends obtained sensitivity of the forebody and LEX vortical flows to by NASA using an F-18 High-Alpha Research Vehi- Reynolds number and Mach number. The degree to cle (HARV) and by the U.S. Navy and McDonnell which the vortical flows in subscale model wind tun- Douglas Corporation on an F-18 airplane with LEX nel testing represent the flow-field behavior in flight fences added to improve the vertical tail buffet en- at full scale is of critical concern. Another issue is vironment. Key issues that were addressed include an apparent model scale effect (ref. 1), when discrep- the sensitivity of the vortical flows to the Reynolds ancies exist between the high-angle-of-attack stabil- number and Mach number; the reduced vertical tail ity characteristics of different scale models tested at excitation, and the corresponding flow mechanism, the same Reynolds number in the same, or different, in the presence of the LEX fence; the repeatabil- wind tunnel facilities. Model surface irregularities ity of data obtained during high-angle-of-attack wind and the scale of the free-stream turbulence relative tunnel testing of F/A-18 models; the effect of parti- to the model size (ref. 2) are factors that may con- cle seeding for flow visualization on the quantitative tribute to the development of global flow fields that model measurements; and the interpretation of off- are sufficiently different to affect the stability lev- body flow visualizations obtained with different illu- els near maximum lift. Techniques for tripping the mination and particle seeding techniques.

wind tunnel model boundary layers at high angles of attack to provide a consistent set of results and Introduction to properly represent the flight characteristics have yet to be developed. The development, interaction, Present-generation fighter airplanes such as the and breakdown of the forebody and LEX vortices, F/A-18 and the F-16 exploit vortex flows for en- their interaction with downstream airframe compo- hanced maneuverability at high angles of attack and nents such as the vertical and horizontal tails, and at subsonic and transonic speeds. The development, vortex-shock interactions are not fully understood.

interaction, and breakdown of the vortices generated The U.S. Navy and the McDonnell Douglas Cor- from the wing leading-edge extensions (LEX's) and fuselage forebodies and the interaction of the vor- poration are also engaged in the development of con- tex flows with shock waves at the transonic speeds cepts to improve the vertical tail buffet environment promote nonlinear aerodynamic and stability char- on the F/A-18 (ref. 3). The interaction of the burst LEX vortices with the twin vertical tails on the acteristics that are difficult to predict and control.

In addition, the interaction of the vortical flows F/A-18 creates a buffet environment that is severe with vertical and horizontal tails can induce a severe enough to cause structural fatigue. The excitation of tail buffet environment leading to structural fatigue. the vertical tails occurs in the range of angle of attack from approximately 16 ° to 44 ° . The most critical ef- F-18 and F-16 derivatives and new-generation fighter fect of the burst vortices is manifested in the tail sec- airplanes will continue to employ vortex-lift concepts.

As a consequence, the understanding, prediction, and ond, or outboard, bending and torsional mode. The control of these phenomena are essential to optimize second mode response is most severe in the range airplane maneuverability and to reduce or eliminate of angle of attack from approximately 20 ° to 30 ° .

TheMcDonnell Douglas Corporation andthe Naval of the off-body flows with the model surface pres-

Air Systems Command conducted anextensive series

sures and forces and moments, and comparisons of of wind tunnelandflight experiments (ref.3) aimed the wind tunnel results to in-flight flow visualizations

at reducing the verticaltail vibrationenvironment

and handling qualities trends on the F/A-18 aircraft.

at high angles of attack. The resultof theseefforts

wasthe development ofa streamwise fence mounted

Experimental Investigation

to the uppersurface of the wingleading-edge exten-

sions.Theseeffortsculminated in a full-scale flight

Model Description and Test Apparatus

validationprogramfeaturingan F/A-18 with LEX

The testing was conducted with a 0.06-scale

uppersurface fences. The fleetairplanes havebeen

model of the F/A-18, which is illustrated in fig-

retrofittedwith the LEX fences. Photographs of the

ure 4. The baseline configuration corresponded to

fences installed ontheF/A-18airplane arepresented

the model with 34 ° leading-edge flap deflection, 0 °

in figure2. The fences weredemonstrated to signif-

trailing-edge flap deflection, -9 ° horizontal stabilizer

icantlyimprove the verticaltail second bending and

deflection, 0 ° rudder deflection, single-place canopy,

torsionalmoderesponse. Verticaltail accelerome-

and wingtip-mounted missiles. The model featured

ter dataobtained on the NASAF-18HARV(ref.4)

flow-through engine inlets and a distorted aft fuselage

are presented in figure3. Theseresultsconfirmed

assembly to allow the installation of a sting between

the significantreductionin the verticaltail buffet

the twin exhaust nozzles.

with the LEX fences installed. The improved ver-

tical tail buffetenvironment, the minimalimpacton

The forward fuselage, consisting of the forebody,

the lateral-directional stability and aircraft perfor-

LEX's, and canopy, was removable and was in-

mance, andtheease of implementation ledto the se-

strumented to measure surface static pressures at

lectionof the LEX fences for installation onthe fleet

141 pressure orifices. The forebody pressures were

airplanes.The flowmechanism associated with the

measured at FS I07 (6.42), 142 (8.52), and 184

LEX fences wasnot identifiedduringthe wind tun-

(11.04), whereas the pressures on the port and star-

nelandflight experiments. In addition,a forebody-

board LEX's were obtained at FS 253 (15.18), 296

LEX flowinteractionoccurred whenthe flight test

(17.76), and 357 (21.42) on the upper surface only.

noseboomandLEX fences wereinstalledthat led

The pressure ports at each station were selectcd to to degraded handlingqualitiesnearmaximum lift.

maximize the resolution in the vicinity of the vorti-

This effectwaseliminated uponthe removal of the

cal flows and in areas of expected large pressure gra-

noseboom. A determination of the corresponding

dients. The wind tunnel model forebody and LEX

flow-fieldinteractions wasnot madeduringthe LEX

pressure measurement stations are indicated in fig- fencedevelopment program.

ure 5. The fuselage station locations are identical to those on the NASA F-18 HARV (fig. 6). The pres-

In supportofthese programs, andtoaddress some

sure port locations at each fuselage station on the

of theseissues, a cooperative experiment involving

0.06-scale model are a subset of those on the full-

NASA,the U.S.Navy,andthe McDonnell Douglas

scale airplane.

Corporation wasconducted with a 0.06-scale model

of the F/A-18in the DavidTaylorResearch Center LEX fences that were representative of those on (DTRC)7- by 10-Foot Transonic Tunnel.Theprin- the fleet airplanes were designed and fabricated. The cipal objectiveof the testingwasto document the fences are fixed devices mounted in a streamwise ori- forebody andLEX vortexflowcharacteristics at sub- entation on the upper surface of the LEX's. The fence is normal to the LEX surface at the fence lead-

sonicandtransonic speeds oftheF/A-18model with

and withoutthe LEX fences anda flight test nose

ing edge (FS 378.17 (22.69)). This results in an angle boom.Thisobjective wasaccomplished by conduct- relative to the vertical plane of symmetry of approx- ing detailed off-bodyflowvisualizations with a laser imately 25 ° . This angular position was maintained

vaporscreen technique andby measuring the fore-

along the length of the fence. The geometry details bodyandLEX surface staticpressures andmodel six- and location of the fences are provided in figure 7.

component forces andmoments at free-stream Mach The LEX fence concept was developed after the fab- numbers from0.20to 0.90,Reynolds numbers based rication of the pressure-instrumented forward fuse-

on the wing aerodynamic chordfrom 0.96 x 106 to

lage component. As a result, there were no LEX 1.75 x 106 , and angles of attack from 10 ° to 50 ° . surface pressure orifices in the immediate vicinity of The present paper emphasizes the improved under- the fences. The aft pressure row on each LEX was standing of the forebody and LEX vortical flows from situated 21.17 in. full scale (1.27 in. model scale) up- the laser vapor screen flow visualizations, correlations stream of the fence leading edge.

The detailsof the flight test nose boomarepre- laser used in the present experiment was situated in- side the wind tunnel control room. The beam was

sented in figure8. The 0.06-scale modelnoseboom

wasrepresentative of that usedin the initial flights directed through an observation window, into the of theNASAF-18HARVandduringtheearlystages plenum, and to a series of mirrors mounted along the of the LEX fence flight validation program conducted tunnel sidewall and ceiling. After passing through a beam contractor, it was then directed to an optics by McDonnell Douglas and the Navy. The nose boom of the wind tunnel model did not include the angle-of- package consisting of a sheet generator and rotat- ing mirror mounted inside a box beam directly above attack and sideslip vanes on the airplane nose boom.

the test section. The laser light sheet was directed The model six-component forces and moments toward the model through a long, rectangular ceil- were measured with an internally mounted strain- ing window that was offset slightly from the tunnel gauge balance. Angle-of-attack measurement devices centerline. Two rotator stages in the optics package were installed in the model support system, and the allowed the continuous variation of the light-sheet measurements were corrected for balance and sting width and its location along the model. Since the deflection under load.

light sheet swept along an arc, orthogonality of the High angles of attack were obtained with the light sheet with the model was precluded except at a DTRC roll sting arrangement shown in figure 9.

preselected condition of a = 30 ° and the 50-percent .

Angles of attack from 10 ° to 20 ° were obtained by wing chord station. This was an acceptable compro- pitching the model about the main support system mise in order to illuminate the entire model ftow field boom pivcit point. Rotating about the roll sting through the ranges of angles of attack and sideslip.

pivot point provided angles of attack from 20 ° to 50 ° .

The laser vapor screen flow visualizations in the Within the latter angle-of-attack range, the model present paper were documented with two color video moved continuously upward through the test section, cameras. A color video camera with remotely con- as sketched in figure 10. At an angle of attack of 50 °, trolled zoom lens was mounted to a tilt/pan mech- the model nose was approximately 12 in. from the anism situated outside the test section. The flow tunnel ceiling.

field was observed through a window located down- The off-body flow visualization was conducted stream of the model as sketched in figure 10. A video with a laser vapor screen technique (ref. 5). Water camera having a fixed, 12.5-mm lens was mounted in sufficient quantity was injected into the settling to the model sting support (specifically, to the sting chamber by using a spray nozzle which increased the adapter shown previously in fig. 9) and viewed di- relative humidity to create condensation within the rectly between the twin vertical tails of the F/A-18 vortical flows about the model. The vortex cross sec- model. The field of view of this camera was fixed and tions were visualized with an intense sheet of laser was independent of the angles of attack and sideslip.

light. For contrast, the model and tunnel test sec- Four additional video cameras and eight 70-mm and tion sidewalls were painted flat black. The conden- 35-mm still cameras were also mounted at selected sation within the vortices was frequently observed locations in the tunnel sidewall and ceiling windows.

along most of the vortex core length at subsonic However, none of these cameras tracked the model speeds by using the wind tunnel test section lights through the complete ranges of a and _3. The results located in the test section corner fillets. Previous obtained with the video camera units were superior NASA experiments in the DTRC facility with the and, consequently, still photographs were taken from vapor screen technique featured a laser light sheet the video monitor. These results are included in the directed from the right side of the tunnel test sec- present paper. The model pressures, forces, and mo- tion (refs. 6 and 7). This approach was inadequate ments were obtained with and without the camera for the present application, since a large portion of mounted to the sting adapter. The upstream influ- the model flow field would be in the shadows created ence of the camera on the high-angle-of-attack flow by the fuselage, canopy, and twin vertical tails. In field was found to be negligible.

addition, the upward movement of the model as the Wind Tunnel Facility and Test Conditions angle of attack increased made it impossible to track the flow field with a fixed set of optics in the test The laser vapor screen results and model force, section window. Accordingly, the laser optics pack- moment, and surface static pressure measurements that are presented in this paper were obtained in the age was modified to allow the light-sheet generation from the tunnel ceiling. Locating the laser head in DTRC (formerly NSRDC) 7- by 10-Foot Transonic the low-pressure environment within the plenum sur- Tunnel located in Bethesda, Maryland. The DTRC rounding the test section was precluded due to laser facility is a continuous-flow, closed-circuit facility operational concerns. Instead, the 5-watt argon-ion capable of operating over a Mach number range

from0.20to 1.17andanequivalent pressure altitude

sualizations are presented along with the model sur-

rangefrom sealevel to 40000feet. A complete

face static pressure distributions and six-component

description of the transonic wind tunnelis provided

forces and moments. The model without the fences

in reference 8. The 0.06-scale F/A-18 modelwith

and nose boom is referred to as the "baseline config-

andwithout the LEX fences is shownmountedto

uration." Comparisons are made of the wind tunnel

theroll stingarrangement in the slottedtestsection

flow-field observations and available in-flight flow vi-

in figure11. Typicallaser-illuminated vorticalflows

sualizations on the NASA F-18 HARV. Experimental arealsoindicated in figure11.

results from other F/A-18 model tests conducted by McDonnell Douglas and NASA are also used on a

The test resultswereobtainedat free-stream

limited basis to support some of the conclusions of

Machnumbers from 0.20to 0.90. The angleof at-

the present investigation.

tackwasvariedin 2.5 ° increments from 10 ° to 50 °.

Pitch polarswereobtainedat sideslipangles of 0°, The forebody pressures at FS 107 (6.42), 142

4°, and8°. Sideslip "sweeps" in 2° increments from

(8.52), and 184 (11.04) are shown as a function of

-10° to +10° wereobtained at selected angles of at-

angular position 0, where the orientation is that of

tackandMachnumbers. Themaximum free-stream

an observer standing in front of the model. A value

dynamic .pressure duringthe testwasapproximately

of 0 of 0 ° corresponds to the bottom centerline; 0 in- 250 lb/ftZdue to a normal force limit of 1000 lb im- creases in the clockwise direction. The LEX surface posed on the DTRC roll sting arrangement. For free- pressures at FS 253 (15.18), 296 (17.76), and 357 stream Mach numbers from 0.20 to 0.40, the testing (21.42) are plotted against the local semispan dis- was conducted at atmospheric conditions. The tun- tance y measured from the LEX-fuselage junction, nel was operated in the evacuated mode (ref. 8) at the normalized by the local distance s from the LEX- higher Mach numbers. The tunnel stagnation pres- fuselage junction to the LEX leading edge. For the sure varied with the Mach number, ranging from ap- starboard LEX, values y/s of 0 and 1 correspond to proximately 1250 lb/ft 2 at Af_ = 0.60 to 750 lb/ft 2 the LEX-fuselage junction and LEX leading edge, at M_c = 0.90. The Reynolds number based on the respectively. Similarly, values of y/s of 0 and -1 co- wing mean aerodynamic chord Re_ varied from ap- incide with the port LEX-fuselage junction and LEX proximately 0.96 × 106 to 1.75 × 106.

leading edge, respectively. Sketches of the LEX and forebody pressure orifice orientations are presented The model force, moment, and surface pressure in figure 12. It is noted that the resolution of the data were obtained on the baseline configuration un- LEX pressure distributions on the port and starboard der "dry tunnel" conditions at the outset of the sides was the same at FS 253 (15.18) but differed at experiment. Thereafter, the data were obtained si- FS 296 (17.76) and 357 (21.42). In addition, the multaneously with the laser vapor screen flow visu- pressure port density along the windward forebody alizations. This allowed an assessment of the sensi- surface was reduced in order to increase the pres- tivity of the quantitative measurements to the water sure resolution in areas where the largest pressure injection.

gradients were expected, namely, near the maximum The model featured boundary-layer trip strips on half-breadth (MHB) and on the lec side underneath the forebody, LEX's, wings, tails, and inlet ducts. the vortices.

The trip strips were developed by McDonnell Dou- The technical discussion is divided into five ma- glas and consisted of epoxy cylinders that were jor sections. The first section compares the model bonded to the model surface. The epoxy cylinders pressure distributions obtained with and without wa- had a nominal diameter of 0.050 in., spacing be- ter injection (for flow visualization) into the tunnel tween cylinders of 0.025 in., and height of 0.0035 in.

circuit. The second section addresses the repeata- (0.06 scale). A trip ring was applied to the forebody bility of the data obtained from other wind tunnel about 0.40 in. (0.06 scale) from the nose tip, and models of the F/A-18. The forebody and LEX sur- a trip strip was installed along the entire forebody face pressures on the 0.06-scale model tested in the length at the bottom centerline. The trip strips on DTRC facility are compared with previously unpub- the LEX's, wings, tails, and inlet ducts were located lished data obtained on the same model in the Lang- 0.40 in. aft of the component leading edges.

ley 14- by 22-Foot Subsonic 'lSmnel. The 0.06-scale model data are also compared with the results ob- Discussion of Results tained in reference 9 with a 0.16-scale F/A-18 model Representative results obtained in the DTRC in the Langley 14- by 22-Foot Subsonic Tunnel. The 7- by 10-Foot Transonic Tunnel are presented in the third, fourth, and fifth sections present the off-body following sections. The laser vapor screen flow vi- flow visualizations, surface pressures, and forces and moments obtained on the baseline configuration, the mum lift, which occurs at an angle of attack of ap- model with LEX fences, and the model with flight proximately 40 °. Bursting of the LEX vortices dom- test nose boom, respectively. inates the flow about the LEX's and wings, and the forebody and LEX vortices interact with each other.

Effect of Water Injection on 0.06-Scale The flow about the forebody of the F/A-18 can be F/A-18 Model Surface Pressures sensitive to model surface irregularities, tunnel flow Figures 13 through 16 present the baseline model conditions, Reynolds number, and Mach number. It forebody and LEX surface static pressure distribu- has been conjectured in reference 10 that even sub- tions at Moc = 0.90 and 0.60, respectively, at se- tle differences in the primary boundary-layer separa- lected angles of attack corresponding to conditions tion along the forebody and, consequently, the fore- when the tunnel circuit was dry at the outset of the body primary vortex strengths and locations may be testing and to conditions when water was injected amplified downstream as the forebody vortices in- for the vapor screen flow visualizations. Water was teract with the LEX wing flow field. Although the injected into the tunnel in a sufficient amount to in- F/A-18 forebody shape is not conducive to the de- crease the relative humidity in the test section such velopment of powerful vortices, they may be of suffi- that the water vapor condensed within the vortical cient strength to affect the flow field about the LEX's flow regions above the model. The flow visualization and wings and, hence, the lateral stability character- run at Mzc = 0.90 was conducted first, since smaller istics. Experience has shown that seemingly minor amounts of water were required to achieve this effect. differences in LEX vortex burst locations in sideslip More water was required at M_ = 0.60. There was at high angles of attack can lead to large differences no instrumentation available in the tunnel to quan- in the lateral stability levels of fighter aircraft models tify the amount of moisture in the flow. (ref. 11).

The region of principal concern was the forebody Tunnel-to-tunnel comparisons. Figures 17 which, in contrast to the LEX, did not have fixed and 18 present the forebody surface static pressures primary boundary-layer separation. The test data in at angles of attack of 40 ° and 50 ° , respectively, ob- figures 13 through 16 show essentially no effect, how- tained on the baseline 0.06-scale F/A-18 model tested ever, of adding water for flow visualization on the in the DTRC 7- by 10-Foot Transonic Tunnel and model pressure distributions. The forebody and LEX the Langley 14- by 22-Foot Subsonic Tunnel. The vortex strengths and locations were unaffected by the Mach number and Reynolds number were identical vapor screen seeding within the range of conditions in both tests (M_ = 0.20, Ree = 0.96 x 106). The considered in the present experiment. At angles of forebody surface static pressure distributions are typ- attack where LEX vortex bursting was known to oc- ically in good agreement. There is a stronger fore- cur over the LEX (a = 30 ° and greater), the vortex body vortex footprint at FS 107 (6.42) and a = 50 ° "footprints" or "signatures" were identical. This re- from the Langley 14- by 22-Foot Tunnel test, but sult is of importance to high-angle-of-attack testing, this effect vanishes at FS 142 (8.52) and 184 (11.04).

since it supports the simultaneous acquisition of off- The model was painted flat black in the DTRC 7- by body flow visualization and quantitative model data.

