Navier-Stokes, flight, and wind tunnel flow analysis for the F/A-18 aircraft

NASA Technical Paper 3478

Navier-Stokes, Flight, and Wind Tunnel Flow

Analysis for the F/A-18 Aircraft

Farhad Ghaffari Langley Research Center • Hampton, Virginia National Aeronautics and Space Administration Langley Research Center • Hampton, Virginia 23681-0001

December 1994

The use of trademarks or names of manufacturers in this report is for accurate reporting and does not constitute an official endorsement, either expressed or implied, of such products or manufacturers by the National Aeronautics and Space Administration.

Acknowledgments Robert M. Hall of the Langley Research Center assisted with the wind tunnel testing of the F/A-18 aircraft configuration.

James M. Luckring and James L. Thomas of the Langley Research Center, Robert T. Biedron of Analytical Services and Materials, Inc., David F. Fisher of the Dryden Flight Research Center, and Brent L. Bates of ViGYAN, Inc., contributed to various aspects of this investigation.

This publication is available from the following sources: National Technical Information Service (NTIS) NASA Center for AeroSpace Information 5285 Port Royal Road 800 Elkridge Landing Road Springfield, VA 22161-2171 Linthicum Heights, MD 21090-2934 (703) 487-4650 (301) 621-0390 -1 :| i, Contents Summary .................................. i o,o III 1 1-1 Introduction Summary Combat aircraft arc often asscsscd on their high- Computational analysis of flow over the F/A-18 angle-of-attack aerodynamic performance for achiev- aircraft is presented along with complementary data ing superior levels of sustained maneuverability and from both flight and wind tunnel experiments. The agility. At high attitude, the flow characteristics over computational results are based on the threc- these often geometrically complex aircraft configura- (timensional thin-layer Navier-Stokcs formulation tions generally become very complicated and (tiffi- and arc obtained from an accurate numerical model cult to predict and control. One such flow charac- of the fuselage, the leading-edge extension (LEX), teristic is the inevitable vortical flow precipitated by and the wing geometry. However, tile constraints flow separation that occurs when the aircraft oper- imposed by the flow solver and/or the complex- ate at high angles of attack. In general, the pres- ity associated with the flow-field grid generation ence of such vortical flow over an aircraft surface can required certain geometrical approximations to be be advantageous as long as it remains organized and implemented in the present numerical model. In par- stable; this flow produces vortex lift, which can en- ticular, such constraints from tile flow solver inspired hance maneuverability, ttowever, with increasing an- the blocking (fairing) of the inlet face, wifich then gle of attack, such a coherent vortex system is sus- precluded the propulsion effects. The grid generation ceptible to instabilities such as breakdown or flow complexity required the removal of the empennage.

asymmetry, which cause undesirable pitching, yaw- The results are computed for three different ing, and/or rolling moment characteristics. As a free-stream flow conditions and compared with flight result, the flight-handling quality and controllability test data for surface pressure coefficients, surface tuft of these aircraft are adversely affected by such flow flow, and off-surface vortical flow characteristics that phenomena; the ability of the aircraft, to maneuver included breakdown phenomena. Excellent surface with high agility is often limited. The fundamental pressure correlations, both in terms of magnitude and understanding and predictability of flow phenomena overall trend, are obtained on the forebody through- for a wide range of flow" conditions are of paramount out tile examined range of flow conditions. Reason- interest from both research and real aircraft design able pressure agreement w_ found over the LEX; the perspectives.

general correlation tends to improve at higher angles The vortical flow formation or, in a more gen- of attack. The surface tuft flow and the off-surface eral form, the initial flow separation can bc classified vortex flow structures compared qualitatively well into two types. The first type is a flow separation with the flight test results.

that primarily occurs over a smooth surface geome- To evaluate the computational results, a wind try because of an adverse pressure gradient interact- ing with the boumtary layer. Fhfid viscosity provides tunnel investigation was conducted to determine the aerodynamic effects of existing configurational dif- the essential mechanism for this type of flow sepa- ferences between the flight vehicle and the numerical ration to occur; this suggests that its formation is model. This study revealed that in most cases, the highly sensitive to the local flow" Reynolds number.

geometrical approximations made to the numerical A typical flow separation of this type occurs over model had very little effect on overall aerodynamic a conic forebody at high angles of attack. Unlike characteristics. Furthermore, to validate the latter the first, the second type of flow separation occurs at, and is primarily induced by, a surfaee discontinu- wind tunnel results at flight flow conditions, a com- putational study was conducted to determine the ity such as the sharp leading edges of a delta wing.

aerodynamic influence of differcnces in the Reynolds Because of diminished sensitivity to the fluid viscos- number. This computational study, which was per- ity, the latter type of flow separation, fixed at the surface discontinuity, is generally considered to be formed on exactly the same grid, showed that an order-of-magnitude difference between the flight and insensitive to the local flow Reynolds number. In wind tunnel Reynolds numbcrs produces negligible this paper, the first and second type of flow sep- effects on the forcbody and the LEX surface pres- aration is referred to as boun(tary layer and sharp sures as well as the longitudinal aerodynamic char- edge flow separation, respectively. In recent years, acteristics. Vcry good surface pressure correlation various numerical (refs. 1 and 2) as well as experi- between wind tunnel and flight data was obtained mental (refs. 3 6) research efforts have been made to on the LEX; however, the wind tunnel pressure data quantify the effects of Reynolds number on different appear to be slightly below thosc of the flight mca- types of flow separation. The results from these stud- surement on the forebody, particularly at high angles ies are particularly important in providing insight of attack.

into the flow physics and the triggering mechanisms results with both the multiblock overset (refs. 21 23)

responsible for the subjectflowseparations. Specifi-

and nonoverlapping grids (ref. 24) are documented.

cally,thebasicknowledge learned fromtheseinvesti-

gations mayleadto the development of newcompu-

The primary objectives of the present investiga-

tational and/or experimental techniques applicable

tion are summarized into the following five categories:

to a conventional wind tunnelenvironment for sim-

1. Expand prior thin-layer Navicr-Stokes computa-

ulatingthe high Reynolds number flowencountered

tions (ref. 24) for the F/A-18 aircraft configu- by the flight vehicle.

ration to include a wider range of flight flow conditions

Uniqueresearch is presently beingconducted by

NASAwittlin tile High-Angle-of-Attack Technology 2. Correlate computational results with flight test data for both on- and off-surface flow character-

Program (HATP)(ref.7), whichhasasits majorob-

istics and surface pressure coefficients

jectivethe exploration of high-attitudeflowcharac-

teristics overa typicalfighteraircraftduringmaneu-

3. Evaluate aerodynamic effects which result from

veringsituations.At the outset,the F/A-18aircraft

the geometrical differences between the flight

waschosen to bc the baseline configuration primar-

vehicle and the numerical model through wind

ily because of its high-angle-of-attack (i.e.,high-a)

tunnel experimentation

capability.Subsequently, anF/A-18 aircraft,desig-

4. Correlate wind tunnel data with flight test re-

natedas ttle High-AlphaResearch Vehicle(HARV)

sults to assess the scale model simulation of high

(ref.8),washighlyinstrumented forsurface pressure

Reynolds number flow; in addition, numerically measurements aswell as in-fligtlt flowvisualization.

assess the effects of Reynolds number on the aero-

(Seerefs.9 and 10.) On- andoff-surface flowvisu-

dynamic characteristics of the configuration

alizations havebeen conducted oil tile HARV;differ-

ent techniques havebeenusedandincludeoneinno-

5. Correlate computational results with appropriate vativeapproaci_ to document surface flow patterns.

wind tunnel data obtained on a configuration that

(Secref. 9.) In additionto the flight experimen-

is more representative of the numerical model

tation, HATP is utilizing ground-based facilitiesto

Symbols

acquire experimental datafromvarious wind tunnel

scalemodels(refs.11 15)aswell as a full-sizecon-

bref reference wing span, 37.42 ft

figuration, whichtinsbeen testedin the Ames 80-by

Drag

120-Foot Wind Tunnel.(Seercf. 16.) Theresultsof

CD drag coefficient, q_cSre f

these experiments have provided thedatabase needed

Lift for tim development and validation of the present computational fluid dynamics (CFD) methodologies. lift coefficient, qocSref pitching moment coefficient In the past few years, significant progress has Pitching moment been made to fulfill the HATP objectives by pro- referenced to 0.25_, q_cSrefC viding detailed analyses of tfigh-angle-of-attack flow over tim F/A-18 aircraft. In particular, the numer- pressure coefficient, p - P_