10-Foot qhannel test. This resulted in increased sur- F/A-18 Model Data Repeatability face roughness in comparison with the glossy black finish on the model in the experiment in the Langley Repeatability of data is a concern in high-angle- 14- by 22-Foot Tunnel. Although no boundary-layer of-attack testing due to hysteresis and the sensitivity trips were utilized during the 14- by 22-Foot Tunnel of the vortical flows to model surface irregularities, test, a boundary-layer trip ring near the nose and the model support system, blockage, wail interfer- a trip strip along the bottom centerline were used ence, tunnel flow angularity, and free-stream turbu- during the DTRC testing. The differences in the lence (ref. 2). Repeat runs in the same wind tun- paint finish and the trip arrangements may account nel entry can yield different results. Similarly, data for the different vortex footprints at a = 50 ° and from separate entries in the same facility, or in dif- FS 107 (6.42). The larger vortex footprint exhibited ferent wind tunnels, with a common model may fail by the 14- by 22-Foot Tunnel data is consistent with to agree. Discrepancies often exist between results a pressure distribution associated with laminar sep- obtained on models of different scales tested in the aration, where the relatively strong vortices formed same or different facilities.

by the laminar separation region are still close to These problems have arisen in the past several the forebody. The much weaker footprint exhibited years during testing of the F/A-18 configuration. Of by the DTRC data is consistent with a transitional principal interest is the lateral stability near maxi- boundary layer. The overall data agreement is

encouraging, however. Theseresultsalsosuggest

(6.42) and 142 (18.52) are insensitive to Mach num-

that the proximityof the modelnose to the DTRC

bers at Met = 0.20, 0.40, and 0.60.

tunnelceilingat a = 50 ° (12 in.) and model block-

Consistent with the forebody pressure distri- age were not significant factors at the low subsonic butions, the 0.16-scale model displays consistently speeds.

higher LEX vortex-induced suction pressures. This effect is shown in figure 23 at a = 40 ° and may be The LEX surface pressures at a = 40 ° in figure 19 due to the increased interaction of the forebody and show reasonable agreement between the 0.06-scale LEX vortices on the larger scale model. The Mach model tests in the DTRC and Langley wind tunnels.

number may also be a factor affecting the surface At this angle of attack, vortex breakdown occurs pressure comparisons. As discussed in later sections, near the second pressure row on the LEX (FS 296 in contrast to the forebody Cp distributions, the LEX (17.76)). The surface pressure signature of the burst vortex-induced surface pressures are sensitive to com- LEX vortex is less pronounced in comparison with pressibility, even at very low Mach nmnbers. This ef- the unburst case. Despite the differences in the test fect is demonstrated in the next section of this paper.

facilities and support systems, the vortex breakdown Nonetheless, these results suggest that the source Of behavior near maximum lift, as inferred from the the model scale effect leading to different high-a sta- surface pressure distributions, was repeatable. It is bility levels on the two models (refs. i and 10) may be noted that the LEX surface pressure resolution was associated with the forebody flow development and not the same in both model tests.

the ensuing interaction of the forebody vortices with the LEX-wing flow field.

Model-to-model comparisons. The forebody Baseline 0.06-Scale F/A-18 surface pressures obtained on the 0.06-scale model in the DTRC tunnel and on the 0.16-scale F/A-18 in Vapor screen flow visualizations. Figure 24 the Langley 14- by 22-Foot Subsonic Tunnel (ref. 9) presents the laser vapor screen cross-flow visualiza- are presented in figures 20 and 21 for angles of attack tions obtained on the baseline model at ct = 20 °, 25 °, of approximately 40 ° and 50 ° , respectively. The free- 30 °, and 35°; M_c = 0.40; and Ree = 1.75 x 106. The stream Mach number in the 14- by 22-Foot Tunnel cross-flow patterns at each angle of attack are shown test of the 0.16-scale model was approximately 0.08, at fuselage stations that bracket the breakdown loca- with a Reynolds number based on the wing mean tion. A stable vortex is manifested as a donut-shaped aerodynamic chord of 0.96 × 106. This was identical structure having low particle density along its axis to the Reynolds number on the 0.06-scale model at and high particle density along the outer core. The_ M_ = 0.20 in the DTRC and Langley tests.

breakdown phenomenon is discernible as an expan- sion, or flaring, of the core, which fills with water Significant differences exist between the forcbody particles due to reverse flow along the axis. The vor- primary boundary-layer separation locations, vortex tex breakdown location at a given angle of attack was positions, and vortex strengths on the two mod- determined from the original videotapes of the laser els. The signatures of the forebody vortices are vapor screen flow visualization. The model is viewed considerably stronger on the 0.16-scale F/A-18 at from a three-quarter right rear position.

FS 107 (6.42) and 142 (8.52) and a = 40 ° and 50 °. The stronger signatures are generally associ- The flow visualization results show the forward ated with more laminar flow. The reason why the advance of the LEX vortex breakdown location as flow is "more laminar" in nature for the 0.16-scale the angle of attack increases. At a = 20 °, vortex model is currently unknown, since both tests had the breakdown occurs at FS 535 (35.1), which is slightly same nominal value of the Reynolds number. Some downstream of the intersection of the vertical tail possible explanations include (1) lower tunnel free- leading edge and the fuselage (FS 525 (31.50)). The stream turbulence at the lower velocities in the test- breakdown position moves to the junction of the LEX ing of the larger model, (2) smoother surface finish on and the leading-edge flap hingeline (FS 434 (26.04)) the larger model, and (3) premature boundary-layer at a = 25 ° . At a = 30 ° , core bursting is over the transition induced on the smaller model by the com- LEX at FS 381 (22.86), which is midway between the paratively larger pressure orifices. The difference in aft LEX pressure row (FS 357 (21.42)) and the LEX- the free-stream Mach number between the two tests wing-leading-edge junction. The latter is defined as is not the likely source of the data disparity. The the point of intersection of the leading edges of the data presented in figure 22, for example, which were LEX and wing when the wing flap is undeflected obtained in the current experiment, indicate that the (SLE = 0°). Vortex breakdown advances forward forebody vortex-induced suction pressures at FS 107 to FS 324 (19.44) at a = 35 °, which is between the second andthird pressure rows(FS296(17.76) and a three-quarter, right rear position at a = 20 ° are shown in figure 27. The details of the cross-flow 357(21.42), respectively).

structure are sensitive to the Mach number, however.

The LEX vorticeson the F-18HARV displaya

The extent of the vortical flow region that is illumi-

similar progression of vortex core breakdown with

nated by the laser light sheet in figure 27 is larger the angle of attack. Figure 25 shows in-flight visu- at the higher Mach number. This is indicated by a alizations from reference 12 taken with a wingtip- growth of the hollow core and a more extensive region mounted 35-ram camera at angles of attack of ap- of condensed water vapor outside the core. Illumina- proximately 20 °, 25 °, 30 °, and 34 °, Moc _ 0.3, and tion of the vortical flows with only the test section Ree _ 13.5 x 106. The sideslip angle in flight var- lights also revealed a larger system of vortices.

ied from approximately 0.25 ° to 0.65 ° . Smoke was injected into the vicinity of the vortex core from a The LEX vortex structure and breakdown behav- port near the apex of each LEX. The smoke flow vi- ior at M_ = 0.60 are shown from a different per- sualization technique employed on the HARV is de- spective in the photographs in figures 28 through 33, scribed in detail in reference 13. It is noted that which were taken from the model sting-mounted the wing flaps on the F-18 HARV are scheduled with camera looking directly upstream between the twin the angle of attack and the Mach number, in con- vertical tails. In each figure, the cross-flow pat- trast to the fixed deflection angles on the wind tun- terns are presented at a fixed light-sheet position nel model (_LE = 34°, 6TE = 0°) • At a = 20 °, at selected angles of attack. The light-sheet loca- the aircraft leading- and trailing-edge flaps are de- tions range from FS 357 (21.42) (aft pressure row on flected to 25 ° and 4 °, respectively. At a = 25.6 ° and the LEX's) in figure 28 to FS 567 (34.02) (near the greater, the flap settings in flight correspond to _LE wing-trailing-edge-fuselage junction) in figure 33. At = 34 ° and _TE = 0°, which coincide with the wind ct -- 15 ° and 17.5 °, the LEX vortices are stable at all tunnel model. The vortex breakdown positions in stations and pass outboard of the vertical tails. The flight at high Reynolds number agree well with the growth of the vortices and their inboard and upward corresponding results obtained on the wind tunnel migration as the angle of attack increases from 15 ° model at much lower Reynolds number. At a = 20 °, to 17.5 ° are particularly apparent at the aft stations.

where the flap settings are different on the wind tun- At _ = 20 °, the vortices continue their migration nel model and the HARV, flow similarity may be im- and burst near the vertical tail (fig. 33). The vortex posed by the presence of the vertical tails in the paths breakdown at c_ = 25 ° is clearly illustrated beginning of the vortices.

at FS 450 (27.00) (fig. 30). Farther aft, the expanded, The LEX vortex breakdown positions obtained on rotating flows envelop the vertical tails. Wind tun- the F-18 HARV (ref. 13) at Moc _ 0.3 and Ree ,_ nel tests and full-scale flight experiments conducted 13.5 x 106 and the 0.06-scale model at Moc = 0.4 by McDonnell Douglas and the Navy (ref. 3) indicate and Ree = 1.75 x 106 are plotted as a function of the that the vertical tail buffet is a maximum at angles angle of attack in figure 26. In the wind tunnel, a of attack of approximately 25 ° to 30 ° . The corre- pulsing of the core, concurrent with the appearance sponding vortex burst positions are contained within of condensate within the core region, was defined as a band, centered about the LEX wing-leading-edge vortex breakdown. This was followed by an expan- junction, of approximately 25 percent of the wing sion of the vortex into a large, funnel-shaped, rotat- centerline chord. Early flights of the F-18 HARV ing flow. The density of condensed water vapor in featured forward-looking video cameras mounted to this turbulent region was very high. The intersection the vertical tails to observe the LEX vortex flows of a line extending from the burst vortex to a surface (ref. 13). The camera vibration was severe at an- normal on the model provided the breakdown loca- gles of attack of about 25 ° and greater, which was an tion x measured along the centerline of the model indicator of the tail excitation. More recent results from the nose. This value was then normalized by obtained in flight on the F-18 HARV (ref. 4), which the model length l extending from the nose tip to included vertical tail accelerometer data, LEX vor- the exhaust nozzle exit plane. The flight results are tex core visualizations, and observations of the ver- presented in a similar manner. The wind tunnel re- tical tail buffet from a chase aircraft, demonstrated sults obtained with the laser vapor screen method fall the severe tail excitation induced by the burst LEX within the band of flight data obtained with natural vortices. The excitation was greatest when the core condensation and smoke injection techniques.

breakdown had advanced to positions similar to those The LEX vortex trajectories and breakdown lo- observed in the wind tunnel. It is interesting to note cations observed in the wind tunnel were similar at that, although clearly visible, the vertical tail dynam- Moc = 0.60. Representative results obtained from ics could not be felt by the pilot.

The flow visualizationresults in figures 28

the inboard movement of the vortical flow along the

through33 alsorevealthe development of numer-

LEX. The constraint on the lateral movement causes ous vorticesalongthe length of the LEX leading an upward displacement of the vortical flow from

edge. They appearin the photographs as distinct

the surface. The overall effects of the boundary

irregularities aboutthe outeredge of the LEX vor-

constraint and the increasing Mach number are a

tical flows,andhavebeenreferred to asshear layer

slight flattening and inboard movement of the vortex

instabilities. Thesevorticesare discussed in refer-

and an upward rotation of the vortex cross section

ences14 through 16 and are characteristic of the

about the LEX leading edge. As a result of the vortex development about slenderlifting surfaces.

altered shape and position of the LEX vortex, the

They havebeenobserved at low Reynolds number

primary flow reattachment induced by the vortex

andlowsubsonic speeds (ref. 14),transonic andsu-

typically occurs along the side or top of the fuselage,

personic speeds (ref.15),andhigh Reynolds number

depending on the angle of attack. At the lower

(ref. 16). At low angles of attack,the vorticesare

Mach numbers, the flow reattaches to the LEX upper

separate and distinct, whereas at higheranglesof

surface. Downstream of the LEX, where the vortex

attacktheymerge to forma central,dominant vorti-

is no longer fed by the boundary-layer separation

cal flow. Therewasnodiscernible movement of the

from the leading edge, the flow cross section becomes

multiplevorticesat a givenmodelstationand an-

approximately circular.

gleof attack. The vortexstructureilluminatedby

The vapor screen results at Mcc = 0.60 and 0.80

thelaserlight sheet in thepresent experiment onthe

from the model sting-mounted camera are shown in

0.06-scale modelis verysimilarto the in-flight (nat-

figure 36. The cross-flow patterns at the vertical

ural condensation) photograph in figure34 (ref. 16)

tail apex (FS 525 (31.50)) are presented at a = corresponding to the full-scale F/A-18aircraft.

15 ° , 17.5 ° , and 20 ° . At each angle of attack, the

Figure 35 presentsthe progression of vortex

condensation pattern enlarges, and the vortex core

breakdown with the angleof attackat free-stream

moves inboard and upward, at the higher Mach

Machnumbers of 0.20,0.40,and 0.60determined

number. At c_ = 20 ° and Mcc = 0.80, there was

fromthe laservaporscreen observations. At Mcc =

no evidence of the vortex core bursting near the tail 0.20, the vortex core and core bursting were first vis- that occurred at M_o = 0.60. At _ = 22.5 ° (flow ible at an angle of attack of approximately 27.5 ° .

visualizations not shown), however, the vortex core At the higher Mach numbers, the cores were vis- exhibited a pulsation, and condensed water vapor ible at much lower angles of attack. The results would intermittently enter the core region beginning in figure 35 indicate that the vortex breakdown at FS 450 (27.00). This effect could be traced characteristics are insensitive to the Mach number downstream to the vertical tails. Concurrently, the fl'om M_ = 0.20 to 0.60. In contrast, the Navier- vortex cross-flow patterns were very unsteady. Thc Stokes computations performed in reference 17 on an onset of the vortex core pulsing advanced to FS 410 F/A-18 forward fuselage component (forebody, (24.60) (near the juncture of the LEX and the wing canopy, and LEX's) at a = 20 ° and M_ = 0.60 rc- leading edge) at _ -- 25 ° and the magnitude of the vealed a high level of compressibility associated with flow unsteadiness increased. These trends coincided the core of the LEX vortex system. The conden- with the development of strong shock waves that sation patterns also change, as noted previously in interacted with the vortices. Unpublished surface oil reference to figure 27. However, the compressibility flow visualizations conducted by NASA in previous effect within the core region is dominated by the ad- testing of the 0.06-scale F/A-18 model in the DTRC verse pressure gradient in the external potential flow 7- by 10-Foot Transonic Tunnel revealed a normal field. The presence of the twin vertical tails may also shock wave situated over the wing and just upstrcam help to "mask" a Mach number effect, particularly of the trailing edge. This shock wave could not at angles of attack where vortex bursting occurs in be discerned directly from the vapor screen flow the vicinity of the tails.

visualizations, however.

Increasing the free-stream Mach number to 0.80 The laser light sheet did reveal a shock wave promotes noticeable changes in the LEX vortex cross- situated above the aft fuselage section between the flow pattern. It has been documented in reference 6 LEX vortices beginning at Mc¢ = 0.80. An intense that the cross section of a wing leading-edge vortex downflow is induced between the vortices, and the is flatter, or lobe-shaped, and the vortex core moves locally supersonic flow recompresses through a shock inboard and closer to the surface at transonic speeds.

wave situated above the fuselage. At c_ = 27.5 ° and However, another factor affecting the vortex cross 30 ° , the cross sections of the burst vortices expanded flow on the F/A-18 is the fuselage, which limits sufficiently over the aft fuselage that they intersected alongtheir inboardedges.Undertheseconditions, of approximately 35 ° . The vortical flows appeared the cross-flow shock wavewasnolongerapparent. larger and stronger at a given angle of attack in com- parison with the result at Moc = 0.60, and they pene-

Theflowvisualizations at Mcc = 0.80 were not of

trated farther into the LEX flow field before being en- sufficient detail to accurately determine the vortex trained by the LEX vortices. The LEX vortices were breakdown location at any angle of attack. This was weaker at the higher Mach numbers, which would al- rendered more difficult by the intermittent nature of low the body vortices to persist farther downstream.

the core flow over the wing. The Ioss of definition of However, the clarity of the forebody vortices at the the vortex cross flow is typical of laser vapor screen transonic speeds is greater than the flow-field visual- flow visualizations at the transonic speeds.

izations on wings at the same conditions. A plausible The effect of sideslip on the vortex cross-flow explanation for the apparent increase in the forebody structure near the twin vertical tails is shown in vortex strength is the development of a cross-flow figure 37 at Moc = 0.60, a --- 20 °, and/3 = 0 ° and 4 °.

shock wave along the forebody side that causes an The asymmetries in the leeward and windward LEX early separation of the primary boundary layer. This vortex core paths and breakdown positions due to is discussed in more detail later.

sideslip are apparent. Sideslip promotes a forward advance of the windward LEX vortex breakdown LEX upper surface static pressure distribu- tions. The effect of the angle of attack on the LEX position and a more extensive wake that envelops upper surface static pressure distributions at FS 253 the windward vertical tail. In contrast, the leeward (15.18), 296 (17.76), and 357 (21.42) is illustrated vortex is stabilized and the core passes outboard of the vertical tail. in figures 39, 40, 41, and 42 corresponding to free- stream Mach numbers of 0.40, 0.60, 0.80, and 0.90, The strengths of the forebody vortical flows on respectively. The LEX vortex breakdown character- the baseline F/A-18 model are less than those of the istics that were derived from the vapor screen flow vi- LEX vortices. The cross-sectional shape and fineness sualizations at Met = 0.40 and 0.60 in figure 24 and ratio of the F/A-18 forebody are not conducive to figures 28 through 33, and the more limited flow-field the development of strong vortex flows. As a conse- information at 5I_ = 0.80 and 0.90, will augment quence, visualization of the baseline model forebody the analysis of the pressure distribution trends.

vortices was limited to angles of attack near and be- The pressure distributions at Mcc = 0.40 and 0.60 yond maximum lift (35 ° and greater), where the vor- in figures 39 and 40, respectively, display a consistent tex strengths were sufficient to create a condensation increase in the vortex-induced suction peaks at a = pattern visible with the aid of the laser light sheet.

10 ° to 25 °. This is consistent with the vapor screen Figure 38 presents a result obtained at an angle of flow visualizations which revealed a stable vortical attack of 50 ° and 3_cc = 0.60, which shows a pair flow over the LEX surface through this range of angle of distinct, donut-shaped vortices at FS 163 (9.78), of attack. The inboard movement of the suction peak which is midway between the second and third pres- as the angle of attack increases reflects the growth sure rings on the forebody. Farther aft, the forebody of the LEX vortex. This migration is less apparent vortices were rapidly entrained into the burst LEX at angles of attack greater than 20 ° due to the vortices and could not be tracked beyond the canopy.

barrier imposed by the fuselage which impedes the At Mach numbers below 0.60, the forebody vor- lateral translation. At a = 30 °, vortex breakdown is tices were not visible, even at angles of attack up between the aft pressure row (FS 357 (21.42)) and the to 50 °. This is a shortcoming of the vapor screen LEX-wing-leading-edge junction (FS 404 (24.24)).

technique, since the condensation of water vapor The approach of core breakdown and the upward is insufficient in the weaker vortices to yield good movement of the vortex promote a decrease in the flow definition. However, alternative particle seeding vortex suction peaks at FS 357 (21.42). At ¢_ = methods can yield vivid visualizations of the fore- 35 ° , vortex breakdown has passed this measurement body vortices. In recent flight experiments on the station. The pressure distributions are flatter, and F-18 HARV (ref. 12), for example, the forebody vor- the maximum suction pressure levels underneath the tices have been clearly seen at Moc = 0.20 to 0.30 vortex exhibit a marked decrease. Similar trends are and a = 30 ° to 50 °. The smoke seeding particles observed at FS 296 (17.76) and 253 (15.18) as vortex were injected from the nose region of the aircraft.

breakdown reaches these measurement stations at This provided sufficient particle density in the region a = 40 ° and 45 °, respectively. The signatures of of the vortex cores to make them visible.

the LEX vortices are weaker at M_ = 0.60. This is consistent with the compressibility effect on the At Mach numbers of 0.80 and 0.90, the forebody vortices were visible beginning at an angle of attack vortex pressure signatures noted in reference 18. It attack of 20 ° and greater. At _ = 40 ° and 50 °

is notedthat the port and starboard LEX surface

(figs. 46 and 47), where vortex breakdown dominates

pressure distributions areasymmetric at M_c = 0.40

the flow about the LEX's, the signature of the burst and a = 50 °. This may be due to asymmetric flow vortex system displays a similar sensitivity to the development from the forebody, which is discussed in Mach number.

the next section.

CFD results have also revealed the compress- Tile compressibility effect on the LEX surface ible nature of the LEX vortex. The density con- pressures is even more apparent at _[_ = 0.80 and tours derived from Navier-Stokes computations on an 0.90 in figures 41 and 42, respectively. The reduced F/A-18 forward fuselage component in reference 17 vortex strength, flatter vortex cross section, and up- corresponding to M_c = 0.60 and a = 20 ° revealed a ward displacement of the vortex are manifested in 40- to 50-percent expansion in the vortex core region figures 41 and 42 as diminished suction levels and compared with that of the free-stream condition. In flatter pressure distributions underneath the vortical addition, the majority of the core flow achieved a flows. The vortex pressure signature is also more con- local Mach number of 0.90 or greater, with a small ical in character at the transonic speeds. This is due supersonic zone occurring near the apex.

to the diminished upstream influence of the trailing- edge pressure recovery at the higher Mach numbers.

Forebody surface static pressure distribu- As a result, the maximum suction pressures are com- tions. Simplified sketches, taken from refercnce 2, parable at a given angle of attack at all three mea- of three basic flow patterns in cross section about surement stations oil the LEX. At FS 357 (21.42), the a body at a high angle of attack are shown in fig- increased suction pressures near the LEX fuselage ure 48. An understanding of these basic flows will junction beginning at (_ = 25 ° coincide with the on- assist in the interpretation of the 0.06-scale F/A-18 set of the vortex-induced primary flow reattachment model forebody pressure distributions. It is conjec- to the fuselage instead of the LEX surface.

tured that the flow about the wind tunnel model fea- In contrast to the results at the lower Mach num- tures elements of all three cross-flow patterns. Fig- bers, it is difficult to infer the position of vortex ures 49 through 56 present thc effect of the angle of breakdown from the LEX pressure distributions at attack on the forebody surface static pressure distri- Met = 0.80 and 0.90. The laser vapor screen vi- butions at Moc ---- 0.40, 0.60, 0.80, and 0.90, cor- sualizations suggested that core bursting occurred responding to FS 107 (6.42), 142 (8.52), and 184 near the LEX wing-leading-edge junction (FS 404 (11.04). To assist in the discussion of the forebody (24.24)) at a = 27.5 ° and approached the aft pressure flow characteristics, the pressure distributions at se- row on the LEX at c_ = 30 °. This would account for lccted angles of attack arc isolated in separate plots.

the loss of the suction peak underneath the vortex at It is noted that the forcbody cross section is circular FS 357 (21.42) as the angle of attack increases from at FS 107 (6.42). At FS 142 (8.52) and 184 (11.04), 25 ° to 30 ° in figures 41 and 42. This progression of thc cross sections are rounded on the top and bottom the burst position with _ is similar to the results at but relatively flat along the sides.

the lower subsonic Mach numbers. At higher angles The first pattern in figure 48 corresponds to the of attack, the vapor screen flow visualizations were case of laminar cross-flow separation (LP), followed inconclusive. The flat pressure distributions along by flow reattachment (R) and subsequent secondary the forward portion of the LEX provide no indication separation (SS). This flow situation may exist along of the approach, or passage, of vortex breakdown.

the nose region of the F/A-18 model. In fact, in- Figures 43, 44, 45, 46, and 47 illustrate the effect flight surface flow visualizations on the F-18 HARV of the Mach number on the LEX upper surface static (ref. 19) suggest the existence of a laminar separation pressure distributions at angles of attack of 10 °, 20 °, zone near the nose tip.

30 °, 40 ° , and 50 ° , respectively. At a = 50 ° , the The second sketch in figure 48 depicts a transi- data were obtained at /¥I_ = 0.20 to 0.60 only. The tional pattern (TRP). Of the three cross-flow pat- DTRC roll sting normal-force limit precluded testing terns, this is the most complicated. Primary lami- to higher Mach numbers at this angle of attack. The nar separation (LS) occurs but the separated shear test results reveal a high level of compressibility of layer becomes turbulent and reattaches to the body the LEX vortex system. A significant decrease in the forming a confined bubble region (B). The flow reat- vortex-induced suction pressures occurs as the Mach taches at the leeside of the bubble and subsequently number increases. The effect of compressibility is separates as a turbulent boundary layer (TS). This apparent at Mach numbers as low as 0.20 to 0.30.

pattern is considered representative of the flow down- The flattening of the pressure distributions at the stream of the nose region of the F/A-18 model.

higher Mach numbers is also apparent at angles of ber, slenderness of the nose, geometric irregulari- Farther aft along the forebody, where the local ties at the nose apex, surface roughness, free-stream body width is greatest, the effective Reynolds num- turbulence, and model support and vibration. The ber is high enough for boundary-layer transition to occur before laminar separation. Thus, no separation asymmetry is amplified at FS 142 (8.52) and 184 bubble forms and the flow separates in a turbulent (11.04). At FS 184 (11.04), a single vortex suction peak is evident. This pressure distribution is consis- manner (TS).

tent with the movement of one primary vortex toward At M_c = 0.40 (figs. 49 and 50), the footprint the forebody surface, accentuating its suction peak, of the forebody primary vortex pair is first apparent and an upward displacement of the second vortex, at all three pressure rings at an angle of attack of with a corresponding loss of its signature in the pres- 35 ° . For reference, the stagnation point at zero sure distribution. The asymmetric flow development sideslip corresponds to an angular position 0 of 0°.

on the forebody was the apparent triggering mecha- Examination of the pressure distribution at a = 40 ° nism for the LEX vortex asymmetries that were man- and FS 107 (6.42) in figure 50 indicates that the ifested in the pressure distributions in figure 39.

flow accelerates around the circular cross section and induces a maximum suction pressure at an angular It is not possible to provide a complete description position 6° above the maximum half-breadth (MHB of the surface flow characteristics on the basis of the at 0 = 90 ° and 270°). A steep pressure recovery, pressure distributions. However, a plausible surface which is indicative of a turbulent boundary layer, flow situation on the F/A-18 model is that the initial occurs on the leeward side of the body, followed by flow separation along the forebody is laminar (LP primary separation about 30 ° from the top centerline cross flow sketched in fig. 48) due to the low local (0 = 180°). The forebody primary vortex footprints Reynolds number. The laminar region is succeeded are indicated by a pair of suction pressure peaks by a transitional pattern (TRP) and finally a fully situated approximately 12 ° on either side of the top turbulent region (TP). (See fig. 48.)

centerline.

The character of the forebody pressure distribu- The character of the pressure distribution at tions at Ms = 0.60 (figs. 51 and 52) is similar to that FS 142 (8.52) is similar. Because of the flatter sides at Moc = 0.40. It is noted that favorable comparisons at this fuselage station, the attached flow suction have been presented in reference 17 of the current ex- pressure maximum is achieved at a lower angular po- perimental results at a = 20 ° and M_ = 0.60 to the sition (approximately 18 ° below the MHB). The en- turbulent flow Navier-Stokes solutions on an F-18 for- suing pressure recovery region is terminated by pri- ward fuselage component. At c_ = 40 ° (fig. 52), the mary boundary-layer separation about 30 ° on either footprint of the forebody primary vortices is appar- side of the top centerline. The forebody vortices in- ent at all three pressure rings. The flow asymmetry duce suction peaks at an estimated angular position that was evident at M_ = 0.40 and a = 50 ° in fig- 15 ° from the top centerline.

ure 50 is reduced at the higher Mach number. A laser vapor screen result corresponding to Moo = 0.60 and At FS 184 (11.04), an attached flow suction pres- = 50 ° was shown previously in figure 38, which sure maximum occurs at 0 -=-60 °. This is followed by revealed a pair of donut-shaped vortices above the a narrow band of uniform surface pressures along the forebody. A flow-field asymmetry was not apparent fiat sides of the forebody and then a second region of in figure 38, which was at a longitudinal station be- local flow acceleration. This pressure ring is in prox- tween FS 142 (8.52) and 184 (11.04).

imity to the apex of the wing leading-edge extension, and the locally accelerated flow is attributed to the Transonic flow mechanisms are manifested in the LEX upwash. Similar to the results at FS 107 (6.42) forebody pressure distributions at Moc = 0.80 and and 142 (8.52), primary separation at FS 184 (11.04) 0.90 (figs. 53 through 56). The principal differences occurs at approximately 30 ° from the top centerline from the results at the lower Mach numbers are and the resultant vortex pair induces suction peaks earlier separation of the primary boundary layer and situated about 9° from the leeward-side centerline.

stronger forebody vortex footprints. At a - 40 ° At higher angles of attack, the pressure distri- and FS 142 (8.52) (figs. 54 and 56), for example, butions are asymmetric. There was no indication the pressure recovery region beginning at 30 ° below the MHB is terminated by boundary-layer separation during the testing that the asymmetry was time de- at an angular position approximately 30 ° above the pendent. At a = 50 ° (fig. 50), a mild asymmetry is MHB. A supersonic expansion occurs along the sides apparent at FS 107 (6.42). Reference 2 has suggested a number of parameters that may cause this asym- of the forebody. This can be seen by comparing metric flow development, including Reynolds num- the experimental surface pressures with the critical pressure coefficient C_ at Mo_ = 0.8 and 0.9. It is ward vortex over the wing surface, the lift remains hypothesized that the flow recompresses to subsonic higher on the windward side, creating a stable rolling conditions through a cross-flow shock wave that is moment. The correlation of the vapor screen re- strong enough to separate the boundary layer. This sults with the yawing-moment characteristics is more flow situation is sketched in figure 57. The shock- straightforward, however. The burst windward vor- induced primary flow separation promotes stronger tex blankets the windward vertical tail, which pro- vortices in comparison with the subsonic results.

motes the directional instability at small sideslip an- The pressure distribution trends are consistent with gles at a = 20 °. At higher angles of attack where the laser vapor screen flow-field observations at the bursting of both LEX vortices occurs, the yawing- transonic speeds, where the vortices became visible moment variation with sideslip becomes increasingly unstable.

at lower angles of attack and appeared stronger and larger relative to their counterparts at the subsonic The 0.06-scale model exhibits a reduction in lat- speeds.

eral stability at small sideslip angles at c_ = 20 ° and Mc¢ = 0.80 and at a = 20 ° and 25 ° and Moc = 0.90 Longitudinal and lateral-directional char- (figs. 60 and 61). This is caused by the interaction acteristics. Figure 58 presents the lift, drag, of the LEX vortex with a rear shock wave over the and pitching-moment characteristics of the baseline wing, which promotes a rapid forward advance of model at free-stream Mach numbers from 0.20 to core bursting on the windward side. It is not known 0.90. In contrast to the sensitivity of the LEX and whether this interaction persists to higher angles of forebody surface pressures, the character of the lift attack, since the vapor scrccn flow visualizations were and drag curves is similar through the range of the not of sufficient detail to identify the flow mecha- Mach number. Maximum lift is obtained at c_ -- 40 ° nisms. It was apparent, however, that the flow-field for all Mach numbers. The principal influence of asymmetries were significantly reduced at the higher the Mach number is reflected in the pitching-moment model attitudes, which compares favorably with the curves. Increasing the Mach number promotes a sta- recovery in lateral stability indicated in figures 60 ble pitching-moment increment at a given lift coef- and 61 at the higher angles of attack.

ficient, and the pitch stability at low lift levels is increased at the transonic speeds (kl_c -- 0.80 and 0.06-Scale F/A-18 With LEX Fences 0.90).

Laser vapor screen flow visualizations. The Figures 59 through 61 show the variations of laser vapor screen results at the subsonic and tran- the rolling-moment, yawing-moment, and side-force sonic speeds indicate that the LEX vortex is situated coefficients with the sideslip angle at selected angles outboard of the fence at a -- 10 ° (flow visualization of attack and free-stream Mach numbers of 0.60, not shown). A single, primary vortex from each LEX 0.80, and 0.90. At Moc = 0.60 (fig. 59), the model is apparent over the wings and outboard of the ver- exhibits a stable variation of the rolling moment tical tails. The inboard and upward movement of with the sideslip angle at all angles of attack from the vortex at a -- 12.5 ° (vapor screen result not pre- 20 ° to 40 ° . This trend is similar to the results sented) places the fence in the path of the core. This obtained in previous testing by NASA, the Navy, and marks the onset of major changes to the cross-flow McDonnell Douglas with the 0.06-scale model, which structure about the wings and vertical tails. The has consistently displayed lateral stability at subsonic solid boundary introduced into the flow changes the speeds through the range of angle of attack. This pressure field about the LEX's and wings. The fence result differs from the data obtained at low speed on also disrupts the secondary boundary-layer separa- a 0.16-scale F/A-18 (ref. 9), which show a reduction tion on the LEX upper surface. The vortex-induced in lateral stability at angles of attack near maximum reattached flow impinges on the inboard surface of lift.

the fence and is diverted upward; this creates a lo- cal upwash. A modified mechanism of vortex devel- The laser vapor screen flow visualizations pre- opment from the LEX is established in response to sented previously in figure 37 revealed asymmetries these effects.

in the LEX vortex core paths and breakdown po- sitions due to sideslip. The qualitative information Figure 62 presents representative off-body flow vi- can be misleading, however, when compared with the sualizations obtained on the 0.06-scale F/A-18 model total forces and moments. The LEX vortex on the with and without the LEX fences at an angle of at- windward side is stronger and closer to the LEX and tack of 20 ° and a free-stream Mach number of 0.40.

wing surfaces relative to the leeward vortical flow The flow field is viewed from a three-quarter, right (ref. 20). Despite the early bursting of the wind- rear position. The LEX vortex is displaced inboard andupward asit passes overthefence. Thefence im- result is a reduction in the normal, lateral, and axial pedes the development of the primaryvorticalflow velocity components in the vicinity of the vertical from the leading-edge extension and effectively re- tails. In the range of angle of attack from 25 ° duces the vortexgenerating length.Thetermination to 30 ° , where the vortex-induced tail excitation is

of the vortexfeedingmechanism weakens the LEX

maximum on the baseline configuration, the LEX

vortexin comparison with the baseline case.In ad-

fences promote a flow at the tails consisting of two dition,the vortexshears awayfromtheleading edge corotating vortices of reduced strength that induce

andbecomes a "freevortex"system upstream of the

lower mean velocities and flow angularity. This effect

LEX wingjunction. Thevortexcross section is dis-

was quantified in low-speed wind tunnel testing by

torted downstream of the fence. Concurrent with

McDonnell Douglas of an earlier version of the LEX

the compression andstretching of theLEX vortexis

fence. As shown in figure 66, the mean velocities and

a downward and outboard movement of the vortical

flow angularity in the vicinity of the vertical tail of a flow as it passes over the wing surface.

0.083-scale F/A-18 model are reduced with the fences on.

The distortion and displacement of the LEX pri- mary vortex downstream of the fence are due to The laser vapor screen flow visualizations pro- its interaction with another vortical flow. A second vided further qualitative evidence of the effectiveness corotating vortex (vortex rotating in the same sense) of the LEX fences in improving the vertical tail .buf- fet environment. The LEX vortex burst phenomenon develops from the LEX leading edge. This is illus- trated in the close-up photographs of the LEX and observed on the baseline configuration at angles of wing regions in figure 63 at M_c = 0.60 and _ = 20 ° attack of approximately 20 ° and greater was char- and 25 °. The ftow phenomena at this higher Maeh acterized by a sudden flaring of the vortex core, the number were similar to those at Moc = 0.40 but were appearance of condensate along the core axis, and a revealed in greater detail. The origin of the second marked flow unsteadiness within the region of burst- vortex is near the point at which the main LEX vor- ing. The breakdown locations from the vapor screen tical flow shears away from the edge. The effective flow visualizations were clearly defined. In contrast, a burst location at c_ = 20 ° on the model with the generating length of the second vortex extends from this point aft to the intersection of the LEX with the LEX fences could not be identified. Repeated sweeps leading-edge flap hingeline. In practice, this generat- of the laser light sheet failed to pinpoint the telltale ing length is not constant, since the leading-edge flap signs of vortex breakdown. At higher angles of at- deflection angle is scheduled with the angle of attack tack, c_ = 25 °, for example, the breakdown was iden- and tile Math number. In the present experiment, tified as a very mild, gradual process, and the level of the leading-edge flap is deflected to its maximum an- turbulence within the expanded, rotating flows was gle of 34 ° , which exposes the longest possible run visibly reduced in comparison with the baseline flow length for the second vortex. The corotating vor- field. At angles of attack of about 30 ° and greater, tex induces downward velocities at the LEX primary where the breakdown position advanced forward to vortex, which compresses and stretches the LEX vor- a position at, or ahead of, the fence, the burst phe- tex and draws it downward toward the wing surface. nomenon was very similar to the baseline case.

This effect is seen from the perspective of the sting An early concern with the fence was the possibil- camera in figures 64 and 65, which present the cross- ity of prematurely bursting the LEX vortices due to flow patterns on the LEX fence and baseline config- the physical obstacle present in the flow. The fence urations at Moc = 0.60 and _ = 20 ° and 25 °, respec- that was tested on the 0.06-scale model was selected tively. As the LEX vortex path moves downward, it from hundreds of configuration modifications tested also bends outward slightly. Tracking the respective by McDonnell Douglas (ref. 3). The fact that the vortex trajectories along the wings and near the tails fence did not significantly impact the longitudinal reveals a slight rotation of the vortical flows about or lateral-directional characteristics obtained in the each other, although the interaction at zero sideslip earlier testing is an indicator of the benign nature of angle is not strong enough to promote a coiling, or the device. Nonetheless, special attention was given wrapping around, of the vortices. The results in fig- in the laser vapor screen flow visualizations to the ures 64 and 65 demonstrate the significant change in primary LEX vortex core stability in the presence of the cross-flow structure at the vertical tails due to the fence. At Moc = 0.60, the results show a delay of the LEX fences.

vortex core breakdown due to the fence at angles of The velocities induced by the corotating vortices attack up to approximately 27.5 ° . The delayed burst- on each other are in opposition. In combination ing is apparent in the sting camera results shown pre- with the reduced vortex strengths, the expected net viously in figure 65 at a = 25 °. At FS 450 (27.00),

the vortexcoreon the baseline modelis burst. The

in comparison to the single, stronger vortex on the

coreisfilledwith condensate asa resultofthereverse

baseline model.

flowalongits axis.With the fences on,the LEX vor-

Prior comparisons of the baseline model LEX vor-

tex at this stationexhibitsa stable,hollowcore.A

tex breakdown behavior obtained with the laser va-

plot of the progression of vortex burstingwith the

por screen method with the flight results obtained

angleof attackfor the baseline andfenceconfigura-

on the NASA F-18 HARV using smoke have been

tionsis shown in figure67.Themorestablebehavior

straightforward. The structure of vortex breakdown

of theLEX vortexin the presence of the fence is due

has been of the classical sense in both cases, namely,

to thefavorable flowgradients induced by thesecond

a sudden expansion of a well-defined vortex core, fol- corotating vortexfromthe rearportionof the LEX.

lowed by a large, turbulent rotating flow downstream.

Oncevortexbreakdown reaches the fence, however, However, differing interpretations of the F/A-18 flow

the vortexbreakdown progression with the angleof

field with the fences installed have arisen as a result attackis similaron bothconfigurations.

of recent flight tests of the F-18 HARV (ref. 4). A representative result from flight at a Maeh number of The favorable interference oftwosuitablyspaced, 0.27 and an angle of attack of approximately 20 ° is

corotatingvorticeshas beendocumented in refer-

shown in figure 68. The smoke particles were intro-

ences 6and 7 on a generalized 55 ° cropped delta wing

duced into the vortex from a port located near the fighter model with chine-like forebody strakes. Under apex of each LEX. The vortex core in flight is well- certain conditions, it was found that the eorotating defined along the LEX, and it moves upward and vortex system delayed vortex breakdown and reduced inboard in the vicinity of the fence. Aft of the fence, the undesirable effects of bursting once it occurred.

however, the smoke particles define a larger, diffused In addition, the altered cross-sectional shape of in- flow that moves downward and outboard over the teracting vortices has been reported in reference 7 on wings. Videotape results from a wingtip-mounted the 55 ° cropped delta wing model and in reference 21 camera revealed a rotating mass downstream of the on a 65 ° cropped delta wing LEX configuration.

fence that was less turbulent than the clearly defined vortex breakdown on the aircraft without the fences.