G

ical analyses based on the thin-layer Navier-Stokes qoc formulation have made important contributions. Ini- wing mean aerodynamic chord, tial computational activities started in parallel at the 11.52 ft Langley and Ames Researctl Centers. Two different center of gravity c.g, approaches were taken to solve the flow over the iso-

lated F/A-18 aircraft forebody which included the Eo total energy per unit volume

leading-edge extension (LEX) geometry. Although flux vectors F,G,H both approaches were based on multiblock structured grid strategies, one method used a nonoverlapping net flux grid-block approach (refs. 17 and 18), and the other Jacobian of coordinate transformation J mett_od was based on an overset grid or Chimera l (ref. 19) approach. (See ref. 20.) Subsequently, both configuration body length measured from nose to exhaust nozzle exit computational methods were expanded to include plane, full scale, 54.4 ft more of the F/A-18 aircraft geometrical components such as the wing, aft portion of the fuselage, and em- Mach number pennage (i.e., vertical and horizontal tails). These free-stream Mach number efforts were very successful and the computational | |11 FS fuselage station, full-scale, in.

static pressure P HARV High-Alpha Research Vehicle free-stream total pressure po HATP High-Angle-of-Attack Technology local total pressure Po,I Program normalized total pressure Po,l/Po H-H grid topology, H streamwisc and circumferential free-stream static pressure Poc H-O grid topology, H streamwise and O

Q state vector, j-1 [p, pu, pv, pw, Eo] r

circumferential free-stream dynamic pressure, lb/ft 2 qoc HST High-Speed Tunnel Re Reynolds number based on e IGES Initial Graphics Exchange Specification reference area of wing planform, Srd 400 ft 2 LEX leading-edge extension MUSCL monotone upstream-centered 8 LEX local-exposed semispan, in.

scheme for conservation laws t time, see NAS numerical aerodynamic simulation body-axis Cartesian velocity "_l , 73, "03 NASA National Aeronautics and Space components, ft/sec Administration V* WT wind tunnel wall friction velocity, V_P" ft/sec A caret (^) over a symbol indicates scaling X axial distance from nose apex, in.

with respect to the Jacobian J of the coordinate transformation.

x/1 fraction of configuration body length Sources of Data distance along LEX local-exposed Y scmispan, in. Flight Experiment y/s fraction of LEX-exposed scmispan The F/A-18 aircraft was chosen as a baseline con- figuration for the HATP, primarily because of its inner-law variable, _ y+ v high-a capabilities. The aircraft (fig. 1), designated tile High-Alpha Research Vehicle (HARV), is instru- O_ angle of attack, deg mented to measure surface static pressures over the flap deflection angle, deg 5f forebody and on both starboard and port sides of the LEX. The three views of the F/A-18 aircraft config- viscosity, lb-sec/ff 2 uration are shown in figure 2. The figure also pro- vides the full-scale reference dimensions in feet. Fig- Y kinematic viscosity, _, ft2/sec ure 3 presents the plalfform of an F/A-18 aircraft azimuthal angle measured clockwise configuration and the corresponding cross-sectional viewed from front, dcg (0 ° located geometry of the forebody and LEX fuselage stations at bottom dead center) (FS) where tile surface pressures are measured. The forebody surface pressure orifices were distributed as (,71, ¢ body-fitted coordinates a function of 0 at longitudinal FS 70, 85, 107, 142, and 184. These fifll-scale dimensions are in inches; for density, slug/ft 3 P reference, the nose apex starts at FS _ 60. When the wall shear stress, lb/ft 2 T_ L, cross section is viewed by looking aft, tile azinmthal angle 0 is measured clockwise from the windward Abbreviations: plane of symmetry of 0 = 0°; the LEX surface pres- CAD computer-aided design sures were measured as a function of LEX-exposed semispan y/s on both the upper and lower surfaces CFD computational fluid dynamics at FS 253, 296, and 357. The semispan parameter C-O grid topology, C streamwise and O y/s = 0 corresponds to the LEX-fuselage juncture circumferential and y/s = 1 corresponds to the LEX leading edge.

tailed computer-aided design (CAD) description in a

Negative values of the parameter y/s correspond to

format known as Initial Graphics Exchange Specifi- the starboard side and positive values to the port side cation (IGES). (See ref. 25.) These data were then of the configuration.

used to extract a high-density surface grid definition The aircraft is also equipped with a smoke- in thc form of cross sections. The CFD database grid generating system (refs. 9 and 10) designed to emit consisted of approximately 30 000 points on the fuse- smoke at appropriate locations along the forcbody lage defined at 60 cross sections and 16 000 points on and the LEX apex for visualization of the off-surface the wing defined at 20 streamwisc cuts. Although vortical flows as well as their interactions with one not used in the present computations, accurate sur- another and/or with the neighboring aerodynamic face definitions of both horizontal and vertical tails surfaces. In conjunction with the off-surface flow were also included in this database.

visualization, surface tufts are used on the wing, This database was subsequently used to generate LEX, fuselage, and tails to assist in correlation of a suitable surface grid for Navier-Stokcs computa- the off-surface flow interactions with the surface flow.

tions by using an Ovcrhauser function (ref. 26) for These in-flight flow visualization images arc recorded the interpolation. This function has been shown to by a camera located either onboard the aircraft or a alleviate the oscillatory behavior inherent in other chase plane. Furthermore, a unique approach has widely used functions (e.g., splines) in the region also been successfully used to capture the in-flight where grid point distributions are not uniform. The surface flow pattern on the forcbody and the LEX.

final computational surface grid was composed of ap- (See ref. 9.) The data gathered from the flight exper- proximately 18 000 points. The geometrical simplifi- iments have been instrumental in helping researchers cations made to the configuration included the fair- understand the subject flow phenomena. All the ing over of the inlet, splitter plate, diverter, and the HARV flight experiments were conducted by NASA LEX slot. Figure 4(a) shows a body cross section at the Dryden Flight Research Center. The flight at FS 401 and illustrates the simplifications made data arc obtained for a wide range of angle of attack, to the splitter plate and the diverter cavity region.

Mach number, and sideslip. (See ref. 10.)

Similarly, figure 4(b) shows a body cross section at FS 441 and illustrates the closing of the LEX slot as Computational Fluid Dynamics well as the fairing of the cavity region between the inlet and the LEX lower surface. The latter simpli- The primary objective of the present compu- fication is made to a limited region to facilitate the tational analysis is to expand the earlier compu- flow-field grid generation in that area. Except for tations (ref. 24) to include a wider range of flow the simplified regions, the two typical cross sections conditions and a more comprehensive flow analy- shown in figures 4(a) and 4(b) demonstrate the accu- sis. Various HARV flight test results (e.g., on- and racy with which the computational grid (_100 grid off-surface flow visualization photographs, surface points/station) represents the surface geometry of pressure data) were initially examined to identify the much finer initial CAD cross-sectional defini- those flow conditions which exhibited the most chal- tion (_500 grid points/station) despite the use of lenging flow characteristics to be simulated numeri- only about a fifth as many grid points. Two ortho- cally. Several flout conditions (table I) werc identified graphic views of the final F/A-18 aircraft CFD sur- which clearly demonstrate the complexity associated face grid definition are shown in figures 5(a) and 5(b) with the overall flow structures (e.g., LEX vortices to illustrate the overall grid resolution. Further- with subsequent breakdown at high-_ conditions, more, the wing geometry is modeled with two dif- forebody vortices, forebody-LEX vortex interactions, ferent leading-edge flaps: the undeflccted and the and/or stalled flow over the wing).

blended flap. The latter will be discussed in the next paragraph.

Table I. Selected Flow Conditions The F/A-18 aircraft wing leading-edge flap de- flection angle varies as a function of angle of attack _], dog and Mach number. For free-stream subsonic Mach

! 034 / la5 ×10°

numbers (Moc _< 0.76), the aircraft control system

5.s I / 10.s

is programmed to vary the flap deflection angle lin- 10.2 early as a function of angle of attack according to the relationship _1- = 34_/25.6. The maximum flap Surface grid. The surface patch definition of the deflection angle of 34 ° is reached when c_ = 25.6 ° complete F/A-18 aircraft was obtained from a de- and the flap angle remains constant for a _> 25.6 °.

7 :| i aries with the wing wedge-shaped region sectioned

As a result, the flap deflection angles, which

out of the field domain. Again, the boundaries of

correspond to the flow conditions(tableI) of this

the various regions and the corresponding block in-

investigation are as follow: @ = 25 ° for c_ = 19 °

terfaces are highlighted with thick, solid and dashed and af = 34 ° for c_ = 25.8 ° and 30.3 ° . To sim- lines, respectively. Three factors contributed to the plify the flow-field grid generation, the surface ge- selection of the grid topology for each region: the ometry of the wing deflected leading-edge flap was consideration of local geometry, the proper resolu- approximated in this computation and is designated as a blended flap. The principle behind the deflected tion of the expected flow structure, and the ap- flap geometry modification was the preservation of propriate grid connectivity between regions. The the wing-body intersection with the undeflected flap, selected topologies should provide good resolution which permitted the same overall blocking strategy (refs. 17 and 18) of all edge flows (e.g., LEX and to be used for both undeflected and blended flap con- wing leading edges and wingtip) and juncture flows (e.g., LEX-body, wing-body, and canopy-body).

figurations. This modification smoothly blended the inboard 15-percent semispan of the flap between the The volume grid is generated with established deflected flap and the undeflected flap wing-body in- transfinite interpolation techniques (refs. 17, 24, tersect.ion. A nose-down front view of the F/A-18 and 27) with sufficient normal clustering near the aircraft CFD surface grid definition with both the surface to adequately resolve the laminar sublayer of blended flap (starboard) and undeflected flap (port) the turbulent-boundary-layer flow for a typical flight is shown in figure 6. In addition, a closeup view of the free-stream condition of c_ = 19 °, Moc = 0.34, and CFD surface definition is shown in figure 7 with shad- Rc = 13.5 x 10 G. This grid produced an average ing to highlight tile surfaces of the baseline F/A-18 normal cell size next to the wall of approximately aircraft that are simplified, blended, and/or modified 7.2 × 10-Gc, which corresponds to 9 + _. 3 for turbu- (e.g., inlet, diverter, splitter plate, and the inboard lent computations; a laminar sublayer generally ex- section of the deflected wing leading-edge flap).

tends out to y+ .-_ 8.5. The radial far-field boundary extends to about 7.6c. A downstream grid extension Flow-field 9rid. Tire selection of the flow-field is created by repeating the grids generated about the grid strategy is primarily dictated by the two dis- base cross section aft to about 4.7c. No grid is gener- tinct types of aerodynamic shapes of the F/A-18 ated for the face of the base geometry (i.e., open), nor aircraft configuration: a slender type, which con- is the flow simulated between the interior surfaces of sists of the front forebody-LEX geometry, and a the model geometry and the exterior flow-field do- high-aspect-ratio type, which contains tile wing com- main. Note that the flow-field grid structure is gen- ponent. An H-O grid topology is selected for the erally designed to be consistent with those of previ- slender part, whereas a C-O grid is chosen for ous computational studies (refs. 17 and 28) on the the high-aspect-ratio wing configuration. A unique isolated F/A-18 aircraft, forebody-LEX configuration global grid strategy is then devised which appro- where tile structures had been found to have ade- priately links various grid topologies while main- quate cell size next to the wall, radial grid stretching, taining tile grid quality. To illustrate the selected circumferential grid resolution, and far-field bomM- global grid strategy, isometric far-field (fig. 8(a)) and ary locations.

near-field (fig. 8(b)) views of the shaded F/A-18 air- craft, surface are shown for the configuration maxi- Computational methodology. The computational mum half-breadth plane along with the field grids in results have been obtained from an algorithm that the plane of symmetry. For clarity, the grid density has been successfully applied to a variety of aero- shown in tile figures has been reduced in both longi- dynamic problems with both simple and complex tudinal and radial directions. The flow-field domain, configurations for a wide range of flow conditions.

which consists of about 1.24 million grid points, is (See refs. 2, 17, 18, 24, and 27.) The algo- divided into five regions with each composed of one rithm (refs. 1, 17, 18, and 29 31) is based on tire or more topologically similar blocks. Ill figures 8(a) compressible, time-dependent, Reynolds-averaged, and 8(b), the region boundary edges arc highlighted Navier-Stokes equations, which are written in a with thick, solid lines and the corresponding block curvilinear coordinate system. The equations are interfaces within each region are denoted by thick, solved with a finite volume approach and are com- dashed lines.

posed in a conservative form as A side view of the flow-field grid for selected sur-

+ @ - + (G - + (fi - = 0

faces is shown in figure 9 to illustrate the various The subscripts with a comma denote partial regions from a different perspective. The figure de- picts tile overall three-dimensional far-field bound- differentiation, the subscript v identifies the viscous and to reduce oscillations in CL to a negligible level

terms,and the caret (") overthe vectorsindicates

as shown in figures 10(a) and 10(b), respectively.

scalingwith respectto the Jacobian J of the coor-

dinate transformation. Details of these terms are in- Subsequent solutions for different angles of attack eluded in reference 17. In addition to the ideal gas were obtained by starting the computations from an assumption in the present study, tile thin-layer ap- existing solution, which then generally reduced the proximation of the governing equations is invoked computational time for a converged result by as much as a half. Similar convergence rates are achieved for (i.e., Fv = (_v = 0) and thus accounts for viscous the computations at higher angles of attack despite flux terms only in the direction _ normal to the body.

the presence of vortex breakdown in the solutions.

Turbulence effects are accounted for by the notion The computations are performed without the use of of eddy viscosity and conductivity. The algebraic mesh sequencing or multigrid iteration. (See ref. 29.)

turbulence model developed by Baldwin and Lomax (ref. 32) is used to evaluate the required turbulence Wind Tunnel Experiment quantities. For separated vortical flow regions, the method introduced by Degani and Schiff (ref. 33) Tile wind tunnel experiment was conducted in the is used to ensure that the proper turbulence length scales are used.

Langley 7- by 10-Foot High-Speed Tunnel (HST).

(See refs. 36 and 37.) This is a closed-circuit, The integral form of the conservation equations continuous-flow, atmospheric tunnel with a solid wall is represented by test section 6.6 ft high, 9.6 ft wide, and 10 ft long.

The tunnel h,'_s an operational Mach number range of 0 to 0.9 with a maximum Reynolds number of about N 4 x 10 6 ft -].

Tile wind tunnel testing was conducted with a where the time rate of change Ot of the state vec- 0.06-scale model of the F/A-18 aircraft configuration.

tor I_ within a cell volume dV is balanced by This wind tunnel model had been used in a previous the net flux f across the cell surface do e with the experimental investigation, and the results arc pub- unit normal ft. The convective and pressure flux lished in reference 12. The model was instrumented quantities are represented by the upwind-biased, for surface static pressure measurements at four sta- flux-difference-splitting approach of Roe (ref. 34), tions on the forebody and three stations on the LEX whereas tile shear stress and heat transfer terms upper surface; LEX lower surface measurcments were are centrally differenced. Tile monotone upstream- taken only at the last station FS 357. The surface centered scheme for conservation laws (MUSCL) of pressures were mcasured on both starboard and port Vail Leer (ref. 35) is used to interpolate state vari- sides of the aircraft to assess flow asymmetry. The ables at the cell interfaces. A detailed discussion fuselage stations on the model at which the surface on the algorithm development for interpolating the pressures arc measured are identical to those of the mass, momentum, and energy across tile various flight vehicle with the exception of the first forebody planar and nonplanar interfaces that separate the station FS 70 where no model data were acquired.

grid blocks is presented by Biedron and Thomas in The sting-mounted wind tunnel model (fig. 11) was reference 31.

equipped with an internally mounted strain gauge balance to me_ure the six component forces and mo- Method performance and convergence. All com- ments. Furthermore, the forebody of the model was putations are performed on the numerical aero- equipped with two transition grit strips positioned dynamic simulation (NAS) Cray-2* computer, lo- longitudinally across the windward plane of symme- cated at Ames Research Center. On this machine, try at 0 = 45 ° and 315 °. Based on the method of ref- the algorithm requires approximately 20 #see per erence 38, the No. 180 grit was found to be adequate iteration per grid point and about 100 million words for tripping the laminar boundary layer to a turbu- of memory. Starting from the free-stream flow condi- lent flow to simulate the flight Reynolds number flow tion, a typical solution converged in about 3000 itera- characteristics at the conditions listed in table I.

tions, which consumed about 20 hr of computer time.