The interpretation of the off-body flow visualiza- This result is sinfilar to the wind tunnel model con- tions warrants special note. The initial wind tunnel densation pattern at M_ = 0.4 illuminated by the results on the model with the LEX fences wcrc ob- test section lights only, as shown in figure 68. These tained at free-stream Mach numbers of 0.30 and 0.40.

results indicate that the traditional interpretation of Good definition of the LEX primary vortex core was vortex breakdown as a rapid expansion of the vortical provided by condensation within the vortex, in com- region does not apply in flow situations such as this.

bination with the lighting of the wind tunnel test The smoke traces in the flight visualization yield a section only. The second corotating vortex could not subset of the overall vortical flow field. Other flow be seen under these conditions, however. Upstream visualization techniques are necessary to extract ad- of the fence, the LEX primary vortex was visible as ditional information contained within the flow field.

a donut-shaped structure, void of condensed water vapor along its axis, with an accumulation of con- LEX upper surface static pressure distribu- densate along the outer edge. Slightly downstream tions. Figures 69, 70, and 71 show the effect of the of the fence, the vortex appeared to expand and the fences on the LEX uppcr surface static pressures at definition of the core region deteriorated, or was lost selected angles of attack and free-stream Mach num- altogether. A similar flow pattern would be obtained bers of 0.30, 0.60, and 0.80, respectively. The last in a smoke flow visualization experiment if the seed- pressure row on the LEX's is approximately 1.27 in.

ing particles were injected near the LEX apex into (0.06 scale) ahead of the fence leading edge.

the vicinity of the vortex core, ms is the case on the F-18 HARV in flight. A likely interpretation of re- At 2_[oc = 0.30 (fig. 69) and angles of attack of sults obtained in this manner is that the fence pro- 20 °, 25 °, and 30 °, the fence promotes a reduction in motes vortex breakdown. Illuminating the model the LEX primary vortex-induced suction peak at the cross flow with an intense sheet of laser light yields aft pressure row (FS 357 (21.42)). This is consistent an entirely different interpretation, however. A dis- with the upward displacement of the vortex core tinct LEX primary vortex core could be discerned that was observed in the laser vapor screen flow downstream of the fence and the second vortex from visualizations. There is no effect of the fence at these angles of attack on the surface pressures at the rear portion of the LEX was visible. The fence promotes a system of two weaker vortices and, conse- the two upstream measurement stations (FS 253 quently, the condensation patterns will be less vivid (15.18) and 296 (17.76)). This trend is maintained at c_----35 °, wherevortex burstinghas advanced angle at selected angles of attack and Moo = 0.60 are upstreamof FS 357(21.42). The fencecauses an shown in figure 75. The data show relatively minor effects of the LEX fences on the lateral-directional

upwarddisplacement of the burst vortexat higher

angles of attack,and the reduced signatureof the stability through the range of angle of attack.

burst vortexpropagates forwardto FS 296(17.76)

O.06-Scale F/A-18 With Flight Test Nose

at c_ = 40 °. There is no indication from the surface

Boom (LEX Fences on) pressures that the fence promotes early bursting of the vortical flow.

Laser vapor screen flow visualizations. Fig- ure 76 shows the effect of the flight test nose boom on The upstream influence of the fence diminishes the forcbody cross-flow pattern at an angle of attack at AIM = 0.60 and 0.80 (figs. 70 and 71). This is of 50 ° and a free-stream Mach number of 0.60. The attributed to the weakening of the vortical flows at light sheet is positioned at FS 184 (11.04). The re- the higher Mach numbers and the development of sult at a = 50 ° was selected because it illustrates to regions of supersonic flow that limit the upstream a larger scale the trends that were observed at angles "communication" of the LEX fence.

of attack near and beyond maximum lift. The flow visualization photograph reveals the boom wake and Longitudinal and lateral-directional char- the forcbody primary vortex envelope. The boom acteristics. The effect of the LEX fences on the lift, wake consists of multiple, asymmetric vortices shed drag, and pitching-moment characteristics is shown from each step increase in the boom local diameter.

in figures 72, 73, and 74 for Mach numbers of 0.30_ The boom wake may induce a downwash on the fore- 0.60, and 0.80, respectively. At M_c = 0.30 (fig. 72), body, which could reduce the size and strength of the the fences promote a sligllt increase in the lift at an- forebody primary vortices. The boom may also re- gles of attack of approximately 20 ° to 27.5 ° and a duce, or eliminate, the laminar separation region near drag decrease at the corresponding lift coefficients.

the nose. The more turbulent nature of the bound- The favorable interference of the two corotating vor- ary layer along the forebody would also be consistent tices with the fences promotes the lift and drag im- with the smaller and weaker vortices.

provements. At (_ = 30 ° and greater, the slight lift The influence of the asymmetric boom wake is decrease and drag increase are due to tim upward dis- also manifested in the LEX vortex structure. With placement of the burst vortices, which was discussed the nose boom installed, the LEX vortices exhibit in the previous section. This reduces the vortex- asymmetric breakdown at zero sideslip at angles of induced suction pressures oil the LEX's and wings.

attack from approximately 30 ° to 40 °. Reversals of The effects of the fence on the lift and drag the asymmetry occur as the angle of attack increases characteristics at the ifigh angles of attack diminish through this range. Typical results from the va- with increasing values of the Mach number (figs. 73 por screen flow visualizations are shown in figure 77, and 74). At M_ = 0.80 (fig. 74), for example, there which illustrate the asymmetric LEX vortex break- is no change in the lift and drag at angles of attack down of opposite sense at a = 30 ° and 32.5 ° and greater than about 25 ° .

/l.Ioc = 0.60. These trends are indicative of a switch- ing of the asymmetric boom wake as the angle of The fences promote nose-down pitching-moment attack increases. In contrast, the model without the increments up to maximum lift at Al_c = 0.30 boom exhibits symmetric LEX vortex bursting at the (fig. 72). The reduced suction pressures on the same angles of attack.

LEX's, which act ahead of the moment reference cen- ter, contribute to the nose-down pitching moments.

The model with LEX fences may be more suscep- The downward displacement of the LEX vortex to- tible to the effects of the boom wake, since the LEX ward the wing surface, resulting from its interaction primary vortex is displaced upward toward the wake with the second corotating vortex, may also promote generated from the forward fuselage, and is weakened higher suction pressures along the rear portion of the as a result of its reduced generating length. During wing. The increased loading aft of the moment ref- the LEX fence flight validation program conducted erence center is a probable source of the larger nose- by McDonnell Douglas and the Navy (ref. 3), the down pitching-moment increments at lift coefficients nose boom degraded the handling qualities of the up to approximately 1.2. The effect of the fences F/A-18 aircraft at angles of atta& near maximum on the pitching moment diminishes at M_o = 0.60 lift. Removal of the nose boom eliminated the de- (fig. 73) and is negligible at M_ = 0.80 (fig. 74).

graded handling qualities.

The variations of the rolling-moment, yawing- This result is a further example of the sensitiv- moment, and side-force coefficients with the sideslip ity of interactive, or coupled, vortices to the flow development nearthe nose. This is of importance visualizations agreed well with similar measurements to futurefighteraircraftthat mayfeatureincreased obtained in flight on the F-18 HARV. These results forebody-wing vortex interactions at high angles of should be interpreted with caution, however, since attack.

the twin vertical tails or leading-edge flaps may mask the Reynolds number effect at angles of attack whcrc Forebody surface static pressure distribu- vortex bursting occurs near the tails.

tions. Figures 78 and 79 present the forebody sur- The LEX vortices were flatter and higher off the face static pressure distributions with and without surface at the transonic speeds. The shear layer tile flight test nose boom at Moc = 0.6 and 0.8 and instabilities that comprised the LEX vortex at the c_ = 50 ° and 40 °, respectively. The presence of the subsonic speeds were also evident at the higher Mach nose boom delavs primary boundary-layer separation numbers. The onset of core breakdown near the twin and reduces the pressure signature of the forebody vertical tails was slightly delayed due to a diminished vortices. This is indicated most clearly in the surface adverse, longitudinal pressure gradient. The vortex pressures at FS 142 (6.42), which show a marked de- bursting phenomenon at the transonic speeds was crease in the primary vortex-induced suction peaks unsteady and exhibited a fore-and-aft movement of with the boom on. This result is consistent with the breakdown position over the wing. This was the reduced size of the forcbody vortices that was due to an interaction with a normal, or rear, shock shown previously in the vapor screen photograph in wave. The locally supersonic flow induced between figure 76.

the LEX vortices reeompressed to subsonic speeds through a shock wave situated above the aft fuselage Longitudinal and lateral-directional char- and between the vertical tails. The flow details acteristics. The nose boom has essentially no effect provided by the vapor screen method deteriorated at on the lift, drag, and pitching-moment characteristics the higher Mach numbers, which may be the result of at subsonic through transonic speeds. Data corre- decreased values of the local relative humidity within sponding to the model with and without the boom the vortical flow regions.

arc shown in figures 80 and 81 at M_o = 0.60 and 0.80, respectively. The forebody vortices were weak in comparison with the LEX vortical flows. For this reason, they The principal effect of the flight test nose boom were not visible with the laser vapor screen technique is manifested in the lateral-directional characteris- until angles of attack near maximum lift and only tics at sideslip angles ranging from approximately at Maeh numbers of 0.60 and greater. These results -4 ° to +4 ° (ref. 9). Figure 82 presents the vari- are in contrast to recent flow visualization results ob- ation of the rolling-moment, yawing-moment, and tained on the F-18 HARV with a sufficient quantity of side-force coefficients with angle of attack at/3 = 0 ° seeding particles injected from the nose region of the and Mm = 0.60. The nose boom promotes asymmet- airplane, which yielded good definition of the fore- ric rolling- and yawing-moment coefficients that re- body vortices at Moc = 0.2 to 0.3 and o_ = 30 ° to 50 °.

peatedly change sign as the angle of attack increases.

The forebody vortices visualized in the wind tunnel and in flight were rapidly entrained into the domi- Summarization of Results nant LEX vortical flows. At the transonic speeds, The laser vapor screen flow surveys revealed a the forebody vortex size and strength increased due complex vortex structure generated from tile LEX's to cross-flow shock-induced boundary-layer separa- of the baseline 0.06-scale F/A-18 model. Multiple, tion. In contrast, the LEX vortex strength dimin- smaller scale vortices were generated along the length ished at the higher Mach numbers. This allowed the of the wing leading-edge extension. These shear layer forebody vortices to persist farther into the LEX flow instabilities rolled up to form the central, dominant field prior to their entrainment.

vortex that is typically observed in water tunnel and The footprints of the LEX vortices were easily wind tunnel testing and in full-scale flight. Evidence discerned in the surface static pressures at the sub- has surfaced recently from smoke flow and condensa- sonic speeds. The progression of vortex bursting over tion patterns on the F-18 airplane that similar vortex the LEX correlated well with the pressure distribu- flow structures exist in flight.

tion trends. In contrast to the vapor screen results, The global characteristics of the F/A-18 LEX vor- the surface pressures revealed a high level of com- tex were insensitive to the Reynolds number at sub- pressibility of the leading-edge vortex, beginning at free-stream Maeh numbers as low as Moc = 0.20 to sonic speeds. The location of LEX vortex breakdown 0.30. The pressure distributions were flatter at the and its progression with the angle of attack that were determined from the wind tunnel vapor screen flow transonic speeds, which made it difficult to identify the LEX vortex footprint or the passage of vortex speeds. At angles of attack beginning at approxi-

breakdown at the higherangles of attack. mately 30 °, the wake from the nose boom promoted

an asymmetry in the LEX vortex breakdown posi-

The forebody vortexsignatures werewell-defined

tions. Reversals of the asymmetry occurred as the

in the mode]pressure distributionsat the subsonic

angle of attack increased toward maximum lift. The

andtransonic speeds at angles of attackof approxi-

LEX vortices in the presence of the fences wcrc sen-

mately30 ° andgreater.Thesurface pressures along

sitive to the asymmetric flow development near the

mostof the forebodywereinsensitive to the Mach

nose caused by the nose boom, due to their reduced

number from M_c = 0.20 to 0.60. At Mcc = 0.80 and

strength and upward displacement toward the boom 0.90, however, primary separation occurred at a lower wake. The ensuing asymmetric rolling moments and position on the body, and the forebody vortex suction yawing moments measured on the wind tunnel model peaks exhibited a significant increase. The cross-flow were consistent with the degraded handling qualities Mach number at angles of attack near maximum lift experienced in the Navy/McDonnell Douglas flight was sufficiently high to promote the development of tests of the F/A-18 with both LEX fences and nose a shock wave along the side of the body, which was boom. Removal of the nose boom eliminated the ad- strong enough to separate the boundary layer.

verse flow interaction. It is noted that the production F/A-18 radome does not incorporate a nose boom.

A fence mounted to the upper surface of the LEX's in a streamwise orientation and near the wing- The forebody and LEX surface pressures ob- LEX junction altered the development of the primary tained on the baseline 0.06-scale model in the DTRC vortex. The LEX vortex sheared away from the lead- 7- by 10-Foot Transonic Tunnel and the Langley ing edge and moved upward and inboard as it passed 14- by 22-Foot Subsonic Tunnel at the same free- over the fence. The migration of the LEX vortex al- stream Mach number (Mc_ = 0.20) and Reynolds lowed the formation of a second corotating vortical number (Ree = 0.96 x 106) were in reasonable agree- flow (rotating in the same sense) from the leading ment up through poststall angles of attack. The dif- edge. The cross-flow structure over the wings and ferences in the tunnel blockage, free-stream turbu- near the vertical tails was significantly changed as lence, model support system, and the proximity of a result of an interaction of the corotating vortices.

the model nose to the DTRC ceiling at the extreme The flow induced by the vortices on each other was angles of attack were not sufficient to affect the data in opposition, which reduced the mean flow veloci- repeatability at the low subsonic speeds. Comparison ties and flow angularity at the vertical tails. This of the 0.06-scale model forebody and LEX pressures effect was observed within the angle-of-attack range with data obtained on a 0.16-scale model tested in (ct = 25 ° to 30 °) where the tail excitation due to LEX the Langley 14- by 22-foot facility at Ree = 0.96 x 106 vortex bursting was greatest on the baseline F/A-18.

and -_foc = 0.08 revealed large differences. The pri- The fences did not adversely affect the LEX vortex mary separation occurred earlier on the forebody of breakdown characteristics. The core breakdown was the larger scale model, with a consequent increase more gradual and the level of turbulence within the in the vortex strengths. The signatures of the LEX burst vortical flow was less with the fences installed.

vortices were also more pronounced on the 0.16-scale The maximum lift was reduced slightly as a result of F/A-18 model. The differences in the strength of a vortex displacement effect. Otherwise, the fences the forebody vortices and their subsequent interac- were benign and had minimal impact on the config- tion with the LEX vortical flows may account for the uration aerodynamic and stability characteristics at different lateral stability characteristics that the two the subsonic through transonic speeds.

models have consistently exhibited at angles of attack ncar maximum lift.

The effectiveness of the fences in improving the vertical tail buffet environment has been quantified The interpretation of flow visualization results by McDonnell Douglas, the U.S. Navy, and NASA obtained with different illumination and/or parti- in full-scale flight tests. Recent flow visualization cle seeding techniques requires care. The vortical experiments conducted by NASA on the F-18 HARV flows on the 0.06-scale model were illuminated by with LEX fences have shown effects of the fence on the wind tunnel test section lights, which provided the vortex location and structure similar to those a three-dimensional perspective of the condensation observed in the wind tunnel.

patterns, and by the laser light sheet, which yielded The flight test nose boom generated a multiple, flow-field cross sections. On the baseline F/A-18 asymmetric vortex wake at the high angles of attack. model, the two illumination techniques provided sim- ilar information. For example, the shear layer insta- The presence of the boom reduced the strength of the bilities and vortex breakdown that were visible with forebody primary vortices at subsonic and transonic the tunnel lightingonly werealsoseen,in greater The LEX vortices are highly compressible, even detail,with the light sheet. The resultsobtained at thc very low subsonic Math numbers. The core breakdown location is insensitive to the Mach num-

with the twolightingschemes compared welland,in

ber, however, until shock waves appear over the

a straightforward fashion,with the in-flight smoke

flowvisualizations obtained on the F-18HARV.

wings that interact with the vortex flows.

Independent analysisof the flow patternsob-

The F/A-18 forebody vortices are comparatively

tainedwith the tunnel lightingand the laserlight

weak and arc dominated by the LEX vortical flows

shceton the modelwith the LEX fences couldlead

at all Mach m_mbcrs. Compressibility effects are

to conflictingconclusions, however. The vortex

not manifested to a significant degree on the fore-

flows,madevisibleby the tunnel lights, expanded

body until the transonic speeds, where shock-induced

downstream of the LEX fences andexhibitedwhat

boundary-layer separation promotes larger and

appeared to be a classical vortex breakdown phe-

stronger vortices.

nomenon. Thesecond corotating vortexwasnot vis-

The 0.06-scale model exhibits high levels of lateral

ible. The laserlight sheetrevealed two stablevor-

stability up to maximum lift at subsonic speeds. A

ticesthat interacted with each otherto formaunique

reduction in lateral stability is apparent at the tran-

cross-flow structuredownstream of the LEX fencc

sonic speeds, however, due to a vortex-shock inter-

with the vortexcoresdefined. This cross flowcould

action. The high-angle-of-attack directional stability

easily have been misinterpreted asanexpanded burst

is nonlinear due to the interaction of the burst LEX

vortexwithoutthe aidof the laser-illuminated cross

vortices with the vertical tails.

flows.Thesmoke injectionmethod employed on the

F-18HARVwith LEX fences yicldcd vortex patterns

Thc LEX fence effectively reduces the vertical tail similar to the wind tunnel observations with the test excitation at high angles of attack by restructuring section lights only. A more detailed description of the the LEX vortex. The fence promotes a system of in-flight flow field was possible when analyzed along two weaker vortices from the LEX, and their mutual with thc laser light-sheet results from the ground- interaction reduces tile mean flow velocities and flow based facility.

angularity at the tails.

Tile injection of water into the tunnel circuit The LEX fence does not adversely affect the vor- in sufficient quantity to conduct the laser vapor tex breakdow_n behavior and has a minimal impact on screen flow visualizations did not affect tile forebody the aerodynamic and stability characteristics. Flight and LEX surface pressures or the total forces and experiments of the F-18 airplane corroborate the moments. This result is of importance to high-angle- wind tunnel observations.

of-attack testing, since it supports the simultaneous A flight test nose boom alters tile development of acquisition of the vapor screen off-body flow-field the forebody vortices at the subsonic and transonic information and quantitative model measurements.

speeds. Transition to turbulent flow may occur Concluding Remarks earlier on the forebody, which promotes smaller and weaker vortices.

A wind tunnel experiment was conducted in the David Taylor Research Center (DTRC) 7- by 10-Foot The F-18 with LEX fences is susceptible to LEX Transonic Tunnel to improve the understanding and vortex burst asymmetries at zero sideslip with the control of the vortical flows about a 0.06-scale model nose boom installed. Wind tunnel and flight experi- of the F/A-18 at high angles of attack and at subsonic ments show that the adverse forebody LEX flow in- through transonic speeds. Laser vapor screen flow teraction and the resultant handling qualities degra- visualizations, model surface static pressures, and dation are eliminated when the boom is removed.

six-component forces and moments were obtained The analysis of the off-body flow visualizations at angles of attack from 10 ° to 50 °, free-stream obtained in the wind tunnel and in flight requires Maeh numbers of 0.20 to 0.90, and Reynolds numbers care. The traditional interpretation of vortex core based on the wing mean aerodynamic chord of 0.96 x breakdown as a rapid expansion of the vortical region 106 to 1.75 x 106 . The model was tcsted in a does not apply to tile F-18 with LEX fences. The baseline configuration and with wing-leading-edge- diffused nature of the interacting vortices near the extension (LEX) upper surface fences and flight test fence resembles a classical breakdown of a single nose boom.

vortex system.