The 3000 iterations were sufficient to reduce the Two different configurations of the model were residuals by a little more than 2 orders of magnitude tested: the first was the baseline F/A-18 aircraft and the second incorporated modifications representative of the numerical model. The second configuration is referred to as the "CFD wind tunnel model" (CFD *Trademark of Cray Research, Inc., Minneapolis, MN 55402. WT) from here on. The data obtained from the :1 Ili CFD wind tunnel model are used to assess the aero- spond to the free-stream static pressure Poc. Surface dynamic effects of empennage removal, inlet fairing, pressure measurements are obtained with electroni- and wing leading-edge flap deflection. Modifications cally scanned pressure (ESP) transducers; the over- of the baseline wind tunnel model were patterned af- all accuracy of this system is about +0.1 percent of the full-load range, which is approximately equal to ter the numerical representation. Dental plaster was 4-0.03 lb/in 2.

used for fairing over regions of the model such as the inlet and diverter. Figures 12(a) and 12(5) show Results and Discussion the CFD wind tunnel model from two perspectives and illustrate the various modifications of the base- CFD Versus Flight Data line F/A-18 aircraft wind tunnel model such as fair- ing over the inlet, splitter plate, and diverter; closing General flow features. The normalized total pres- of the gap between the deflected flap and fuselage; sure Po,1/Po contours in various cross-flow planes as and the removal of the empennage.

well as the LEX and forebody vortex core stream- The CFD wind tunnel model was a good represcn- lines (where applicable) computed at the selected tation of the numerical model except in the regions flow conditions (table I) are shown in figures 13(a) of the blended flap and the splitter plate. Unlike the 13(c). The magnitude associated with each normal- numerical model, where surface modifications were ized total pressure contour is displayed with the ap- made to the inboard 15-percent scmispan of the flap propriate color bar. The normalized total pressure by a smooth blending of the deflected flap geometry function is used to highlight the viscous losses within to the undeflected flap wing-body intersection, the a separated flow structure such as a vortex. For tile CFD wind tunnel modcl did not incorporate flap sur- same purpose, this function has also been success- face blending. However, the gap between the inboard fully used in an experimental investigation reported face of the deflected flap and the fuselage was closed in reference 41. The results shown in figures 13(a ) with a metal sheet. (fig. 12(b)) that vertically joined 13(c) are all obtained with a fully turbulent bound- the two edges. This gap renmined closed for the de- ary layer model at flight flow conditions with the flected flap configurations of the present CFD wind blende(t flap configuration. The computations arc tunnel model. Furthermore, dental plaster was also performed for half the configuration, but the results used to fair over the cavity region between the splitter are presented for the fifll configuration by using the plate and the fuselage but in a slightly different nmn- mirror-image principle of symmetry. Although both ner than the numerical model approach (fig. 4(a)) in Mach and Reynolds numbers vary slightly, the com- which the lower part of the splitter plate was trun- putational results presented in figures 13(a) 13(c) re- cated. This difference between the numerical model veal tile effects of angle of attack on the flow charac- and the CFD wind tunnel model in the geometric teristics. These figures, discussed in the subsequent representation of the splitter plate and diverter cav- paragraphs, highlight the following three general flow ity regions is illustrated for a typical cross section at.

features and their interactions: LEX w)rtex system, FS 401 in the lower right corner of figure 12(b).

wing flow field, and forebody riot" field.

The measured wind tunnel data are corrected The normalized total pressure contours in fig- for the effects of angle of attack, wall interference, ures 13(a) 13(c) clearly indicate the presence of a and model base. drag. The model support system well-organized LEX vortex riot" structure up to the incorporated an a.ccelerometer to measure the an- LEX-wing leading-edge juncture. Over this longitu- gle of attack and is subsequently corrected to ac- dinal extent, the overall LEX vortical riot, structure count for tile balance and sting deflection under load. generally remains similar even as the angle of attack The wall interference effects are accounted for by is increased. At c_ = 19 °, the LEX primary vor- tex system appears to remain coherent and maintain the principles of 51ockage (ref. 39) and jet bound- ary (ref. 40) corrections. The model base pressures its tight core structure over the entire configuration are measured and subsequently integrated to obtain body length. However, with increasing angle of at- tile resulting force acting normal to the base plane tack, figures 13(b) and 13(c) illustrate that the LEX of the model. This normal force is then subtracted vortex core region (highlighted by" the lower levels of from the total axial force component measured by the normalized total pressure, which signify higher the internally mounted strain gauge balance to ex- levels of viscous loss) expands dramatically aft of clude the pressure drag caused by the local wake about the wing root midchord. This sudden core flow on the base of the model. As a result, the expansion in the LEX vortical flow system is gen- model base pressure drag contril)ution to the configu- erally associated with a phenomenon referred to as vortex burst or breakdown. The vortex breakdown is ration total forces and moments is adjusted to corm- because of a massive flow separation and the LEX

oftencharacterized by an abruptreductionin veloc-

vortex breakdown, respectively.

ity (particularlytheaxialcomponent) andtile lossof

cohesiveness within the vorticalflowstructure.The

Computational results for LEX vortex core

latter effectis clearlydemonstrated by the LEX vor-

streamlines superimposed on surface tuft flow pat-

tex corestreamlines at c, = 25.8 ° and 30.3 °. The

terns are presented in figures 15(a) 15(e) for the flow predicted location of the LEX vortex breakdown is conditions listed in table I. For the higher angles of discussed later in conjunction with flight and wind attack (i.e., a = 25.8 ° and 30.3°), the forebody vor- tunnel test results.

tex core streamlines are also shown to highlight their paths and influence on the overall flow structures The normalized total pressure contours at both on and off the surface. Surface tuft flow pat- c_ = 19 ° clearly illustrate the separated flow region terns are simulated computationally with tile method over the wing upper surface. This massive flow sep- of unrestricted streamline tracing introduced and dis- aration over tile wing is essentially a confined region cussed in detail in reference 24. Unlike tile conven- of retarded airflow with a chaotic behavior which is tional method of tracing the experimental surface oil discussed in reference 24. When the angles of at- flows and tuft patterns, this new approach does not tack are increased to 25.8 ° and 30.3 °, the separated impose the restriction that tile streamline calcula- flow region over the wing appears to move outboard and exhibits lower levels of viscous toss as indicated tions lie in a particular grid plane near the surface.

The method (ref. 24) has demonstrated the capability by the higher levels of the normalized total pressure.

One contributor to this flow change is tile expansion of simulating surface flow patterns in regions of at- tached as well as separated flows and is particularly of the LEX vortex flow which extends spanwise onto applicable to a stalled flow environment. Because tile wing with increasing angle of attack.

of the stalled flow characteristics over the wing, the With increasing angle of attack, the flow within method applied here initiated tile particle tracing at the boundary layer over the smooth leeward side of a grid plane slightly off the surface (_0.02 in. full the forebody separates and leads to a well-organized scale, which is _0.00014_:) where the flow velocity vortical flow structure as shown in figures 13(b) magnitudes become sufficiently large to produce a and 13(e). At about the middle (fig. 13(b)) or im- visible tuft flow pattern within a reasonable number mediately aft (fig. 13(c)) of the canopy, the forebody of time steps. Note that the number of time steps vortex migrates downstream into a region where its used in computing tile unrestricted streamline traces trajectory becomes affected by the much stronger is constant, which results in variable length traces be- neighboring LEX vortex system. The forebody vor- cause of the nommiform distribution of the velocity tex flow is initially drawn into the LEX-fuselage magnitudes in a given flow region. Tile tuft flow pat- juncture from which it is entrained outboard by the terns of the model shown in figures 15(a) 15(c) quali- LEX vortex system. Note that just aft of the wing tatively simulate those patterns including the stalled LEX leading-edge juncture, the streamlines originat- flow region over the wing observed on the flight ing from the forebody vortex core split; some wrap aircraft.

around the LEX vortical flow and the rest interact The simulated tuft flow patterns appear to have with the wing flow field.

been influenced by the off-surface flow structures The HAI/V in-flight photographs (ref. 9) of the such as the LEX vortices. This effect is particularly tufts as well as the LEX primary vortex core smoke evident over the wing in tile aft inboard region where visualization are presented in figures 14(a), 14(b), tile tufts indicate a spanwise flow pattern caused by" and 14(e) for c_ _ 20 °, 25 °, and 30 °, respectively.

the flow expansion around the LEX vortex break- Tile photographs clearly illustrate the LEX vortex down at o_ = 25.8 ° and 30.3 °. This flow expansion breakdown just ahead of the vertical tail at ct around the LEX vortex breakdown and the result- 20 ° and its upstream progression with increasing ing interaction with the wing flow field is also evi- angle of attack. In addition, the tufts show the dent in the computational results shown earlier in fig- surface flow patterns over the wing, LEX, fuselage ures 13(b) and 13(c). The accuracy of the predicted aft of the canopy, and the vertical tail. In general, longitudinal location of the LEX vortex breakdown for this range of angle of attack, the tufts reveal as a function of angle of attack is discussed in the a fairly orderly flow pattern over the LEX up to next two paragraphs.

the LEX-wing leading-edge juncture. However, the Several approaches, publieally available in the sci- tufts clearly indicate a chaotic flow pattern over the entific literature, have been devised to locate the wing and the vertical tail with some tufts standing vortex breakdown in a given flow structure. One up off the surface; these disordered flow structures such method that has been widely investigated and are directly attributed to stalled flow over the wing ::1 |1 ration at the selected flow conditions (table I) with

is adoptedheredefines the onsetof vortex break-

downat a point in the corewherethe axial veloc-

(51 = 25 ° are shown in figures 17(a) 17(e). The sur- face pressure coefficients arc contoured at constant

ity component becomes zero(i.e.,u = 0) or reverses

values ranging from 1.0 to -3.0 for all three angles of

direction(i.e., u _<0) fromthat of the free-stream

component. (See ref. 29.) The present numerical so- attack. (See the color bar.) At high angles of attack, the suction peaks in two small regions of the LEX lutions are examined one cross-flow plane at a time to determine the magnitude of u within the LEX apex and over the blended flap exceed the lower con- primary vortex core. By this analysis, no evidence toured limit of -3.0; the pressure coefficients in these exists that a vortex breakdown occurs at (_ = 19°; two regions are represented by solid white. Limita- tion of the pressure coefficient contours to a narrower however, at a = 25.8 ° and 30.3 °, the LEX vortex range would allow more color variation, which wouht breakdown develops longitudinally at x/l _ 0.72 and accentuate the pressure gradients. The following sta- 0.65, respectively.

tions at which both flight and wind tunnel data have been measured are highlighted in white: FS 85, 107, Tile predicted LEX primary vortex breakdown lo- 142, 184, 253, 296, and 357.

cations arc presented in figure 16 along with those ob- tained from different flight (ref. 42) and wind tunnel The effects of angle of attack on the com- (refs. 12 and 15) experiments at subsonic conditions puted surface pressure coefficients as presented in for various angles of attack. Tile experimental inves- tigation of reference 15 was conducted in a low-speed figures 17(a) 17(c) appear to be most pronounced tunnel with a 7- by 9-ft test section on a 1/9-scale in two regions. These regions over the LEX and model of the F/A-18 aircraft at Re _ 1 × 106. The the wing upper surface are directly influenced by wind tunnel data for the longitudinal location of the the neighboring off-surface LEX vortex system and LEX vortex breakdown are presented over a range stalled flow over the wing, respectively. With increas- of a = 21.5 ° to 29 ° for the configuration with and ing angle of attack, the LEX vortical flow" appears to without the empennage (i.e., vertical and horizontal accumulate more strength as evidenced by the higher suction-peak footprint. At high angles of attack (i.e., tails). The data for the baseline configuration (i.e., with empennage) indicate that the vortex break- _> 25.8°), the increase in the LEX vortex strength down location moves upstream with increasing angle not only affects the aerodynamic loads on the LEX of attack and that the overall characteristics gener- surface, but it also has significant influence on the ally correlate well with the data gathered from other adjacent surfaces. For example, figure 17(c) clearly sources. However, the data (rcf. 15) presented in illustrates regions of low pressures acting on the fuse- lage aft of the canopy; these pressure levels are com- figure 16 clearly indicate that the LEX vortex break- down moves further aft without the empennage par- parable in magnitude to those computed on the LEX ticularly for the lower range of angle of attack (i.e., upper surface.

21.5 ° < a _< 24°). As expected, the empennage and, in particular, the vertical tails, which arc lo- At a = 19 °, figure 17(a) indicates that a major cated downstream in the path of the LEX vortical area of the wing upper surface, aft of the wing-flap flow, induce a pressure-fieht disturbance that prop- hinge line, has pressure coefficient levels of about agates upstream and precipitates vortex breakdown.

-0.7 _< Cp _< 0.4. As discussed earlier in conjunc- The vortex breakdown location is predicted farther tion with the tuft patterns (figs. 14(a) 14(c) for flight aft than those obtained experimentally at c_ _ 26 ° tests and figs. 15(a) 15(c) for numerical simulation), and 30 °. Although no data are presented in refer- this portion of the wing exhibited chaotic flow char- cnce 15 for the LEX vortex breakdown location at acteristies attributed to stalled flow. With increasing c_ = 19 °, which corresponds to the angle of attack angle of attack, the surface pressure coefficients com- of interest in the present computation, the trend of puted on the wing upper surface show an extended the data reported for the tailless (i.e., without em- region of lower Cp levels (i.e., Cp <_ -0.7) because pennage) configuration indicates a strong possibility of the localized flow expansion. Also evident was a of a coherent vortex system over the entire length suction-peak footprint associated with a leading-edge of the configuration. The absence of a LEX vortex vortex flow, which developed over the blended flap breakdown at c_ = 19 ° is consistent with the present region at o_ = 19 ° and intensified significantly with computational prediction as discussed in the previous increasing angle of attack. At c_ = 19 °, the surface paragraph. pressure coefficients indicate a small suction-peak footprint associated with the wingtip vortical flow, Surface pressures. The static surface pressure co- which does not appear in the solutions at higher an- efficients computed for the F/A-18 aircraft configu- gles of attack.

The surfacepressure coefficients computedfor (See fig. 15(c).) The LEX lower surface pressure dis- the F/A-18 aircraftconfiguration at all threeangles tribution shows increased compression at the higher of attackare presented in figures18(a)and 18(b). angles of attack.

Tile pressure coefficients at the forcbodystations

The correlations of the computed surface pressure

are shown in figure 18(a) and are plotted as a

coefficients with the flight data for the forebody and

functionof 0. (See fig. 3.) In general, the forebody

the LEX are presented in figures 19, 20, and 21 for surface pressure distribution shows an increasing = 19 °, 25.8 °, and 30.3 °, respectively. Note that the suction-peak level with increasing angle of attack.

flight data shown in figures 19(a) and 19(b) are ob- The computed surface pressures suggest an incipi- tained at slightly different flow conditions and with ent flow separation at 0 _ 150 ° (starboard) and 210 ° some geometrical differences between the numerical (port side) between FS 142 and FS 184 for c_ = 25.8 ° model and the flight vehicle. Experimental aero- and FS 107 and FS 142 for a = 30.a °. These flow sep- dynamic effects from the latter geometrical differ- arations would subsequently form the leeward fore- ences have been found to bc small and arc discussed body vortices with clearly defined suction-peak foot- in detail in the following section.

prints (i.e., 0 ,,_ 155 ° and 205 °) at FS 184 for both c_ = 25.8 ° and ao.3 °. Because the forebody geom- Computational and flight pressure coefficient etry is composed of a smooth curved surface with data for the entire forebody length are in excellent no discontinuities (or limited to within the numeri- agreement throughout the examined range of flow cal discretization error), the triggering mechanism for conditions. The computational results not only pre- the resulting flow separation is an adverse pressure dict the overall pressures as well as the trends but gradient within tile boundary layer. accurately simulate tile pressure distributions that correspond to small flow features such as the leeward forebody vortices. The pressure data disagreements The computed LEX surface pressure coefficients at FS 142 (0 _ 90 ° and 270 °) are caused primarily by are plotted in figure 18(b) as a function of LEX- an antenna fairing on the HARV that was not mod- exposed semispan y/s. The LEX pressure distribu- eled numerically. This antenna fairing can clearly be tions are presented for the same range as in the pre- seen in the flight photograph of the ttARV (fig. 14(a)) vious color contour figures 17(a) 17(c). In general, just ahead of FS 142 (highlighted in white). Note the LEX upper surface pressure distribution can be that for c_ = 30.3 °, the suction peak associated with characterized by a large suction-peak footprint, at tile primary vortex flow at FS 142 (0 _ 158 ° and y/s _ -t-0.50 associated with the primary vortex sys- 202 °) is slightly underpredicted.

tem. At y/s ,_ -t-0.80 just outboard of this large suction-peak footprint at FS 296 and FS 357, areas The computed upper and lower surface pressure of smaller suction-peak footprints exist that corre- coefficients for the LEX are generally in good agree- spond to the LEX secondary vortex system. Note ment with the flight data at all three angles of attack.

that the sharp spikes in the LEX upper surface pres- However, a more detail assessment of the pressure sure distribution just inboard of the leading edge correlations reveals some differences and the possible (i.e., y/s _ ±1) result from numerical artifacts and causes. In general, the correlations tend to degrade in have occurred previously in numerous computational the outboard region, which is essentially dominated studies of vortical flow separations from sharp-edged by the LEX secondary vortex flow. The LEX pri- configurations. (See refs. 17, 20 24, and 43 44.) As mary vortex suction peak is predicted to be slightly with tile forebody, the LEX pressure distributions outboard at FS 253 and the magnitude is under- also indicate higher suction peaks with increasing estimated at FS 296 throughout the examined range angle of attack, except at FS 357, where the lack of c_. However, at the last LEX station FS 357, the of increase in the LEX primary suction peak for magnitude of the primary vortex suction peak is pre- > 25.8 ° can be attributed to the influence of vortex dicted very well at the higher angles of attack of 25.8 ° breakdown. IIowever, this effect does not appear to and 30.3 ° but is underestimated at 19 ° .

impact the secondary vortex suction peak as evident from its consistent increase with increasing angle of Finally, the complete computational results are attack. Note that the computed surface pressures at correlated with the corresponding flight data for the c_ = 30.3 ° clearly indicate a small low-pressure re- forebody and the LEX in figures 22(a) and 22(b).