The high Reynolds number behavior of the LEX A model scale effect exists in wind tunnel testing vortex flows on the F/A-18 aircraft can be simulated of the F/A-18 even when the models are tested in at lower Reynolds numbers in the wind tunnel.

doesnot adversely affect the quantitativemodel

the samefacility at the samenominalvalueof the

measurements.

Reynolds number.The boundary-layer development

on the forebodymay be sensitive to the surface

The detailsthat areextractedfromthe off-body

finish,trip arrangements, andscale ofthefree-stream

flow visualizations are dependent on the illumina-

turbulence. These effects couldpromote significantly

tion and particleseeding techniques employed.A

differentforebody and LEX vortexsignatures and

combination of flowvisualization methods canhelp

corresponding differences in the high-angle-of-attack

to ensure the correctinterpretationof the vortex-

stabilitycharacteristics.

dominated flowfields.

Injectionof waterinto the tunnelcircuit in suf-

NASA Langley Research Center

ficient quantity to promotelocal condensation in

Hampton, VA23665-5225

the vortex flow regionsabout the F/A-18 model September 30,1991

References Volume L Technical Discussion and Analysis of Results.

AFFDL-TR-78-94, Vol. I. (Available from DTIC as AD 1. Fisher, David F.; Richwine, David M.; and Banks, A069 646.)

Daniel W.: Surface Flow Visualization of Separated Flows on the Forebody of an F-18 Aircraft and Wind-Tunnel Volume IL Data Base. AFFDL-TR-78-94, Vo]. II. (Avail- Model. NASA TM-100436, 1988.

able from DTIC as AD A069 647.)

2. Hall, Robert M.: Influence of Reynolds Number on Fore- 12. Del Frate, John H.; and Zuniga, Fanny A.: In-Flight body Side Forces for 3.5-Diameter Tangent-Ogive Bod- Flow Field Analysis on the NASA F-18 High Alpha ies. A Collection of Technical Papers AIAA 5th Applied Research Vehicle With Comparisons to Ground Facility Aerodynamics Confcrvnce, Aug. 1987, pp. 63 73. (Avail- Data. AIAA-90-0231, Jan. 1990.

able as AIAA-87-2274.)

13. Fisher, David F.; and Meyer, Robcrt R., Jr.: Flow 3. Frazier, F. Alan: F/A-18 Hornet LEX Fence Flight Test Visualization Techniques for Flight Research. NASA Results. 1988 Report to the Aerospace Profession - TM-100455, 1988.

Thirty-Second Symposium Proceedings, Soc. of Experi- mental Test Pilots, 1988, pp. 72 89.

14. Payne_ F. M.; Ng, T. T.; Nelson, R. C.; and 4. Schneider, Edward T.; and Meyer, Robert R., Jr.: F-18 Schiff, L. B.: Visualization and Flow Surveys of the Lead- tIigh Alpha Research Vehicle Description, Results, and ing Edge Vortex Structure on Delta Wing Planforms.

Plans. I989 Report to the Aerospace Profession Thirty- AIAA-86-0330, Jan. 1986.

Third Symposium Proceedings, Soc. of Experimental Test 15. Squire, L. C.; Jones, J. G.; and Stanbrook, A.: An Pilots, 1989, pp. 135 162.

Experimental Investigation of the Characteristics of Some 5. McGregor, I.: The Vapour-Screen Method of Flow Vi- Plane and Cambered 65 _ Delta Wings at Mach Numbers sualization. J. Fluid Mech, vol. 11, pt. 4, Dec. 1961, From 0.7 to 2.0. R. & M. No. 3305, British Aeronautical pp. 481 511.

Research Council, 1963.

6. Erickson, Gary E.; Rogers, Lawrence W.; Schreincr, 16. Campbell, James F.; Chambers, Joseph R.; and Rumsey, John A.; and Lee, David G.: Subsonic and Transonic Vor- Christopher L.: Observation of Airplane Flow Fields by tex Aerodynamics of a Generic Forebody Strake-Cropped Natural Condensation Effects. AIAA-88-0191, Jan. 1988.

Delta Wing Fighter. AIAA-88-2596, Jm_e 1988.

17. Ghaffari, Farhad; Luckring, James M.; Thomas, James L.; 7. Erickson, Gary E.; Rogers, Lawrence W.; Schreiner, and Bates, Brent L.: Navier-Stokes Solutions About the John A.; and Lee, David G.: Further Studies of the F/A-18 Forebody-LEX Configuration. AIAA-89-0338, Subsonic and Transonic Vortex Flow Aerodynamics of a Jan. 1989.

Close-Coupled Forebody-Slender Wing Fighter. AIAA- 88-4369, Aug. 1988.

18. Kiichemann, D.: The Aerodynamic Design of Aircraft.

8. ASED Staff: Transonic Wind- Tunnel Facility at the Naval Pergamon Press Inc., c.1978.

Ship Research and Development Center. ASED Rep. 332, 19. Fisher, David F.; Banks, Daniel W.; and Richwine, David W. Taylor Naval Ship Research and Development David M.: F-18 High Alpha Research Vehicle Surface Center, June 1975. (Available from DTIC as AD A014 Pressures: Initial In-Flight Results and Correlation With 927.)

Flow Visualization and Wind-'lhmnel Data. A Collection 9. Banks, Daniel W.: Wind-Tunnel Investigation of the Fore- of Technical Papers, Part I AIAA 8th Applied Aerody- body Aerodynamics of a Vortex-Lift Fighter Configura- namics Conference, Aug. 1990, pp. 421 451. (Available tion at High Angles of Attack. Advanced Aerospace Aero- as AIAA-90-3018-CP.)

dynamics, SP-757, Soc. of Automotive Engineers, Inc., Oct. 1988, pp. 101 123. (Available as SAE 881419.)

20. Lorincz, Dale J.: A Water Tunnel Flow Visualization Study of the F-t5. NASA CR-144878, 1978.

10. Erickson, Gary E.: Water Tunnel Flow Visualization and Wind Tunnel Data Analysis of the F/A-18. NASA 21. Erickson, Gary E.; Schreiner, John A.; and Rogers, CR-165859, 1982.

Lawrence W.: On the Structure, Interaction, and Break- 11. Headley, Jack W.: Analysis of Wind Tunnel Data Per- down Characteristics of Slender Wing Vortices at Sub- taining to High Angle of Attack Aerodynamics. U.S. Air sonic, Transonic, and Supersonic Speeds. AIAA-89-3345, Force, July 1978. Aug. 1989.

2O ORIGINAL PAGE BLACK AND WHITE PiiOTO(_l-tAWt-_ @ F-18 High-Alpha Research Vehicle.

Figure 1. NASA ORIGINAL PAGE BLACK AND WHITE PHOIO_,RAI.-'h I.EX fence--_ ......................... .......... .... -- (a) Top view of F-18.

(b) Close-up of LEX.

(c) Carrier landing of F/A-18.

L-91-66 Figure 2. Navy fleet airplanes with LEX fences.

off (15 Hz) 10 m _----Fence on (15 Hz) Vertical stabilizer 1 acceleration, g2/Hz s s .1 s

I I I

.01 20 30 40 tx, deg Effect of LEX fence on vertical tail buffet (from ref. 4).

Figure 3.

Reference dimensions

S= 400ft2 (1.440 ft2)

b = 37.417 ft (2.245f-t)

_= 11.517 ft (0.691 ft)

c.g.= 25%E

Figure4. F/A-18geometry details.Dimensions arein feetfull scale (0.06scale).

FS 296 (17.76) FS 357 (21.42) FS 253 (15.18) FS 184 (11.04) LEX fence FS 142 (8.52) FS 107 (6.42) Figure 5. Wind tunnel model forebody and LEX pressure measurement stations. Dimensions are in inches full scale (0.06 scale).

ORIGINAL PAGE dLACK AND WHITE P_OIO(_R_H -1 Pressure- JlIStFUm ented forebody L-91-67 Figure 6. Forebody and LEX pressure measurement stations on F-18 HARV. Dimensions are in inches full sea]e,

FS378.17 (22.69)

LEX fence

BL /¢.,,, \ _- BL 42.00

, (0.00) -/

0.00 _65o (2.52)

-'"LEX fence (nominal) 8.30 22.58 (0._0) ---_ [------(1.36) 60 ° Fence LE Mounting _"_ F-$237689}7 r- Fence TE pad (2) L--"/-_ 32.18 I"--- _ t • )-] / FS410_33 FS 60.60 (1.93) _z4.oz)

A (3.63) _ _-II A

FS 404.05 (24.24) // LEX- wing junction /

(LE flap un_nected) --/

Figure 7. Details of LEX fences on forward fuselage component. Dimensions are in inches full scale (0.06 scale).

l ti°nA A°f gu

68.62 (4.12) Figure 8. Flight test nose boom geometry details. Dimensions are in inches full scale (0.06 scale).

Roll stingassembly

Sting

adapter

Main

•. support

system

boom Strain-gauge

balance

2-

Mainsu

boom

Roll sting

34 20

22----_ -'-- 3.6

93.6

i

i

l E E E Figure9. DTRCroll stingarrangement used for high-angle-of-attack testing.Dimensions arein inches.

T 39° --_ Video camera location (three-quarter, right rear) Figure 10. Model movement through test section at selected angles of attack.

-, '_' PAqE OR,G_ ,At.

• ," pHOTC_R_PH BLACK AND WH, i E Figure 11. Sting-mounted 0.06-scale F/A-18 model in the DTRC 7- by 10-Foot Transonic RXmnel.

FS 107 (6.42) 180 ° | = | FS 253 (15.18) 270 ° 90 ° -y --.-_ y u !

=- o i

..,s,-]

|

FS 142 (8.52) , ] 180 ° i == FS 296 (17.76) E w !

lie 90 ° 270 ° r -y --_y _R ___ m o B FS 184 (11.04) 180 ° FS 357 (21.42) -y y 90 ° 270 ° m_ E O (a) Forebody angular position. (b) LEX spanwise position.

Figure 12. Forebody and LEX surface static pressure orifice orientations. Dimensions are in inches full scale (0.06 scale).

o No Water Injection [] Water Injection FS 107 (6.42) 180 ° -2.0 -1.5 90 ° 270 ° -1.0 0 o Cp -.5 .5_ I I I I I I I I I I I I 1.0 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg -2.0 -1.5 -1.0 Cp -.5 .5 I I I__1 I I I I I I I I 1.0 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg (a) FS 107 (6.42); a = 30 °.

= 0.90 and Ree = 1.02 x 106 .

Figure 13. Forebody surface pressures with and without water injection at Moc Dimensions are in inches full scale (0.06 scale).

= o No Water Injection [] Water Injection FS 184 (11.04) -2.0 180 °

-1.5f

270 ° 90 ° Cp _.

-1.0! __0_ ° 1.01 I ...] I I I I I I I I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 i O, deg -2.0 - -1.5 - -1.0 - Cp -.5 - .5 1.0 I 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg (b) FS 184 (11.04); c_ = 40 °.

Figure 13. Concluded.

o No Water Injection [] Water Injection -3.0 FS 253 (15.18) -2.5 -2.0 Cp,u -1.5 -1.0 -.5 I I I I I I 0 I I I I 0 .2 .4 .6 .8 1.0 -.8 -.6 -.4 -.2 -1.0 y/s -3.0 - FS 296 (17.76) -2.5 - -2.0 - Cp,u -1.5 - -1.0 - -°5 - I I I I I 0 1 I I I I .2 .4 .6 .8 1.0 -1.0 -.8 -.6 -.4 -.2 0 y/s -3.0 - -2.5 - FS 357 (21.42) -2.0 - Cp,u -1.5 -1.0 -.5 I I I I I I 0 I I I I 0 .2 .4 .6 .8 1.0 -°8 -°6 -°4 -.2 -1.0 y/s Figure 14. LEX surface pressures with and without water injection at Mec = 0.90, Ree -- 1.02 × 106, and a = 30 °. Dimensions are in inches full scale (0.06 scale).

o No Water Injection = [] Water Injection FS 107 (6.42) 180 ° -2.0 -1.5 -1.0

i

T

q, -.5 .5 1.0!

R

I I J I I I i J I I I I 30 60 90 120 150 180 210 240 270 300 330 360 i

|

O, deg -2.0 - -1.5 -1.0 Cp -.5 E f .5 L 1.0 _ F E I t I I I I I I I I I t 0 F 30 60 90 120 150 180 210 240 270 300 330 360 O, deg (a) FS 107 (6.42); a = 40 °.

Figure 15. Forebody surface pressures with and without water injection at -_Ioc = 0.60 and Ree = 1.32 x 106.

Dimensions are in inches full scale (0.06 scale).

o No Water Injection [] Water Injection FS 184 (11.04) -2.0 - 180° - 1.0 - 90o "270° -1.5 Cp 0 ° .5 1.0 0 30 60 90 120 150 180 210 240 270 300 330 360 O, deg -2.0 -1.5

E

-1.0 Cp -,5 .5 1.01 I I I I 1 t I I I I I I 30 60 90 120 150 180 210 240 270 300 330 360 O, deg (b) FS 184 (11.04); a = 50 °.

Figure 15. Concluded.

o No Water Injection O Water Injection -3.0 - -2.5 - FS 253 (15.18) -2. 1_;O _ -1.0 -,5 CP, u -1.5 -1.0 -.8 -.6 -.4 -.2 0 .....

y/s -3.0 - i 0 I I _ i 6 8 110 FS 296 (17.76) -2.5 - i -1.0 - ._ -2.0 - -.5- i 0_...----1 I, I I I I A I I • -1.0 -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 y/s -3.0 FS 357 (21.42) -2.5 - -2.0 - CP, u -1.5 -1.0 -,5 I .1 I i I J I I _l I 0 .2 .4 .6 .8 1.0 -.8 -.6 -.4 -.2 -13 y/s Figure 16. LEX surface pressures with and without water injection at M_ = 0.60, Ree = 1.32 x 106, and a = 40 °. Dimensions are in inches full scale (0.06 scale).

M_ Tunnel Configuration O 0.20 DTRC 7 x 10 0.06-scale F/A-18 [] 0.20 LaRC 14 x 22 0.06-scale F/A-18 -2.0 -1.5 FS 107 (6.42)

F

180° -1.0 _- 90 ° Cp -.5 270 ° 0° .5 1.0 I t I I I t I I I I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg -2.0 FS 142 (8.52) 180° -1.5 • -1.0 Cp -.5 900@ 270° OO .5 1.0 I I I I I I I I I I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg -2.0 FS 184 (11.04) -1.5 180° -1.0 Cp -.5 0 90°_ 270° .5 0o I I I I I 1 I I I I I I 1.0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg Figure 17. Forebody surface static pressures on 0.06-scale F/A-18 model at M_ = 0.20, Ree = 0.96 x 106, and a = 40 °. Dimensions are in inches full scale (0.06 scale).

Moo Tunnel Configuration O 0.20 DTRC 7 x 10 0.06-scale F/A-18 [] 0.20 LaRC 14 x 22 0.06-scale F/A-18 -2.0 - = -1.5 - _ ..fl__ FS 107 (6.42) i

#,_--%_ _ _ ¢_-_, ,_oo

-1.0 -

/ 70o

0-

.5 - 0° 1.0 Cr* I l I I I 1 I I I ! I I

o 30 60 90 120 150 180 210 240 270 300 330 360

!

-2.0 _ 0, deg !

,_ _,_o(_,_

Cp -.5 o 90 ° 270 °

_10 @ |

.5 0o m 1.0 I [ 1 I I 1 I I 1 I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg FS 184 (11.04) -1.5 180 ° -2.0 ; -1.0 - Cp _.5 90°_ 270° .5 0o I I I t 1 I I I I I I I 1.0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg Figure 18. Forebody surface static pressures on 0.06-scale F/A-18 model at Moo = 0.20, R% = 0.96 x 106, and a = 50 °. Dimensions are in inches full scale (0.06 scale).

M_o Tunnel Configuralion O 0.20 DTRC 7 x 10 0.06-scale F/A- 18 [] 0.20 LaRC 14 x 22 0.06-scale F/A- 18 B -3.5 -3.0 FS 253 (15.18) -2.5 -2.0 Cp,u -1.5 -1.0 - -.5 0 I I I I I I 1 I I 1 -1.0 -.8 -o6 _°4 -.2 0 .2 .4 .6 .8 1.0

y/s

-3.0 - FS 296 (17.76) Cp,u I I I I I I 0 .2 .4 .6 .8 1.0

y/s

FS 357 (21.42) ,U

@

I I I I I I 0 .2 .4 .6 .8 1.0 y/s Figure 19. LEX surface pressures on 0.06-scale F/A-18 model at 3Ice = 0.20, Ree = 0.96 x 106, and cr = 40% Dimensions arc in inches full scale (0.06 scale).

Moo Tunnel Configuration

O 0.20 DTRC7 x 10 0.06-scaleF/A-18

[] 0.08 LaRC 14x 22 0.16-scale F/A-18

-2.o _ : -1.5 FS 107 (6.42) --

-_ _ /__ _ o _oo

cp ___ _'__/_'__ 90 ,_ ,7o. i

.51 pC 0° 1.0 I I I I I I I I._ I I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 i 0, deg _i -2.0 i

-_._ _ __ _ _ %(_,._

Cp -" 90 ° 270 1.0 I I I I I I I I I I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg -2.0 E FS 184 (11.04) _-m -1.5 180 ° -1.0 Cp -.5 90 ° 270 ° .5 o 1.0 I I I I __l I I I I I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg Figure 20. Forebody surface pressures on 0.06- and 0.16-scale F/A-18 models at a = 40 ° and Rec = 0.96 x 106.

Dimensions are in inches full scale (0.06 scale).

Moo Tunnel Configuration

o 0.20 DTRC7× 10 0.06-scaleF/A-18

0.08 LaRC 14×22 0.16-scale F/A-18

[]

-2.0 - -1.5 - FS 107 (6.42) 180° -1.0 - Cp -,5 90° _ 270° 0° .5 I I I I I i I I I I I I 1.0 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg -2.0 - FS 142 (8.52) -1.5 - 180° -1.0 Cp -.5 90o_270o .5 0o 1.Ot I I I I I I I I I I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg -2.0 - FS 184 (11.04) 180° -1.5 - -1.0

Cp -.5

90°_ 270° 0 ° .5 1.0 I I I 1 I I 1 I I I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg Figure 21. Forebody surface pressures on 0.06- and 0.16-scale F/A-18 models at c_ = 50 ° and Ree = 0.96 × 106.

Dimensions are in inches full scale (0.06 scale).

Moo o 0.20 [] 0.40 i <> 0.60 i A 0.80 ) FS 107 (6.42) i 180 ° -I.0 - _ = % -.5 90 ° 270 ° 0 o .5 1.0 I l 1 I I I I I I 1 1 I 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg -2.0 V -1.5 L FS 142 (8.52)180 ° !

cp -.5 90 ° 270 °' .5 0o i z I.O I I I I I I I I I I I I m 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg ira: FS 184 (11.04) B -1.5 180 ° m -2.0 f -1.0 _ C..p _.50 i ------ _ 90° 270° .5! 0o 1.0 I l l I I I I I t _ I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 L 0, deg Figure 22. Effect of Mach number on 0.06-scale F/A-18 model forebody surface pressures at c_ = 40 °.

Dimensions are in inches full scale (0.06 scale).

Configuration Moo Tunnel 0.06-scale F/A- 18 O 0.20 DTRC 7 x 10 0.16-scale F/A-18 [] 0.08 LaRC 14 x 22

-4.5 -

-4.0

-3.5

FS 253 (15.18)

-3.0

-2.5

Cp,u

-2.0 -

-1,5 -

-1.0 -

-.5 - 0 I 1 I I I I I I I I -1.0 -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 y/s -3.0 FS 296 (17.76) -2.5 -2.0 Cp,u -1.5 -.5 -1.0 t I I I I I I I I I 1 -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 -1.0 y/s FS 357 (21.42) -2.0 -1.5 Cp,u -1.0 -.5 0 I I I I 1 I I I I I -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 -1.0

y/s

Figure 23. LEX surface pressures on 0.06- and 0.16-scale F/A-18 models at a = 40 ° and Ree = 0.96 x 106.

Dimensions are in inches full scale (0.06 scale).

ORrGINAL PA_E BLACK AND WHITE PHOTO_RAp_ (a) _ = 20°; FS 525 (31.51).

i E_ I !

v F (b) c_ -- 20°; FS 567 (34.02).