The results clearly demonstrate the accuracy with gion over the upper surface LEX-fuselage juncture which tile theoretical solutions predict the sensitiv- at FS 357 (y/s _ 4-0.1). This low-pressure region, which has also been seen both in wind tunnel and ity of the surface pressures to changes in angle of flight data, is chiefly attributed to the entrainment of attack. The incremental changes and trends of the computed surface pressure distributions as a function the forebody vortices at the LEX-fuselage juncture.

":! |li range of flow conditions at 5y = 25 ° and 34 °, respec- of angle of attack appear to agree well with the cor- tively. Generally, the forebody pressures remain in- responding flight data. In particular, note the fairly sensitive to the empennage presence regardless of the good prediction for the LEX primary vortex suction flap deflection angle. However, the LEX pressures peak at FS 357, which reveals the upstream influ- begin to be influenced by the presence of the tails, ence of the blockage precipitated by tile vortex break- particularly in the aft LEX region. In general, the down. A favorable correlation is also presented for experimental data indicate that the mlgmentations the LEX lower surface pressure distributions, which of the vertical-horizontal tails result in tile following: clearly demonstrates pressur9 sensitivity to the angle of attack. As expected, the computed lower surface 1. Insignificant cffcct on the forcbody pressures pressure distributions at FS 357 exhibit excess com- throughout the examined ranges of 5f and (_ pression caused by fairing and closing off the inlet 2. Negligible effect oil the LEX pressures measured face and are discussed in the next section.

at FS 253 and FS 296 throughout the examined The present computational results are encourag- ranges of 5.[ and ct ing for simulating the overall flow features and the 3. A slight decrease in the LEX vortex suction peak pressure distributions for the forebo(ty and LEX con- at FS 357; the effect is greater with increasing 5f' figuration at various flight conditions. Nonetheless, particularly for c_ _> 25.8 ° a wind tunnel experiment was initiated to evaluate Effect of inlet fairing. The effect of the in- the aerodynamic effects of various configurational dif- let fairing on tile measured forebody and LEX sur- ferences between the flight vehicle and the numer- face pressure coefficients is presented in figures 26(a) ical model, such as inlet flow simulation, empen- and 26(b), respectively, for the flow conditions of nage, and the deflected flap geometry. As mentioned interest. These aerodynamic data were obtained on earlier, these simplifications of the numerical model the configuration that included the empennage and were incorporated because of the limitations imposed by either the flow solver and/or the grid generation 5f = 0 °. TILe results presented in the figure clearly demonstrate that fairing over the inlet face has very complexity.

little effect on the measured forebody surface pres- Wind Tunnel Data sure coefficients throughout the examined range of a _.

However, the fairing over the inlet appears to slightly As discussed earlier, the experimental data pre- decrease (i.e., more negative) the measured pressure sented here were obtained with a 0.06-scale F/A-18 coefficients associated with the LEX primary and the aircraft model which was tested in the Langley 7- secondary vortex suction peak, particularly at the ad- by 10-Foot High-Spee<t Tunnel. The primary objec- jacent aft stations. Note that tile latter effects seem tive of the test was to validate the present compu- to diminish at higher angles of attack. As expected, tational results by providing experimental data on a the fairing over the inlet causes only slight flow com- configuration that was more representative of tile nu- pression under the LEX as reflected in tile lower sur- merical model. The experimental data analysis was face pressure coefficient measurements at FS 357.

conducted to isolate tile acrodynamic effects of the Effect of flap deflection. The effect of the wing empennage (vertical and horizontal tails), inlet, and leading-edge flap deflection on tile forebody and LEX various flap deflection settings with and without tile pressure coefficients measured on tile CFD wind tun- empennage by evaluating the surface pressure coeffi- nel model is presented ill figures 27, 28, and 29 for cients measured on the forebody and the LEX.

a _ 19 °, 26 ° , and 30 ° , respectively. The results clearly indicate that tile flap deflection angle has neg- Effect of empennage. The forebody and LEX pressure coefficients measured on the CFD wind tun- ligible influence on tile forebody and the LEX pres- sure coefficients throughout the examined range of a nel model, with and without empennage, are pre- except at FS 357 for a _ 30 °. (See fig. 29(b).) At this sented in figures 23(a) and 23(b) for 5f = 0 ° at LEX station, tile data reveal only a slight increase in three angles of attack. In general, tile removal of both tile primary as well as tile secondary vortex suc- tim empennage has minimal effect on measured pres- sure coefficients of tile forebody as well as the LEX; tion peaks (i.e., more negative) with increasing flap deflection. Because of these small effects, the ap- at a = 19 ° tile difference is almost indistinguishable.

Also, note that at FS 357, the removal of the tails proximation (figs. 20 and 21) made in computing the causes a very small increase in the suction level at flow over the configuration with @ = 25 ° (instead of a = 25.8 °.

5I = 34°) for c_ _> 25.6 ° is considered reasonable.

One of the objectives of this study was to ascer- These tail effects on the forebody and LEX pres- tain whether the tails of tile F/A-18 aircraft CFD sures arc presented in figures 24 and 25 for the same

wind tunnelmodelwouldalter tile previous conch>

pressure measurements on tile forebody (figs. 33 35) sionwith regard to theeffectofflapdeflection onthe reveal pressures that are slightly higher (i.e., more

forebodyand LEX pressures. Figures 3O 32 show

positive) than the flight data with the correlation for the experimental data obtained from tile F/A-18 air- the aft stations FS 142 and FS 184 degrading with craft CFD wind tunnel model with the empennage.

increasing angles of attack (c_ = 25.8 ° and 30.3°).

The aerodynamic effect of wing leading-edge flap de- The degradation in the surface pressure correlations flection on tile forebody and LEX pressure measure- is also apparent in the suction-peak regions of lee- ments appears to be insignificant over the examined ward forebody vortices at 0 _ 158 ° and 202 ° . As range of c_ except at FS 357. The pressure distribu- compared earlier (figs. 19 21), the pressure disagree- tion at the last LEX station shows a small sensitiv- ments at FS 142 (0 _ 90 ° and 270 °) are primarily ity to the flap deflection at all three angles of attack. caused by an antenna fairing on the HARV that was Unlike the results (fig. 29(b) for FS 357) discussed in not incorporated on the CFD wind tunnel model.

the previous paragraph, tile slight increase in both Figures 33 35 clearly indicate excellent correla- tile primary and the secondary vortex suction peaks tion between the wind tunnel and flight data for all at FS 357 for o_ _ 30 ° is no longer achieved with the LEX stations in terms of both magnitude and general empennage installed. Actually, at this angle of at- trends throughout the examined range of c_. These tack, the pressure distribution over the last LEX sta- favorable correlations are attributed to the inviseid tion (fig. 31(b)) indicates a slight drop in the primary flow characteristic of the LEX primary vortex sep- and secondary vortex suction peak when increasing aration line (i.e., fixed at the sharp leading edges), 5f from 25 ° to 34 °. Tile reduction in the LEX vor- which leads to the development of the leeward vorti- tical flow suction peak at a _ 30 ° with 5f = 34 ° cal flows. As a result, the pressure distribution of the can be attributed to the nearby LEX primary vor- LEX primary vortex flow indicates only a small sen- tex breakdown at x/l _ 0.45 (fig. 16) precipitated sitivity to the difference between the flight and wind by the empennage. In general, the effect of wing tunnel Reynolds numbers. Tile excellent correlation leading-edge flap deflection on the forebody and LEX also extends to the LEX outboard region where the surface pressure distribution over the F/A-18 aircraft flow separation that leads to the formation of tile sec- CFD wind tunnel model with and without the em- ondary vortex structure is generally considered to be pennage is small.

a boundary-layer phenomenon. Note that the slight pressure data disagreement on the lower surface of Flight Versus Wind Tunnel Data the last LEX station is primarily from additional compression caused by the fairing of the inlet face As discussed in tile previous section, the wind of the experimental wind tunnel model.

tunnel data are primarily used to determine the aero- dynamic effects of the various geometrical differences CFD Versus Wind Tunnel Data between the numerical model and the HARV. Be- cause the ultimate objective is to validate tile com- The computational solutions were all obtained at putational results with the flight data, the accuracy flight Reynolds numbers which were generally about with which the wind tunnel data simulated the flight an order of magnitude higher than those achieved Reynolds number flow characteristics is important.