Figure 24. Laser vapor screen flow visualizations on baseline F/A-18 model at Mcc = 0.40 and Ree = 1.75 x 106.

Camera is in three-quarter, right rear position; dimensions are in inches full scale (0.06 scale).

44 z BLACK Ai_D WHITE F_i_iOiCGRAPH (c) a = 25°; FS 411 (24.67}.

(d) a = 25°; FS 450 (31.51).

Figure 24. Continued.

ORIGINAL PAGE 8LACK AND WH,,E PHOTOGRAPH

|

l

[

i

m (e) c_ = 30°; FS 357 (21.42).

i E E m m (f) c_ = 30°; FS 411 (24.67).

Figure 24. Continued.

BLACK AND WHITE PHOTOQRAPN (g) c_ = 35°; FS 296 (17.76).

(h) a -- 35°; FS 357 (21.42).

Figure 24. Concluded.

ORIGINAL p::,:__ BLACK AND WHITE PHO-i'O_ffAF_-: (a) c_ = 19.65°; fl = 0.35 °.

(b) c_ = 25.28°; fl = 0.65 °.

Figure 25. In-flight smoke flow visualizations of LEX vortcx breakdown on F-18 HARV at ,¢'Izc _ 0.3 and Rc_ _ 13.5 × l06 (ref. 12).

BLACK AHD WHITE PHOTOGRAPH (c) o_ = 29.85°;/_ = 0.25°.

i

(d) _ = 33.50°; ,3 = 0.60°.

Figure 25. Concluded.

© Flight, smoke [] Flight, natural condensation ii = Wind tunnel, vapor screen = = I i FS 404 (24.24) (LEX-wing-LE junction) i __1 FS 434 (26.04) (LEX-wing-LE-flap-hingeline junction) = _= FS 525 (31.50) (Vertical tail LE-fuselage junction) 35 (15.18) l FS253 l_ N FS 296 i (17.76) deg 25 FS 357 (21.42) T_ = = E !

= l 1 I I i I I I I I I 0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 x/l Eigure 26. LEX vortex breakdown progression with angle of attack on baseline 0.06-scale F/A-18 model (M_c = 0.40, Rec = 1.75 × 106 ) and F-18 HARV (M_: _ 0.3; Ree _ 13.5 × 106). (Flight results are from ref. 13.)

OI_IG!r'AL PAGE BLACK AND WHITE PHOI-OGRAPH (a) FS 411 (24.66).

(b) FS 450 (27.00).

Figure 27. Laser vapor screen flow visualizations on baseline F/A-18 model at M_c = 0.60, Re_ = 1.32 x 106, and a = 20 °. Dimensions are in inches full scale (0.06 scale).

ORIG R_IAL PAGE BLACK AND WHITE PHOTOG_RAPh (c) FS 483 (28.98).

(d) FS 525 (31.50) (just prior to breakdown).

Figure 27. Concluded.

• j # _; ,-- OI'_IG!r'_AL F -,_L: BLACK AriD WHJTE P;"IOi'OQRAPH Light-sheet tion (a) a = 15 °.

Light-sheet tion (b) a=17.5 ° • Figure 28. Effect of angle of attack on cross flow about baseline F/A-18 model at Moo = 0.60, Rez. = 1.32 x 106, and FS 357 (21.42). Dimensions are in inches full scale (0.06 scale).

ORIGINAL PAGE BLACK AND WHITE PHO-i-OGRAPI.4 /_ Light-sheet tion

i

(c) a = 20 °.

Z

m

i m !

B Light-sheet II

_N

tion (d) _ = 25 °.

Figure 28. Concluded.

ORTG fr,_,_'_,L PAGE BLACK AND WHITE Pi-_©FOGRAPH Light-sheet location (a) a= 15 °.

Light-sheet . lQcation (b) (_ = 17.5 °.

Figure 29. Effect of angle of attack on cross flow about baseline F/A-18 model at M_ = 0.60, Re_ = 1.32 × 106, and FS 411 (24.66). Dimensions are in inches full scale (0.06 scale).

ORIGINAL PAGe.

8LAC.,K AND WHITE PHOi-OGRAPr, Light-sheet location (c) a = 20 °.

Light-sheet location (d) c_ = 25 °.

Figurc 29. Continued.

ORIGINAL P,r-',G E BLACK AND WHITE PHOTOGRAPH (e) c_ = 30 °.

Figure 29. Concluded.

BLACK AND WH!TE Pi--IOi-O_A_H Light-sheet (a) _ = 15 _.

Light-sheet L._catia°_7 (b) o_ = 17.5% Figure 30. Effect of angle of attack on cross flow about baseline F/A-18 model at 2t1_ = 0.60, Re_ = 1.32 × 106, and FS 450 (27.00). Dimensions arc in inches full scale (0.06 scale).

BLACK AND WH!TE _'rIuI_RAPH (c) a = 20 °.

Light-sheet (d) a = 25 °.

Figure 30. Continued.

P _ q,,,..! 15.

OR!GINAL i ," ___ 8LACK AND WH!TE PHO-I'OGP_PH m

i

E (e) a = 30 °.

Figure 30. Concluded.

6O _' ^ .... 3 .- BLACK AND WHITE PHOTC'3RAPH Light-sheet

on/

(a) a = 15 °.

Light-sheet (b) c_ = 17.5 °.

Figure 31. Effect of angle of attack on cross flow about baseline F/A-18 model at Mcc -- 0.60, Ree = 1.32 × 106, and FS 483 (28.98). Dimensions are in inches full scale (0.06 scale).

Light-sheet ORiGir,',_,.,,,. P ::2 BLACK AND WH',TE F'i_UTOGRAPH Light-sheet (a) o_ = 15°.

ig ht'sheet (b) a=17.5 ° .

Figure 32. Effect of angle of attack on cross flow about baseline F/A-18 model at M_ = 0.60, Ree = 1.32 x 106, and FS 525 (31.50). Dimensions are in inches full scale (0.06 scale).

ORIGINAL p._'_ _ 8t-ACK AND WHITE PHO1"OGRAP_

Ll_ight-sheet

(c) c, = 20 °.

m L

_Light-sheet

(d) ct = 25 °.

Figure 32. Concluded.

BLACK Ai_D WHITE I-t_iC, GRAPH _Light-sheet (a) _ = 15 °.

Light-sheet (b) c_ = 17.5 °.

Figure 33. Effect of angle of attack on cross flow about baseline F/A-18 model at Moc = 0.60, Ree = 1.32 x 106, and FS 567 (34.02). Dimensions are in inches full scale (0.06 scale).

ORtG_NAL PAGE AND WHITE PHOTOI2RAPg Light-sheet z (c) a = 20 °.

Light-sheet E .m E .e= (d) a = 25 °.

Figure 33. Continued.

ORIGINAL P_;_E BLACK AND WHITE i'r_otOORAPFI (e) a = 30 °.

Light-sheet (f) a = 35 °.

Figure 33. Concluded.

|

i

m B B Figure 34. In-flight flow visualization (natural condensation) of the F/A-18 LEX vortices (ref. 16).

M_ © 0.20 [] 0.40 5O © 0.60 FS 404 (24.24) (LEX-wing-LE junction) 4O FS 434 (26.04) (LEX-wing-LE-flap-hingeline junction) FS 257__, 35 FS 525 (31.50) (Vertical tail LE-fuselage junction) (15"18) I "_, X (17.76)] 1 C_, deg FS 357 (21.42) , l I I I 1 I I I t 1 I 0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 x/l Figure 35. Effect of Mach number on LEX vortex breakdown characterizations of baseline F/A-18 model.

Dimensions are in inches full scale (0.06 scale).

ORIGINAL p,_._E BLACK AND WHITE PHoI'OGiRAPi4 Light-sheet (b) M_ = 0.80, a = 15 ° .

Figure 36. Cross-flow patterns about baseline F/A-18 model at Moc = 0.60 (Ree = 1.32 × 106) and Mac = 0.80 (Ree = 1.02 × 106) at FS 525 (31.50). Dimensions are in inches full scale (0.06 scale).

ORIG!F,t,AL i:'2 :: BLACK AND 3,t,"HITE f:b_b]-OGRAPH Light-sheet (c) Moc = 0.60, o_ = 17.5°.

I Light-sheet

on!

(d) Moc = 0.80, a = 17.5 °.

Figure 36. Continued.

ORIGINAL PAGE BLACK AND WHITE PHOTOGRAPh Light-sheet (e) Moc = 0.60, a = 20 ° .

Light-sheet (f) M_ = 0.80, a = 20 °.

Figure 36. Concluded.

w L

t

ORIGIHAL PHOTOGRAPH BLACK AND WHITE Light-sheet

n 7

(a) fl = 0°, FS 525 (31.50).

Light-sheet (b) fl = 4 °, FS 525 (31.50).

Figure 37. Effect of sideslip on cross-flow patterns about baseline F/A-18 model at Moo = 0.6; Ree = 1.32 × 106; and c_= 20 °. Dimensions are in inches full scale (0.06 scale).

ORIGINAL PAGE BLACK AND WHITE PHOTOGRAP_ Light-sheet (c) fl = 0°, FS 567 (34.02).

!

- Light-sheet Figure 37. Concluded.

i (d) fl = 4°, FS 567 (34.02). _ i=- 74 _ !-- ORIG ;,".,_,_ !_ P.'..3E BLACK AND WHI_£ I'f_oi-OGRAPH t (a) FS 163 (9.78).

t {0) Close-up of vortex pair.

Figure 38. Forebody cross flow about baseline F/A-18 model at Mec -- 0.60, Ree -- 1.32 × 106, and a = 50 °.

Dimensions are in inches full scale (0.06 scale).

a, deg

O 10.03

[3 15.03

O 20.05 -3.0 - j,, 24.92 tx 30.05 FS 253 (15.18) -2.5 -

Cp,u

-1.5 -2.0 _/%_ -1.0 B I

0 t "_---°r-°--r °

.2 .4 .6 .8 1.0

-1.0 -.8 -.6 -.4 -.2

y/s E -3.0 :-=

|

-2.5 FS 296 (17.76)

E

Cp,u r

o I "_T-_-rO J °-r--°-_ I J J

-1.0-.8-.6-.4-.2 0 .2 .4 .6 .8 1.0 y/s -3.0 -2.5 FS 357 (21.42) -2.0 / Cp,u -1.5 -1.O -.5

b--o-q

1 I I I -1.0 -.8 -.6 -.4 -.2 0 .2 ,4 .6 .8 1.0 y/s (a) a=10 °to30 ° .

Figure 39. Effect of angle of attack on baseline F/A-18 model LEX surface static pressures at Mzc = 0.40 and Ree = 1.75 × 106. Dimensions are in inches full scale (0.06 scale).

= OC, deg 35.04 o 39.93 [3 45.09 49.99 FS 253 (15.18) -3.0 - zx

-2.5

-2.0

Cp,u -1.5

-1.0

m -.5 I ______1 _ A_t-----A I I I I 0 .2 .4 .6 ,8 1.0 -1.0 -.8 -.6 -.4 -.2 yls -3.0 FS 296 (17.76) -2.5 -2.0 Cp,u -1.5 -1.0 -.5 I I I I I I I I I I .8 1.0 0 .2 .4 .6 -1.0 -.8 -.6 -.4 -.2 y/s -2.5 FS 357 (21.42) -3.0 f -2.0 Cp,u-1,5 -1.0 m -.5 L I I I ] A__I I [ 1 J -1.0 -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 y/s (b) a=35 ° to 50 ° • Figure 39. Concluded.

\ \ a, deg O 35.02 [] 39.97 45.05 FS 253 05.18) 49.95 -2.5 A -3.0 f <> Cp,u -I.5 -2.0 -1.0 ".5 0 I I I I I 1 I I I I -1.0 -.8 -.6 -.4 -,2 0 .2 .4 .6 .8 1.0 y/s -3.0 - -2.5 FS 296 (17.76) m -2.0 Cp,u -1.5 n -1.0 -.5 I 1 I I -1.0 -.8 -,6 -.4 -.2 y/s -2.5 FS 357 (21.42) -3.0 f -2.0 Cp,u -1.5 =.5 -1.0 --i_1 _ 0 I I I I I I -1.0 -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 y/s (b) _=35 ° to50 °.

Figure 40. Concluded.

(z, deg O 10,11 14.90 [] -3.0 - 20,16 <> FS 253 (15.18) A 24.93 -2.5 - -2.0 - m Cp,u ol.5 -1.0 -.5 ° I--+7°-q_ I i

I i

.2 .4 .6 .8 1.0 -.8 -.6 -.4 -.2 -1.0 y/s

|

-3.0 FS 296 (17.76) -2.5 -2.0 Cp,u -1.5 -

-1.oL _"-,'-__-

i

o--_---r°"_ I

I I 1.0 0 .2 .4 .6 .8 -1,0 -.8 -,6 -.4 -.2 y/s -3.0 FS 357 (21.42) -2.5 -2.0 Cp ,u -1.5 -.5 -1.0 0 I I .2 .4 ,6 .8 1.0 -1.0 -.8 -.6 -.4 -.2 y/s (a) c_=10 ° to 25 ° • m Figure 41. Effect of angle of attack on baseline F/A-18 model LEX surface static pressures at Moo = 0.80 and R% = 1.02 x 106. Dimensions are in inches full scale (0.06 scale).

!

i m B

i

|

!

!

a, deg 30.05 -3.0 - [] 35.04 Fs 253 (15.18) <> 40.07 -2.5 -2.0 Cp'u -1.5 -1.0 ".5 I I 1 1 I I I I I I -1.0 -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 y/s -3.0 - -2.5 FS 296 (17.76) -2.0 Cp,u -1.5 -1.0 -.5 I I I I [ I • -1.0 -.8 -.6 -.4 -.2 0 .2 .4 -3.0 -2.5 FS 357 (21.42) -2.0 Cp,u . 1.5 -1.0 B %5 I 1 I I I I I I I 1 -I.0 -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 y/s (b) a=30 ° to 40 ° .

Figure 41. Concluded.

ct,deg

o 10.03

[] 15.15

-3.0- O 20.16

-2.5- A 25.05

-2.0 - !

Cp,u -1.5 - -1.0 r A A _^--a_ _/A_A--_-a_ A i -.5 _ _ 01 I I 1 I / I -I.0 -.8 -.6 -.4 -.2 0 .2 A .6 .8 1.0 y/s -3.0 | -2.5 -2.0 i Cp,u -1.5 - 1 0 AA_A--A--A--A_'A A IA--A_ A.A-A AAA 0 I I -1.0 -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 y/s

i

-3.0 - FS 357 (21.42) -2.5 - -2.0 - Cp,u -1.5 - i -.5 -1.0 __ 0 I .2 ,4 .6 .8 1.0 -1.0 -.8 -.6 -.4 -.2 0 IE= y/s m (a) a=10 ° to 25 ° .

Effect of angle of attack on baseline F/A-18 model LEX surface static pressures at M_ = 0.90 and Figure 42.

E Ree = 1.02 x 106 Dimensions arc in inches full scale (0.06 scale).

!

!

m m 3o,3O o 3_91 39.99 .Z A.O F'o" _ _.0 0,0 "'_ "'6 "'¢ .3.0 T" _Is "ES 'Z96 01 .%.)

.'Z,5 _jIs 0A,0 "'_' ..6 -A, ,/Is _,:_F_e 4tZ. Co_CI_,6e6" Moo o 0.20 t [] 0.30 0 0.40 A 0.60 -_ -3.0 - _ 0.70 t_ 0.80 -2.5 - o 0.90 -2.0 - Cp,u -1.5 i -.5 i -l.0_ E ?1.O-.8-.6-.4-.2 0 .2 .4 .6 .8 1.0 | y/s -3.0 - -2.5 - FS 296 (17.76) Cp,u -1.5 - -2.0 - _ i -.5 _

o

0 I I -1.0 -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 _- y/s -3.0 - FS 357 (21.42) Cp,u .8 1.0 y/s Figure 43. Effect of Mach number on baseline F/A-18 model LEX surface static pressures at a = 10 °.

Dimensions are in inches full scale (0.06 scale).

Moo O 0.20 0.30 [] 0.40 <> A 0.60 FS 253 (15.18) t,. 0.70 0.80 0.90 -2.5 c3 -3.0 f th -2.0 Cp,u -1.5 -.5 -l.0 I I I I I .2 .4 .6 .8 1.0 -1.0 -.8 -.6 -.4 -.2 y/s -3.0 FS 296 (17.76) -2.5 -2.0 -.5 I L I I I I I I I I 1.0 0 .2 .4 .6 .8 -1.0 -,8 -,6 -,4 -,2 y/s -3.0 - -2.5 - FS 357 (21.42) Cp ,u -1.5 -1.0 -2.0 -.5 I I I I I 0 I I I I .2 .4 .6 .8 1.0 -1.0 -.8 -.6 -.4 -.2 y]$ Figure 44. Effect of Mach number on baseline F/A-18 model LEX surface static pressures at c_ = 20 °.

Dimensions are in inches full scale (0.06 scale).

Moo

o 0.20

i-1 0.30 <> 0.40 A 0.60 0.70 -3.0 - D 0.80 FS 253 (15.18) -2.0 0.90 CP ,u -1.5 -2.5 __ -1.0 -o -.5 - 0 I I I 1 I I I • I I -1.0 -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 y/s FS 296 (17.76) I I I I y/s -3.0 I -2.5 FS 357 (2 t.42) Cp,u -1.5 -2.0 __ -1.0 -,5 - 0 I I I I I I I I I I -1.0 -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 y]s Figure 45. Effect of Mach number on baseline F/A-18 model LEX surface static pressures at a = 30 °.

Dimensions are in inches full scale (0.06 scale).

Moo o 0.20 0.30 [] <> 0.4O A 0.60 Ix 0.70 [3 FS 253 (15.18) -3.0 -2.5 -2.0

Cp,u

-1.5 _-'_ IA.---'- _[3-.- [3--43..lA : Cy.£>__.-.._---_----____ _ -1.0 N -.5 I I I I J I I I I 1 -1.0 -.8 -°6 -.4 -°2 0 .2 .4 .6 .8 | .0 y/s -3.0 - FS 296 (17.76) B -2.5 -2.0 Cp,u -1.5 -1.0 -.5 - . I I I I I I 0 I I I I 0 .2 .4 .6 .8 1.0 -1.0 -.8 -.6 -.4 -.2 y/s -3.0 -2.5 FS 357 (21.42) -2.0 Cp ,u -1.5 -1.0 -.5 I I I I • I I I I I I 0 .2 .4 .6 .8 1.0 -.8 -.6 -.4 -.2 -1.0 y/s Figure 46. Effect of Mach number on baseline F/A-18 model LEX surface static pressures at a = 40 °.

Dimensions are in inches full scale (0.06 scale).

Me, o o 0.20 t3 0.30 <> 0.40 FS 253 (15.18) -3.0 t t -2.5

i

-2.0 q Cp,u -1.5 IE -1.0 -.5 I I J

_B

0 .2 .4 .6 .8 X .0 -X ,0 -,8 -.6 -.4 -,2 yls i FS 296 (17.76) E y/s -3.0 FS 357 (21.42) -2.5 -2.0 Cp ,u -1.5 0 [____1--_ l I I 0 .2 .4 .6 .8 1.0 -1.0 -.8 -.6 -.4 -.2 y/s Figure 47. Effect of Mach number on baseline F/A-18 model LEX surface static pressures at a = 50 °.

Dimensions are in inches full scale (0.06 scale).

R

ss \ _- ss ss

R R R LS TS TS B

I I

LP TRP TP Figure 48. Sketches of cross-flow patterns about a body at a high angle of attack (from ref. 2).

a, deg 0 10.03 [] 15.03 FS 107 (6.42) 180° <> 20.05 A 24.92 270 ° -2.0 r b. 30.05

900@

-1.5 0° -1.0

i

Cp

.5 I

i

1.0 1 I I I I I I I I I I I 0 30 60 90 120 150 180 210 240 270 300 330 360

E

O, deg a, deg = 35.04 E c3 39.93 E P -.= o 45.09 r -2.0 - 0 49.99 =m -1.5 - 000,,,, .,.000_ / v'o _ \ -1.0 - o

q, -.5

.5 1.0 ]' " 1 I I I 1 I I I I I I 1 0 30 60 90 120 150 180 210 240 270 300 330 360 O, deg (a) FS 107 (6.42).

Figure 49. Effect of angle of attack on forebody surface static pressures at Moc = 0.40 and Ree = 1.75 × 106.

Dimensions are in inches full scale (0.06 scale).

9O a, deg o 10.03 FS 142 (8.52) [] 15.03 180° zx 24.92 tx 30.05 90 ° 270 ° 0 20.05 0o I I I I I I I I I I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 O, deg a, deg 35.04 D 39.93 o 45.09 0 49.99 -2.0 I -1.5 ,4- 0 0-0- -1.0 Cp -.5 .5 I I I I I I I I I I I I 1.0 0 30 60 90 120 150 180 210 240 270 300 330 360 O, deg (b) FS 142 (8.52).

Figure 49. Continued.

o_, dcg

i o 10.03 FS 184 (11.04) -- D 15.03 180 ° <> 20.05 A 24.92 /' ] '_ -2.0 - tx 30.05 90o____4_270 o I I I

• i

-.5 -- :

i. r

1.0 _ _ ! _ _ i _ i _ _ , i i _

0 30 60 90 120 150 180 210 240 270 3oo 330 360

|

O, deg _, deg c3 39.93 o 45.09 t3 35.04 i -2.0 F 0 49.99 /

1.o -o-o-_ __

Cp -.5

.i

I I 1.0 l 1 I I I I 1 I I I I z 0 30 60 90 120 150 180 210 240 270 300 330 360 O, deg (c) FS 184 (11.04).

Figure 49. Concluded.

-2.0 - FS 107 (6.42) m -1.5 180 `> -1.0 90 _ cp -.5 _70o 0 o .5 I I I I I I I I I I I I 1.0 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg FS 142 (8.52) 180 ° q, 90° @ 270° 0 o 0, deg -2.0 - FS 184 (11.04) 180 ° -1.5 - m -1.0

q, .5

90°_ 270° 0`> .5 I I I I I [ I I I I I I 1.0 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg (a) a = 40 °.

Figure 50. Forebody surface static pressures on baseline F/A-18 model at Mo_ = 0.40 and Ree = 1.75 x 106.

Dimensions are in inches full scale (0.06 scale).

-2.0F

-1.5 -

FS 107 (6.42) 180 °

-1.0

cv -.5 90 ° _ 270o 0 ° .5 1.0 < I 1 I I I I I I I I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg -2.0 - -1.5 FS 142 (8.52) 180 ° 90° _ 270° .5 0 ° 1.0 I I I I I l I I I I I 1 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg f FS 184 (11.04) 180 ° -1.0 Or, -.5

o oo

.5 1.0 ( I I I I I I 1 I 1 1 1 I 0 o 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg (b) a=50 °.

Figure 50. Concluded.

a, deg O 10.00 FS 107 (6.42) [] 15.01 180 ° Z_ 25.01 <> 20.05 -2.0 tx 30.05 90° _ _70° -1.5 0o -1.0 Cp -.5 .5 I I I I I I I I I I I I 1.0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg a, deg th 35.02 O 39.97 O 45.05 -2.0 - 0 49.95 -1.5 -1.0 Cp -.5 .5 I 1.0 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg (a) FS 107 (6.42).

Figure 51. Effect of angle of attack on baseline F/A-18 model forebody surface static pressures at M_c = 0.60 and Ree -- 1.32 x 106. Dimensions are in inches full scale (0.06 scale).

0_, dog

FS142 (8.52)

o 10.00 180 °

20.05

A 25.01 90 ° 270 °

v,, 30.05

o 15.01

0° i I I I I I I I I I 1 I I 0 30 60 90 120 150 180 210 240 270 300 330 360 O, deg a, deg 35.02 o 39.97 o 45.05 0 49.95 -1.5 -2.0 i -1.0 Cp -.5 E .5 I I 1 I I I I I I I I I 1.0 0 30 60 90 120 150 180 210 240 270 300 330 360 O, deg (b) FS 142 (8.52).

Figure 51. Continued.

or, deg

o 10.00

FS 184 (11.04) 180"

[] 15.01

0 20.05

/x 25.01

-2.0 -

tx 30.05

-1.5 -

90°_ 270°

-1.0 -

Cp -.5 -

110 I 0 30 60 90 120 150 180 210 240 270 300 330 360 O, deg cx,deg 35.02 39.97 o 45.05 -2.0 0 49.95 -1.5 -1.0 Cp -.5 .5 1.0 I I I I I I I I I I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 O, deg (c) FS 184 (11.04).

Figure 51. Concluded.

= -2.0 - FS 107 (6.42) 180 ° -1.5 -1.0

q, -.5

90° _ 270° 0 o .5 I I I I I I I I I I I I 1.0 0 30 60 90 120 150 180 210 240 270 300 330 360 O, deg -2.0 - FS 142 (8.52) i 180° -1.5 -1.0 Cp -.5 90°_ 270° ii o .5 [ I 1 I I I I I I I I 1.0 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg -2.0 - FS 184(11.04) 180 ° -1.5 -1.0 -.5 90°_ 270° .5 OO ( I I I I I I I I I I I 1.0 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg (a) c_ = 40 °.

Figure 52. Forebody surface static pressure on baseline F/A-18 model at Moo = 0.60 and 1Ree = 1.32 x 106.

Dimensions are in inches full scale (0.06 scale).

-2.0-

-1.5 - FS 107 (6.42) 180 ° 270 ° q, -1.0 - / 90° 0 ° 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg -2.0 - -1.5 FS 142 (8.52) -1.0 180 ° -.5 90°@ 270° .5 1.0 _ I I I I I I [ I I I I I OO 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg -2.0 - FS 184 (11.04) -1.5 180 ° -1.0 Cp -.5 90°_ 270° .5 0 o 1.0' I I I I I I I I I I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg (b) a = 50 °.

Figure 52. Concluded.

a, deg FS 107 (6.42) O 10.11 180° [] 14.90 O 20.16 270 ° -2.0 | Z_ 24.93 90° -1.5

]

0° | -1.0

q, -.5

.5 R I I I I I 1 1 1 1 I I I 1.0

E

0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg

F

a, deg t3 30.05 c_ 35.04 -2.0 o 40.07 -1.5 w

F

-1.0

Cp -.5

.5 1.0 I I 1 I I I I I I I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg (a) FS 107 (6.42).

Figure 53. Effect of angle of attack on baseline F/A-18 model forebody surface static pressures at Moo = 0.80 and Ree = 1.02 × 106. Dimensions are in inches full scale (0.06 scale).

a, deg FS 142 (8.52) 0 10.11 180° [] 14.90 20.16 -2.0 -- zx 24.93 270 _ 90° -1.5 - 0 o -1.0 - Cp -.5 -

i

1.0 I 0 30 60 90 120 150 180 210 240 270 300 330 360 O, deg or,deg 30.05 C3 35.04 -2.0 Q 40.07 -1.5 -1.0 Cp -.5 .5 I I I I I I I I I I I I 1.0 0 30 60 90 120 150 180 210 240 270 300 330 360 O, deg (b) FS 142 (8.52).

Figure 53. Continued.

i FS 184 (11.04) a, deg 180° J o 10.11 _ !- [] 14.90

/ I '_ o '

-_o - ? -_°'!6 _°°-IBI_k _°

-1.0 - i Cp -.5- : l

- 1.0 I I l I I, I l I I [ I l i

0 3-0 60 90 120 150 180 210 240 270 300 330 360 0, deg a, deg th 30.05 35.04 -2.0 r o 40.07 / -1.5 -1.0 Cp -.5 1.0l I I I I I I { I I 1 I I 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg (c) FS 184 (11.04).

Figure 53. Concluded.

{ 1 " k

-2.0 - -1.5 FS 107 (6.42) 180 ° -1.0 C_ -.5 90° _ 270° 0 ° ( 1.0 I I I I I I 1 I I I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg -2.0 -1.5 FS 142 (8.52) 180 ° -1.0

q, -.5

90°_ 270° .5 I OO I I I [ I I I I I I I I 1.0 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg -2.0 -1.5 FS 184 (11.04) 180 ° c_ -.5 c* 0 P "x_7_t3Ct3_OOOOOOOO_ - _ 90° 270° -1.0 __o .5 0 o 1.0l I I I I I I 1 I ! 1 I I 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg Figure 54. Forebody surface static pressures on baseline F/A-18 model at M_o = 0.80, Ree = 1.02 x 106, and a = 40 °. Dimensions are in inches full scale (0.06 scale).

a, deg o 10.03 FS 107 (6.42) 180 ° 20.16 -2.0 270 ° lx 25.05 9°°

J

[] 15.15 -1.5 0o -1.0

%

-.5 - -zx.... _I- - _ i

|

0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg

P

a, deg th 30.30 C3 34.92 39.99 -1.5

m

-2.0 f -1.0 Cp -.5 .5 I I I t I I I I I I I I 1.0 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg (a) FS 107 (6.42).

Figure 55. Effect of angle of attack on baseline F/A-18 model forebody surface static pressures at M_ : 0.90 and Ree = 1.02 × 106. Dimensions are in inches full scale (0.06 scale).

a, deg 0 10.03 FS 142 (8.52) [] 15.15 180° -2.0 - ZX 25.05 90° 70° -1.5 - 0 20.16 _2 -1.0 - Cp -.5 - 1.0 I 0 30 60 90 120 150 180 210 240 270 300 330 360 O, deg a, deg th 30.30 a 34.92 -2.0 0 39.99 -1.5

F

-1.0 Cp -.5 .5 1.0 I I I I I I I I ] I I J 0 30 60 90 120 150 180 210 240 270 300 330 360 O, deg (b) FS 142 (8.52).

Figure 55. Continued.

a, deg FS 184 (11.04) o 10.03 180 ° [] 0 20.16 -2.0 - A 25.05 90 ° 270 c -1.5 - 15.15 -1.0 - i i- Cp -.5 - i i I 1.01 1 I I I I I I I I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg or,deg th 30.30 C3 34.92 -2.0 - o 39.99 -1.5 - -I.0 - Cp -.5 - .5 I I I I I 1 1 I I I I I 1.0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg (c) FS 184 (11.04).

Figure 55. Concluded.

-2.0 FS 107 (6.42) -1.5 180 ° -1.0 270 ° 90 ° cp -.5 0 ° q 1.0 I I I I I 1 I I I 1 I I 0 30 60 90 120 150 180 210 240 270 300 330 360 O, deg -2.0 FS 142 (8.52) -1.5 180 °

-1.0 F

Cp -.5 270 ° 0 o .5 ( 1.0 I I I I l I.. I I [ I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 O, deg -2.0 - FS 184 (11.04) 180 ° -1.5 - -1.0 - Cp -.5- 90°_ 270° O

.5 _.' 'b

1.o. , t I I , , , I , I , , 0 30 60 90 120 150 180 210 240 270 300 330 360 O, deg Figure 56. Forebody surface static pressures on baseline F/A-18 model at M_ = 0.90, Ree = 1.02 × 106, and a = 40 °. Dimensions are in inches full scale (0.06 scale).

Boundary-layer separation 7

Primary vortex pair --_ /

Shock wave i Subsonic Transomc Hypothesized cross-flow shock-induced boundary-layer separation on forebody of 0.06-scale F/A-18 Figure57.

model.

E E = E D E Mo_ 0.20 o 0.30 1-1 0.40 /x 0.60 b. 0.70 0.80 c5 0.90 2.4 c?

2.0 1.6 CL CL 1.2 .8 .4 I I I I I I 0 I 20 30 40 50 60 0 .4 .8 1.2 1.6 2.0 2.4 a, deg CD Cm Figure 58. Effect of Mach number on lift, drag, and pitching-moment characteristics of baseline F/A-18 model.

a, deg o 20.02 [] 25.03 0 30.04 .03 - .01 - .02 - _ Cl 0 -.01 - _ -.02 - ] i | , -.03 I I I I _ I I I • I I i -1( -8 -6 -4 -2 0 2 4 6 8 10 i i i ! 13, deg .03 - I . .02 - N -- .01 - _ ___, _o___-o - Cn 0 _ -.01 - -_ = -.02 - -.03 1 I I I I I " I J I I -10 -8 -6 -4 -2 0 2 4 6 8 I0 2 l_, deg -= .2 0 "" Cy -.4 I I 1 I i I I i I I _- -10 -8 -6 -4 -2 0 2 4 6 8 10 --'- 13, deg (a) c_ = 20° to 30°.

Figure 59. Variations of rolling-moment, yawing-moment, and side-force coefficients with sideslip at M_ = 0.60 and Ree = 1.32 x 10 6.

110 : i t_, deg o 30.04 [] 35.05 0 40.02 .03 - .02 - .01 - CI 0 -.01 - -.02 - -.03 I I I I I I I I I I -1_ -8 -6 -4 -2 0 2 4 6 8 10 13, deg .03 - .02 - .01 - Ca -.01 - -.02 - -.03 I.. I I I 1 I I I I I -10 -8 -6 -4 -2 0 2 4 6 8 10 13, deg .2- Cy -,2 i -.4 I I I I I 1 I I I I -10 -8 -6 -4 -2 0 2 4 6 8 10 Figure 59. Concluded.

_, deg

0 20.05

O 25.04

0 30.07

.02

.01

0 ---_

Cl

.o1

-.02 - -.03 I I I l I I t I I 1 -11 -8 -6 -4 -2 0 2 4 6 8 10 _, deg .03 F .0 Cn _ -_ ..._/_.. - - -.01 - % -.02 - _'_ -.03 I I I I t 1" I I I I -10 -8 -6 -4 -2 0 2 4 6 8 10 _], deg .2 Cy -.2 1 I I I I I I I I I -,4 0 -8 -6 -4 -2 0 2 4 6 8 !0 6, deg (a) a=20 ° to 30 ° .

Figure 60. Variations of rolling-moment, yawing-moment, and side-force coefficients with sideslip at M_o = 0.80 and Ree. = 1.02 × 106.

a, deg o 30.07 [] 35.03 O 40.03 .02 .01

.03

ci 0

-.01 - -.02 - -.03 ] 1 I I I 1 I I I I -11 -8 -6 -4 -2 0 2 4 6 8 10 13, deg .01 -_'__ Ca -.01 - _'0_ -.02 - --x_ -.03 I I I I I I I I I I -10 -8 -6 -4 -2 0 2 4 6 8 10 13, deg Cy -.2 I I I I t I I .1 I I -,4 -10 -8 -6 -4 -2 0 2 4 6 8 10 13, deg (b) a=30 ° to 40 ° .

Figure 60. Concluded.

a, deg o 20.01 [] 24.94 0 29.97 .02

°3f

,01

o

CI

i

-.01 - O_-_ -.02 - -.03 t 1 I I I 1 I I I I

|

-11 -8 -6 -4 -2 0 2 4 6 8 10 = 13, deg m = .03 z .02 .01

Ca F

O' -.01 m -.02 t I I I I I I I I 1 -.03 -1t -8 -6 -4 -2 0 2 4 6 8 10 13, deg .2 Cy -.2 t I ] I I I t I I I -.4 -8 -6 -4 -2 0 2 4 6 8 10 -10 13, deg (a) a=20 ° to 30°.

Figure61. Variations ofrolling-moment, yawing-moment, and side-force coefficients with sideslip at Moo = 0.90 and Re_ = 1.02x 106.

cx, deg o 29.97 [] 35.04 0 39.98 .02

o3_

Cl

O_o3, , , ,;--___'_

-10 -8 -6 -4 -2 0 2 4 6 8 YO 13, deg .03 .02 .01 Ca -.01 0 -.03 I I t 1 I I I I "_ -.02 _0 -11 -8 -6 -4 -2 0 2 4 6 13, deg Cy -.2 -.4 I I I I I I I I -10 -8 -6 -4 -2 0 2 4 6 8 10 _, deg (b) a=30 ° to 40 ° .

Figure 61. Concluded.

ORIGINAL PAGE BLACK AND WHITE PHOTOGRAPh

|

D

m

(a) LEX fence; FS 411 (24.66).

F E

J

(b) Baseline; FS 411 (24.66).

= Figure 62. Laser vapor screen flow visualizations on LEX fence and baseline configurations at 2_,f_ = 0.40, _ Ree = 1.75 x 106, and a = 20 °. Dimensions are in inches full scale (0.06 scale).

O__!C-_INAL PAGE AND WHITE PHOTO(2RAPH (c) LEX fence; FS 450 (27.00).

(d) Baseline; FS 450 (27.00).

Figure 62. Continued.

ORIGINAL PAGE BLACK AND WHITE PHOTOGIRAPn (e) LEX fence; FS 525 (31.50).

(f) Baseline; FS 525 (31.50).

Figure 62. Concluded.

O.4GINAL PAGE BLACK AND WHITE PHOTOGRAPH LEX vortex LEX vortex (a) FS _ 357 (21.42).

Figure 63. Close-ups of cross flow about LEX fence configuration at _[ac = 0.60 and Ree = 1.32 x 106.

Dimensions are in inches full scale (0.06 scale).

i ORIGINAL PAGE BLACK AND WHITE PHOTOGRAPH = F (b) FS _ 411 (24.66).

Figure 63. Continued.

F i ORIGINAL PAGE BLACK AND WHITE PHOTO_3RAPH (c) FS _ 450 (27.00).

Figure 63. Continued.

ORIGINAL PAOE R.I.ACK AND WHITE F_T_pk_ = (d) FS _ 483 (28.98).

Figure 63. Concluded.

ORIGI'.'_r BLACK AND WHITE PHOTOGRAPH Light-sheet _n (a) LEX fence; FS 411 (24.66).

_Light-sheet location (b) Baseline; FS 411 (24.66).

Figure 64. Cross flow about LEX fence and baseline configurations at Moo = 0.60, Ree = 1.32 × 106 , and o_ = 20 °. Dimensions are in inches full scale (0.06 scale).

ORIGINAL PAGE

BLACK AND WHITE PHOTOGRAPH (c) LEX fence; FS 450 (27.00).

t Light-sheet (d) Baseline; FS 450 (27.00).

Figure 64. Continued.

BLACK AND WHJTE _'r_OiOGRAPH Light-sheet (e) LEX fence; FS 483 (28.98).

Light-sheet (f) Baseline; FS 483 (28.98).

Figure 64. Continued.

ORIGINAL PAGE BLACK AND WHITE PHOTO(3RAPH i i B LI__ m fl onTight_sheet !

Corotating vortices

B (g) LEX fence; FS 525 (31.50).

Z Light-sheet (h) Baseline; FS 525 (31.50).

Figure 64. Concluded.

BLACK AND W_H.T_ PHOTOGRAPH Light-sheet _on (a) LEX fence; FS 411 (24.66).

_Lilgohtat_hoe: t (b) Baseline; FS 411 (24.66).

Figure 65. Cross flow about LEX fence and baseline configurations at 2riot = 0.60, Re_ ---- 1.32 × 106, and = 25 °. Dimensions are in inches full scale (0.06 scale).

ORIGINAL P =,!_S BLACK AND WHITE PHOTOGRAPh _Light-sheet == (c) LEX fence; FS 450 (27.00).

!

t Light-sheet -_cat_7

g

(d) Baseline; FS 450 (27.00).

Figure 65. Continued.

OR!Gi:,;:',L 7, ...:; BLACK AND WHIT_- P_JO]OGRAPH

Downward movement of LEX vortex

Light-sheet

7:z

(e) LEX fence; FS 483 (28.98).

Vortex breakdown

Light-sheet (f) Baseline; FS 483 (28.98).

Figure 65. Continued.

ORIGINAL PAGE BLACK AND WHITE PHOIOGRAP_ Light-sheet

_on/

(g) LEX fence; FS 525 (31.50).

Light-sheet

__on/

(h) Baseline; FS 525 (31.50).

Figure 65. Concluded.

I10 110 I 100

,2o/(

4ofps / so

Wing TE Baseline LEX LEX + fence (a) Mean velocity contours.

15 ° kxS_ _"_ .._. 0 ° _k_'O _

/ _oo",_°',, _.G/

10 o j / 5° 2 ° 1

Y_-oo I

.-/¢a-7-:_ -_° ,o

'x \'1 t_)°°l _ __oo

"_ -20° I

LEX + fence Baseline LEX (b) Angularity contours.