experimentally in the wind tunnel investigation. As As mentioned earlier, the forebody of the scale model a result, a computational study was performed to had grit strips that were positioned longitudinally examine the effect of Reynolds number on the solu- across the windward plane of symmetry at azimuthal tions. After assessing the Reynolds number effect, angles 0 = 45 ° and 315 ° to trip tile expected laminar the measured surface pressure coefficients are corre- boundary layer to a turbulent flow and thus simulate lated with the computational results for the forebody tile assumed flight flow characteristics.

and the LEX configuration. In addition, the aero- dynamic effects of the wing leading-edge flap deflec- The forebody and LEX surface pressure coeffi- tions are evaluated experimentally as well as compu- cients measured oil the CFD wind tunnel model with tationally. Finally, the measured forces and moments empennage are presented in figures 33, 34, and 35, are correlated with the computational results.

for a _ 19 °, 26 °, and 30 °, respectively. As discussed earlier (figs. 26(a) and 26(5)), the aerodynamic ef- Reynolds number effect. The primary objective of fects on tile forebody and LEX pressures from the this section is to determine if the results of a typical fairing of the inlet face when compared with the existing solution computed at flight flow conditions flow-through inlet were experimentally very small, are sensitive to an order-of-magnitude reduction in confined only to the last LEX station, and diminished Reynolds number. A new solution with a fully tur- quickly at higher angles of attack. The wind tunnel bulent flow assumption was computed with the same I | 1 = lent agreement with the flight data throughout tile flow-field grid by continuing thc solutions from the examined range of (_. (Sec figs. 19 21.) In addition, results that had been obtained earlier at flight flow measured surface prcssure distributions at FS 184 do conditions. As expected, the new converged solutions not indicate the expected suction-pcak footprints as- at the wind tunnel Reynolds numbers indicate that sociated with the presence of the vortical flows at the flow-field grid provides finer resolution of the higher angles of attack (i.e., c_ _> 25.8°).

boundary layer than that obtained earlier at flight Reynolds numbers. This finer grid resolution of the The primary and secondary vortex suction-peak turbulcnt boundary layer flow is naturally extended footprints in the measured LEX upper surface pres- onto the laminar sublayer region where it was rc- sure distributions indicate that the expected overall solved with y+ .._ 1 instead of y+ _ 3 for tile car- flow physics of the LEX configuration has been ex- lier computations at flight Reynolds number. On perimentally simulated. The computed pressure co- the basis of the prior solutions (rcf. 17) obtained on efficients on the LEX upper surface appear to be in thc isolated F/A-18 aircraft forebody-LEX configu- reasonable agreement with the corresponding wind ration, this order of grid refinement is not expected tunnel data for the cxamincd range of c_. As cx- to have any significant effect oil the prescnt compu- petted, the subject correlations rcvcal some exist- tational results.

ing differenccs that are generally very similar both in trcnd and magnitude to those discussed earlier in The effect of Reynolds number on the computed conjunction with tile computational-flight data com- forcbody and tile LEX surface pressure coefficients is parison. (Sec figs. 19 21.) Tile computed lower sur- shown in figures 36(a) and 36(b). The results clearly face pressure distribution at FS 357 is clearly in good indicate that thc computed surface pressure coeffi- agreement with the measured wind tunnel data at all cients arc insensitive to the change in Reynolds num- ber at c_ = 19 ° and Mzo = 0.34. At this flow con- three angles of attack. This favorable correlation on the LEX lower surface can be attributed mainly to dition, the Reynolds number effect on the computed thc similarity of the geometry representation for the forces and pitching moment was also small and is dis- inlet fairing in both the numerical and the CFD wind cussed later in conjunction with the measured wind tunnel models.

tunnel data. To assess thc sensitivity of the com- putational results to changes in Reynolds number at Finally, the complete computational results arc the higher angles of attack, a similar computational correlated with the corresponding wind tunnel data study was performed at c_ = 30.3 ° and Moc = 0.24.

over the forebody and tile LEX in figures 40(a) and At this flow condition, the results also indicated that 40(b), respectively. The results clearly show the sen- the surface pressure cocfficients, forces, and pitching sitivity of tile forcbody and the LEX surface pres- moment were generally insensitive to the change in sure distribution to the changes in angle of attack Reynolds number. Note that the small sensitivity for both the computed and the measured wind tun- of the surface pressure distribution to the change in nel data. Similar to the earlier comparisons between Reynolds number for a comparable range and magni- the computed and the flight test results (figs. 22(a) tude has also been reported in reference 45 for a tan- and 22(b)), the present wind tunnel data correlate gent ogivc configuration at c_ = 30 ° and M_c = 0.2.

reasonably welt with the computational results in These findings justify the surface pressure correla- trends and incremental changes of the surface pres- tions between the present wind tunnel data and the sures as a flmction of angle of attack except for the computational rcsults that have been obtained at vortex flow simulation at the aft forebody stations flight Reynolds number flow conditions.

for the range of higher c_. As discussed earlier, the discrepancies of surface pressures for the forel)ody, Surface pressures. Tile computed forebody and which had a smooth surface geometry with no dis- the LEX surface pressure coefficients are compared continuities, are attributed chiefly to the lack of scale with the experimental data obtained on the CFD simulation of the high Reynolds number flow, partic- wind tunnel model (figs. 37 39) for the range of c_ of ularly in the separated flow regions where the viscous interest. Tile computed results are the same as those effects dominate tile ensuing flow characteristics.

correlated earlicr with the flight test data. Also, note tile consistency of tile configuration geometrical Effect of flap deflection. The primary objec- representation used for both sets of data such as tile tive in this section is to investigate the capabil- flap deflection angle, empennage, and inlet.

ity of the present computational method to predict the aerodynamic effect resulting from different wing The pressure coefficients for the forebody indicate leading-edge flap deflections. The computational re- that the wind tmmel data measurements are slightly sults as well as the experimentally measured surface higher (i.e., more positive) than the computational pressure coefficients for the forebody and the LEX predictions, which wcrc shown earlier to be in excel- configuration arepresented for both _ = 0° and25° for a range that bounds the overall available longi- in figures41(a)and 41(b). The experimental data tudinal aerodynamic characteristics. Among others, were obtained at flowconditions that wereveryclose two specific conclusions can be drawn from the re- to thoseof the computations with the exception of sults with respect to the effects of Reynolds number on the computed forces and moments and the correla-

tile Reynolds number.However, at theseflowcon-

ditionsdiscussed earlier(figs.36(a)and36(b)),the tion betwccn predicted and measured data. Similar computed surfacepressure coefficients for the fore- to the earlier findings in conjunction with tim sur- bodyandtile LEX configuration wereinsensitive to face pressures, the cffect of Reynolds number on the

the Reynolds number. computed total forces and moments also appears to

be very small. However, note the slight increase in

The aerodynamic effectof wingleading-edge flap

the drag coefficient, which is attributed directly to

deflection on tile computed surfacepressure coeffi-

the reduction of the Reynolds number to match that

cientsoil the forebody appears to be verysmalland

achieved in the wind tunnel expcrimcnt. Further-

nearlyconstant. Similarly,this insensitivityof the

more, the computed results presented in figure 42

forebody pressures to the flap deflection is alsoev-

agree reasonably well with the measured wind tun-

ident from the experimental wind tunnel data pre-

nel data except for the total lift coefficient, which

sentedill figure41(a). However, the aerodynamic

appears to have been slightly underpredicted. Thc

influence of flap deflection on the computed surface

lift underprediction at a = 19 ° is not surprising be-

pressure coefficients for the LEX configuration ap-

cause as shown earlier with regard to the surface

pearsto be slightly morepronounced, particularly

pressure coefficients (fig. 37(b)), the computational

at the aft stationswherethey become physically

results also underpredicted the measured LEX pri-

closer to the flapconfiguration. In general, tile com-

mary vortex suction peak in the aft stations. As a

putedresultsshowthat the flapdeflection causes an

result, the LEX vortex lift contribution to the total

increase in the LEX suction-peak levelin a region

lift has probably been compromised.

whichessentially liesbelowthe primaryvortexflow.