Figure 66. Effect of mean velocity and flow angularity contours near vertical tails of 0.083-scale F/A-18 model at 3_c = 0.07 and a _ 29 °. (Unpublished McDonnell Douglas data.)

o LEX fence off [] LEX fence on FS 404 (24.24) (LEX-wing-LE junction) FS 434 (26.04) (LEX-wing-LE-flap-hingeline junction) FS 253 % FS 525 (31.50) (Vertical tail LE-fuselage junction) (15"18F)S2!6 _ w 30 ] (17.76) ]u x deg 25 FS 357 (21.42) O B " / "l 0 I I I I I I I I I I 0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 x/l Figure 67. Effect of LEX fence on vortex breakdown progression with angle of attack at Mcc = 0.60 and Rea = 1.32 x 106. Dimensions are in inches full scale (0.06 scale).

ORIGINAL [ _,,,_,_=" BLACK AND WHITE PHOTOGRAPH (a) F-18 HARV; Moc _ 0.27; Ree _ 13 x 106.

(b) 0.06-scale F/A-18; Mz¢ : 0:40; Ree : 1.75 x 106.

Figure 68. Flight and wind tunnel flow visualizations at c_ = 20 ° with LEX fences on.

0 LEX fence off LEX fence on [3 FS 253 (15.18) -2.5 -2.0 -3.0 f Cp,u -1.5 -,5 aop: 0 1 I I I I -1.0 -.8 -.6 -.4 -.2 .2 .4 .6 .8 1.0 y/s -3.0 - FS 296 (17.76) Cp,u -.5 0 I I 1 I I I I I I I -1.0 -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 y/s -3.0 - -2.5 - FS 357 (21.42) m -2.0 Cp,u -1.5 -1.0 -.5 I I I I i I I I I I -1.0 -.8 -.6 -.4 -.2 0 ,2 .4 .6 .8 1.0 y/s (a) a = 20 °.

Figure 69. Effect of LEX fence on LEX surface static pressures at M_ = 0.30 and Ree = 1.40 x 106. Dimensions are in inches full scale (0.06 scale).

o LEX fence off [] LEX fence on -3.0 - -2.5 - -.5 [-- 0 I I i I I -1.0 -,8 -.6 -.4 -.2 0 .....

i

y/s -3.0 r- .2.5 L Fs 296 (17.76) Cp,u -1.5 - x -1.0 - ".5 0 I I I I I I I "l.0 ".8 -.6 -,4 ".2 0 ,2 ,4 y/s -3.0 - FS 357 (21.42) -2,5 -2.0 ,U -1.5 -1.0 -.5 I 1 I I I I I I I I -1.q -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 y/s (c) a = 30° z: Z Figure 69. Continued.

11.o

C_4 -

-,.or __ '---\ ,>

.6 "

-_V ,o___' °'°

01 0 ""_ -.6 '

-,°V ,_.o

.6 " -._ 04.0 ..S ..6 "" _/Is _g_e 69. C°_t_ec_' LE% _ence off O LEX _once on _3 _y 7.

.%0 C_,u A,5 - -1.0 - 1.0 0_ 0 "" ._..0 "" ' -,jIs -2.5I

.%0

Cp,u .1.5 " Cs FS 357 (._-_'4T)

\

._ .5 -3.0 ___ C_,_ .

\ '"-_5_" , _, -.?. 0 0 _ -.6 "" A.O ". _ls 1B8

LE,X fence off

o L_X_ence on

C3

.?..5 -2.0 1.0 O_ 0 .z " -1 u -,v • ' _'/s -2.5 -2.0 Cp,u .'t _=_ 't,O O_ 0 .2 "* " .1 0 -- "' • ,jIs (_) c__ _o°.

0,60 and Ree _ 1.32 x 106. DimenSions f LEX {ence on LEX stXr{_ce st_tkc preSslares at Moo ' are 70. F_ectO o,._ e (0.06 sc_e)- Fxg are in _ncb-es trot _,_al .

_3 -3.o ._,51 .I, __.0 \ -2.5 .'_ . ' ,6 '_ 1` .0 -t.O ". ' ,jIs 1.,_o o LEX fence off LEX fence on [] -3.0 - FS 253 (15.18) -2.5 - -2.0 Cp,u _1.5 -1.0 -.5 I I I I I 0 I I I I .2 .4 .6 .8 1.0 -1.0 -.8 -.6 -.4 -.2 0 y/s FS 296 (17.76) -.5 0 I I I I I I 1 I I I -1.0 -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 y/s -3.0 - -2.5 FS 357 (21.42) w -2.0 Cp,u -1.5 -Zl..._-O_O -1.0

@

=.5 I I I I I I 0 I I I I 0 .2 .4 .6 .8 1.0 -1.0 -.8 -.6 -.4 -.2 y/s (c) a = 30 °.

Figure 70. Continued.

o

LEX fence off

[]

LEX fence on

-3.0-

FS 253 (15.18)

-2.5-

Cp,u -1.5

-1.0 -

-,5 0 I I I I I I I I I I -1.0 -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 y/s -3.0 - FS 296 (17.76) -2.5 - -2.0 - Cp,u. 1.5

_

-1.0 - -.5 - 0 I I I I I I I I I I -1.0 -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 y/s -3.0 - -2.5 - FS 357 (21.42) -2.0 - Cp,u . 1.5 -1.0 -.5 = I I l I 0 1 I I I I I -1.0 -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 m y]s Z = (d) a = 40 °.

Figure 70. Concluded.

LEX fence off o LEX fence on [] FS 253 (15.18) -3.0 -2.5 -2.0 Cp ,u. 1.5 -1.0 -.5 _L____L__ I I J [___L___L_-- t I .2 .4 .6 .8 1.0 -1.0 -.8 -.6 -.4 -.2 y/s -3.0 FS 296 (17.76) -2.5 -2.0 Cp,u -1.5 -1.0 - CKI_.t3.M3---"I_--_ Ct_ -.5 A I I I I d I I I I 0 .2 .4 .6 .8 1.0 -.8 -.6 -.4 -.2 -1.0 y/s -3.0 - -2,5 FS 357 (21.42) -2.0 Cp,u . 1.5 -1.0 _o5 0L I 1__1 I 0 .2 .4 ,6 .8 1.0 -1.0 -,8 -.6 -.4 -.2 y/s (a) _ = 20 °.

Figure 71. Effect of LEX fence on LEX surface static pressures at M_o = 0.80 and Ree = 1.02 x 106. Dimensions are in inches full scale (0.06 scale).

o LEX fence off [] LEX fence on P -3.0 - -2.5 - Cp,u -1.5 - -2.0 - _ _ i -.5 0 I I I I I I I I I I -1.0 -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 y/s -3.0 - _ -2.5 - FS 296 (17.76) _ Cp,u -1.5 - -1.0 - cr_"_"c"'_-_ _-4_.--,_:_ -.5 - _ 0 • 1 I I I 1 1 I I I -1.0 -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 y/s m -3.0 m -2.5 FS 357 (21.42) -2.0 B Cp,u -1.5 _"-gl -1.0 -.5 m I I 1 I 1 I 0 I I I I 0 .2 .4 .6 .8 1.0 -1.0 -.8 -.6 -.4 -.2 y/s (b) a = 25 °.

Figure 71. Continued.

m O LEX fence off C2 LEX fence on -3,0 - -2.5- -2.0- Cp,u -1.5 - __..__ -1.0 - -.5 -1.0 -.8 -.6 -.4 y/s FS 296 (17.76) -3.0 - -2.5 - -2.0 - Cp,u -1.5 Cl¢:_a..-_-'_===_ -1.0 -.5 -I,0 -.B -.6 -.4 y/s -2.5 -2.0 Cp,u -1.

-.5 1_0 0 -.2 -I.0 -.B -,6 -.4 fls (c) a = 30°.

Figure 71. Continued.

O LEX fence off [] LEX fence on -3.0 -2.5 FS 253 (15.18) -2.0 Cp,u -1.5 -1.0 -.5 I I I I I I 0 I I I I -1.0 -.8 -.6 -.4 -.2 0 .2 .4 .6 .8 1.0 y/s -3.0 - -2.5 - FS 296 (17.76) -2.0 Cp,u _1.5 - _::_.-.._,,_=_,.Q -1.0 -.5 I i I I 1 I 0 I I I I 0 .2 ,4 .6 .8 1.0 -1,0 -,8 -.6 -.4 -.2 y/s -3.0 - -2.5 - FS 357 (21.42) -2.0 Cp,u _1.5 -I.0 -.5 I I I I I 0 I I 1 I I .2 .4 .6 .8 1.0 -1.0 -.8 -.6 -.4 -.2 0 y/s (d) c_ = 40 °.

Figure 71. Concluded.

O LEX fence off H LEX fence on 2.4 -- 2.4 2.0- 2.0 1.6- 1.6

CL 1.2 - Q

1.2 .8 -- .8 .4 .4 0 I I I I I I 0 I I 1 I I I 0 10 20 30 40 50 60 .4 .8 1.2 1.6 2.0 2.4 _, deg

CD

2.4 2.0 1.6 CL 1.2 .8 .4 -- 0 I I I I I I .2 ,1 0 -.1 -.2 -.3 -.4 Cm Figure 72. Effect of LEX fence on lift, drag, and pitching-moment characteristics at Moc = 0.30 and Ree = 1.40 x 106 .

147'

0 LEXfence off

_

[] LEXfence on

2.4 -- 2.0- 1.6- CL 1.2- ,8 .4 I I I I I I 0__ 10 20 30 40 50 _, deg 2.4 E 2.0- 1.6- EL 1.2 - .8 -- .4 -- .2 Cm Figure 73. Effect of LEX fence on lift, drag, and pitching-moment characteristics at Moc -- 0.60 and Ree = 1.32 x 106.

o LEX fence off LEX fence on [] 2.4 2.4 2.0 2.0- 1.6 1.6- CL CL 1.2 1.2- .8 .8 -- l .4 .4 -- I I 0 I I I I I I 0 I 1 I I 50 60 0 .4 .8 1.2 1.6 2.0 2.4 0 10 20 30 40 CD _, deg 2.4 -- 2.0- 1.6-

/

CL 1.2- .8 --

/

.4 -- 0 I I I 1 I I .2 .1 0 -.1 -.2 -.3 -.4 Cm Figure 74. Effect of LEX fence on lift, drag, and pitching-moment characteristics at M_o = 0.80 and Re_ = 1.02 x 106.

0 LEX fence off [] LEX fence on .O3 .02 .01 Cl 0 -.01 -.02 -.03 __3._ I I I I I I I I I -10 -8 -6 -4 -2 0 2 4 6 8 10 l_, deg .03 - .02 - .01 Cn 0 -.01 -.02 -.03 I I I I I I I I I I -11 -8 -6 -4 -2 0 2 4 6 8 10 13, deg m .2 D-o----_-.....o.____ - /",rl Cy -.2 -.4 .1 I I I I I I I I I -10 -8 -6 -4 -2 0 2 4 6 8 10 15, deg (a) a = 20 °.

Figure 75. Effect of LEX fcnce on rolling-moment, yawing-moment, and side-force coefficient variations with sideslip at M_c = 0.60 and Ree = 1.32 x 106.

o LEX fence off [] LEX fence on

03[-

.01 .021 __O_ Cl -.01 -

o

-.02 - _ -.03 t I I I I I I I I I -10 -8 -6 -4 -2 0 2 4 6 8 10 13, deg .02

°3 t

Cn 0 ,_-- ._ - ----u-_-_._.B::_..___i 3 -.01 - -.02 - -.03 I I I I I I I I t I -10 -8 -6 -4 -2 0 2 4 6 8 10 _, deg .2- Cy -,2 - -.4 , I I I I 1 I 1 I I I -10 -8 -6 -4 -2 0 2 4 6 8 10 _, deg (b) a = 25 °.

Figure 75. Continued.

0 LEX fence off ffl LEX fence on .03 - .02 .01 C1 0 -.01 -.02 - I I I I I I t I I I -.03 -8 -6 -4 -2 0 2 4 6 8 10 -10 [3, deg Cn I I I I I I I I I I -8 -6 -4 -2 0 2 4 6 8 10 9, deg .2- Cy -,2 - -.4 I I I I I I I I I I -10 -8 -6 -4 -2 0 2 4 6 8 10 Figure 75. Continued.

0 LEX fence off [] LEX fence on

03F

.01

Cl 0

-.01 - e__ 0 -.02 - "-D -.03 I I I I I I I I I I -II -8 -6 -4 -2 0 2 4 6 8 10 I], deg .03 - Cn I I I I I I I I I I -8 -6 -4 -2 0 2 4 6 8 10 1_, deg .2- Cy -.2 - -.4 I I I I I I I I I I -10 -8 -6 -4 -2 0 2 4 6 8 10 1_, deg (d) oe= 35 °.

Figure 75. Continued.

O LEX fence off -- [] LEX fence on 5 .02 - .01 _ i .03 _ G =

r "

-.03 [ I I I I I I I I I I i -10 -8 -6 -4 -2 0 2 4 6 8 10 9, deg .03 .02 .01 • Cn 0 - -"--_ -.01 -.02 -.03[ I I I I 1 I I I I I -10 -8 -6 -4 -2 0 2 4 6 8 10 [3, deg Cy 0 ___ -.2_ -.4 1 I I I I I I I I I . -10 -8 -6 -4 -2 0 2 4 6 8 10 [3, deg N (e) a = 40 °.

Figure 75. Concluded.

i

ORIGINAL PA':3 E BLACK AND WHITE PHOTOGRAHh (a) Nose boom on.

t (b) Nose boom off.

Figure 76. Laser vapor screen flow visualizations with nose boom on and off at ]Y/o_ = 0.60, Ree = 1.32 x 106, = 50 °, and FS 184 (11.04). Dimensions are in inches full scale (0.06 scale).

ORIGINAL PAGE Bt.AE;K AND WHIT[. PHOTOGRAPh (a) a = 30°; FS 411 (24.66).

=== i ==< (b) c_ = 32.5°; FS 357 (21.42).

Figure 77. LEX vortex breakdown asymmetries at it{oo = 0.60 and Rea = 1.32 x 106 with nose boom and LEX fences on. Dimensions are in inches full scale (0.06 scale).

o Nose boom off 2.2. [] Nose boom on ': _-_ -2.0 - FS 107 (6.42) -1.5 -- 180 ° 270 ° 90 _

CO -,5 -

-1.0 -- /

o

0 °

i i i

0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg -2.0 - FS 142 (8.52) -1.5 - 180 ° -1.0

Co -.5

90 ° 270 ° 0 o .5 1.0 t I I I I I 1 l 1 1 I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg -2.0 FS 184 (11.04) 180 ° -1.5 -1.0

Co

90°_ 270° OO 1.0_ I I I I I I I I I I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 . -, .

0, deg Figure 78. Effect of nose boom on forebody surface static pressures at AI_c = 0.60; R% = 1.32 x 10 (_, and = 50 °. Dimensions are in inches full scale (0.06 scale).

O Nose boom off

t3 Nose boom on

-2.0

FS 107 (6.42) 180 °

-1.5

-1.0

90 ° 270o

Cp -.5

0 °

.5

1.0 I I I I I 1 1 1 I I I I

0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg -2.0 FS 142 (8.52) 180 ° -1.5 -1.0 C__.p -.5 90°_ 270° o .5 1.0 1 I I I I I I I I I I I 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg -2.0 - FS 184 (11.04) 180 ° -1.5 -1.0

% -.5

90°_ 270° .5 [ 1.0 I 1 I I I I I 1 I 1 I 0 30 60 90 120 150 180 210 240 270 300 330 360 0, deg Figure 79. Effect of nose boom on forebody surface static pressures at M_c = 0.80; Ree = 1.02 x 106, and c_ = 40 °. Dimensions are in inches full scale (0.06 scale).

O Nose boom off [] Nose boom on 2.4 m 2.4 m 2.0- 2.0- 1.6- 1.6 CL 1.2 - CL 1.2 .8 -- .8 .4 -- .4 -- I I I I I I 0 I I I I I I 0 10 20 30 40 50 60 0 .4 .8 1.2 1.6 2.0 2.4 a, deg co 2.4 -- 2.0 - 1.6- CL 1.2- .8 -- .4 -- I I I I I I .2 .1 0 -.1 -.2 -.3 -.4 Cm Figure 80. Effect of nose boom on lift, drag, and pitching-moment characteristics at M_ = 0.60 and Ree = 1.32 × 106 .

= | 0 Nose boom off | [] Nose boom on ,_

|

2.4 - 2.4 E i i 2.0 2.0 1.6 1.6 i CL 1.2 CL 1.2 !

i .8 .8 0 I I 0 1 I I ii .4 .4 10 20a, 3_eg 40 50 60 0 .4 .8 1.2CD 1.6 2.0 2.4 i 2.4 2.0 1.6 CL 1.2 .8 .4 I I I I I I .2 .1 0 -.1 -.2 -.3 -.4 Cm Figure 81. Effect of nose boom on lift, drag, and pitching-moment characteristics at M_c = 0.80 and Roe = 1.02 x 106.

m m o Nose boom off Nose boom on [] .03 .02 .01 Cl 0 -.01 w -.02 -.03 I I I I I I I I I I I I 0 5 10 15 20 25 30 35 40 45 50 55 60 a, deg .03 .02 .01 Ca - - v V -_.t:l u--_ -.01 m -.02 -.03 I I I I I I I I I I I I 0 5 10 15 20 25 30 35 40 45 50 55 60 a, deg .2 Cy -,2 °°4 1 I 1 I I I t I I I I I 0 5 10 15 20 25 30 35 40 45 50 55 60 _, deg Figure 82. Effect of nose boom on lateral-directional characteristics at Moo = 0.60, Re_ -- 1.32 x 106, and /?=0% Form Approved REPORT DOCUMENTATION PAGE OMB No. 0704-0188 PuNic reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the data needed, and completing and reviewing the collection of information Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project {0704-0188), Washington, DC 20503 1. AGENCY USE ONLY(Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED December 1991 Technical Paper 4. TITLE AND SUBTITLE 5. FUNDING NUMBERS Wind Tunnel Investigation of Vortex Flows on F/A-18 at Subsonic Through Transonic Speeds WU 505-68-30-03 6. AUTHOR(S) Gary E, Erickson 8. PERFORMING ORGANIZATION 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) REPORT NUMBER NASA Langley Research Center ttampton, VA 23665-5225 L-16799 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORING/MONITORING AGENCY REPORT NUMBER National Aeronautics and Space Administration NASA TP-311I Washington, DC 20546-0001 11. SUPPLEMENTARY NOTES 12b. DISTRIBUTION CODE 12a. DISTRIBUTION/AVAILABILITY STATEMENT Unclassified Unlimited Subject Category 02 13. ABSTRACT (Maximum 200 words) k wind tunnel experiment was conducted in the 7- by 10-Foot Transonic Tunnel at the David Taylor Research Center of tile wing leading-edge extension (LEX) and forebody vortex flows at suhsonic and transonic speeds about a 0.06-scale model of the F/A-18. The primary goal was to improve the understanding and control of the vortical flows, including the phenomena of vortex breakdown and vortex interactions with the vertical tails.

The wind tunnel results were correlated with in-flight flow visualization and handling qualities trends obtained by NASA using an F-18 High-AIpha Research Vehicle (HARV) and by the U.S. Navy and McDonnell Douglas Corp. on an F-18 airplane with LEX fences added to improve the vertical tail buffet environment. Key issues that were addressed include the sensitivity of the vortical flows to the Reynolds number and Math number; the reduced vertical tail excitation, and the corresponding flow mechanism, in the presence of the LEX fence; the repeatability of data obtained during high-angle-of-attack wind tmmel testing of F/A-18 models; the effect of particle seeding for flow visualization on the quantitative model measurements; and the interpretation of off-body flow visualizations obtained with different illumination and particle seeding techniques.

14. SUBJECT TERMS 15. NUMBER OF PAGES Vortex flows; Subsonic flow; Transonic flow; Fighter aircraft; Forebody; Vertical tail 164 buffet; Vortex breakdown; Flow visualization; Leading-edge extension 16. PRICE CODE A08 17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATIOI_ 20. LIMITATION OF REPORT OF THIS PAGE OF ABSTRACT OF ABSTRACT Unclassified Unclassified _ISN 7540-01-280-5500 tandard Form 2gS(Rev. 2-89) Prescribed by ANSI Std Z39-18 298-102 NASA-Langley, 1991 ml Es E I !

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