However, theexperimental windtunneldataindicate

Experimental and computational longitudinal

only a minimalchange in the LEX-measurcd prcs-

aerodynamic characteristics for the entire range of suredistributionresultingfrom the flap deflection.

flow conditions arc presented in figure 43. Thc ex-

The effects of flapdeflection on the LEX lowersur-

perimental data arc presented for the CFD wind

facepressures alsoappearto bc small,as indicated

tunnel model and for the baseline F/A-18 aircraft

t)5'both thecomputational resultsandthemeasured

configuration without geometrical alteration, which data.

providcd the necessary datum for the force and Forces and moments. In this section, the com- moment data analysis. Similarly, figure 43 shows puted longitudinal aerodynamic characteristics are corresponding computational results that have been correlated with those experimentally measurcd on obtained with the numerical model. Although the the CFD wind tunnel model. Because all the com- latter two sets of data are consistent with one an- putational results were obtained initially at flight other as a function of Mach numbcr (i.e., Mo _ 0.34, Reynolds numbers, which were gcncrally about an 0.25, and 0.24 for c_ _ 19 ° , 25.8 °, and 30.3 ° , order of magnitude greater than those achieved in respectively), they differ slightly from the constant the experiment, tile effect of Reynolds number on M_ = 0.30 at which the data for the baseline F/A-18 the computed forces and moments for a typical casc aircraft configuration were obtained. However, pre- is cvaluated. Note that a similar analysis on tile vious experimental data (ref. 12) obtained from the surface pressure distributions, discussed earlier, in- same 0.06-scalc F/A-18 aircraft wind tunncl model dicated that an order-of-magnitudc reduction in the clearly indicate that small variations (i.e., +0.05) Reynolds number had negligible effects on the com- in Maeh number, particularly in the low subsonic puted surface pressures on the forcbody and tile LEX range, do not have significant influence on overall configuration.

longitudinal aerodynamic characteristics. The latter cffeet as well as the carlicr finding of the influence The longitudinal aerodynamic characteristics, computed at both the flight and the wind tunnel of Reynolds number on the computed longitudinal Reynolds numbers, are presented in figure 42 for aerodynamic characteristics validates the data cor- a = 19 °. Figure 42 also includes the correspond- relations presented in figure 43 despite the inconsis- ing experimental data point obtained for the wind tencies in Mach and Reynolds numbers. The data analysis of the forces and moments is presented in tunnel model that matched the geometry of the nu- merical representation. To be consistent with the the following three categories: the experimental aero- subsequent data analysis, plotting scales are selected dynamic characteristics for thc baseline configuration, I II_ the experimental aerodynamic characteristics for the the forebody flow structures, the wing leading-edge CFD wind tunnel model, and the correlations be- extension (LEX) vortex system, and the subsequent tween the computational results and the correspond- vortex breakdown with increasing angle of attack, ing wind tunnel data. forebody and LEX vortex interactions, and deflected flap leading-edge flow separation leading to a stalled The experimentally measured lift, drag, and flow over the wing upper surface.

pitching moment coefficients for the baseline F/A-18 aircraft configuration indicate essentially stable aero- A wind tunnel experiment was conducted with dynamic characteristics throughout the examined a 0.06-scale F/A-18 aircraft model to ascertain the range of o_. However, some degradation in aero- aerodynamic effects of the geometrical differences be- dynamic characteristics is apparent, particularly in tween the flight aircraft and the numerical model.

the reduced rate of increase in lift coefficient with The wind tunnel data revealed isolated aerodynamic increased angle of attack beyond _18 °, which can effects of the empennage, fairing of the inlet face, and be attributed to tile LEX vortex breakdown (fig. 16) wing leading-edge flap deflection angles. In general, and/or the stalled flow over the wing. (See figs. 14(a) analyses indicated that the geometrical differences an(] 15(a).) The pitching moment characteristics for have only minimal influence on the surface pressure the baseline configuration remain fairly stable (i.e., distributions on the forebody throughout the exam- dCm/da < 0) even at the range of higher c_ despite ined angle of attack range. However, the LEX sur- tile loss of lift caused by the LEX vortex breakdown face pressures are affected slightly by the geometrical ami stalled flow over the wing.

changes, particularly in the aft region and with in- creasing angle of attack.

The experimentally determined effects of configu- ration geometrical simplifications on the longitudinal The experimental wind tunnel data are also com- aerodynamic characteristics are evident ill figure 43 pared with the flight test results to determine the by tile difference between the data presented with capability of the ground-based facility to simulate open and solid circular symbols. Tile fairing of the the flight Reynolds number flow characteristics. This inlet face and the removal of the empennage cause a study revealed that the LEX surface pressure coeffi- slight decrease in lift and an increase in drag coeffi- cients measured on the wind tunnel scale model cor- cients only at the higher angles of attack of 25.8 ° and relate very well with the flight test data. However, 30.3 °. These geometrical simplifications also result in data analysis of the surface pressures on the fore- a reversal of tile pitching moment characteristics (i.e., body indicates some disagreement between the flight dCm/dc_ > 0). Note that the latter change in the and wind tunnel data, particularly in the aft stations pitching moment characteristics can be attributed di- where the flow is separating at tile higher angles of rectly to the absence of the horizontal tail.

attack.

The computed longitudinal aerodynamic charac- The wind tunnel data are presented for the longi- teristics (i.e., solid square symbol) for the geomet- tudinal aerodynamic characteristics measured on the rically simplified F/A-18 aircraft configuration com- baseline F/A-18 aircraft configuration as well as the pare favorably with the corresponding wind tunnel CFD wind tunnel model. These data reveal that the data (i.e., solid circular symbol). Tile computed lift, fairing of the inlet face and the removal of the empen- drag, and pitching moment coefficients correlate rea- nage cause a slight decrease in lift and an increase in sonably well with experimental data throughout the drag coefficients only at higher angles of attack. As examined range of ct.

expected, the experimental wind tunnel data also in- dicate that these geometrical simplifications resulted Concluding Remarks in a pitch-up moment characteristic. Furthermore, the computed longitudinal aerodynamic character- Flow analyses of results from a variety of flight istics correlate reasonably welt with the experimen- tests, wind tunnel experiments, and thin-layer tal measurements obtained on the CFD wind tunnel Navier-Stokes flow simulations are presented for model.

the F/A-18 aircraft configuration. Tile computa- tional results are compared with flight test data of off-surface flow features, surface tuft flow patterns, and surface pressure distributions for three angles of NASA Langley Research Center attack. In general, the computational results cor- Hampton, VA 23681-0001 rectly predict the major flow characteristics such as August 30, 199.1 References of Pitch Rate and Yaw on LEX Generated Vortices of an F/A-18 Fighter Model. AIAA-91-0280, Jail 1991.

1. Thomas, James L.; and Newsome, Richard W.: Navier- 15. Martin, C. A.; and Thompson, D. H.: Scale Model Stokes Computations of Lee-Side Flows Over Delta Wings.

Measurements of Fin Buffet Due to Vortex Bursting on AIAA-86-1049, May 1986.

F/A-18. Maneuvering Aerodynamics, AGARD-CP-497, 2. McMillin, S. Naomi; Thomas, James L.; and Murman, Nov. 1991, pp. 12-1 12-10.

Earll M.: Navier-Stokcs and Euler Solutions for" Lee-Side 16. Meyn, Larry A.; Lanser, Wendy R.; and James, Kevin Flows Over Supersonic Delta Wings A Correlation With D.: Full-Scale High Angle-of-Attack Tests of an F/A-18.

Experiment. NASA TP 3035, 1990.

AIAA-92-2676, June 1992.

3. Polhamus, Edward C.: A Review of Some Reynolds Num- 17. Ghaffari, Farhad; Luckring, James M.; Thoma.s, James L.; ber Effects Related to Bodies at High Angles of Attack.

and Bates, Brent L.: Navier-Stokes Solutions About the NASA CR-3809, 1981.

F/A-18 Forebody-LEX Configuration. AIAA-89-0338, 4. Keener, Earl R.: Flow-Separation Patterns on Symmetric Jan. 1989.

Forebodies. NASA TM-86016, 1986.

18. Thomas, James L.; _,Valters, Robert W.; Reu, Taekyu; 5. Fisher, David F.; Banks, Daniel W.; and Richwine, David Ghaffari, Farh_ut; Weston, Robert P.; and Luckring, M.: E-18 High Alpha Research Vehicle Surface Pressures: James M.: A Patched-Grid Algorithm for Complex Con- Initial In-Flight Results and Correlation With Flow Visu- figurations Direeted Towards the F/A-18 Aircraft. AIAA- alization and Wimt-'ISmnel Data. A Collection of Techni- 89-0121, Jan. 1989.

cal Papers AIAA 8th Applied Aerodynamics Conference, 19. Benek, J. A.; Steger, J. L.; Dougherty, F. C.; and Part 1, Aug. 1990, pp. 421 ,151. (Available as AIAA-90- Buning, P. G.: Chimera: A Grid-Embedding Technique.

3018-CP.)

NASA TM-89246, AEDC-TR-85-64, 1986. (Available 6. ttall, Robert M.: Influence of Reynolds Number on Fore- from DTIC as AD A167 466.)

body Side Forces for 3.5-Diameter Tangent-Ogivc Bodies.

20. Sehiff, Lewis B.; Cummings, Russell M.; Sorenson, Reese AIAA-87-227.t-CP, Aug. 1987.

L.; and Rizk, Yehia M.: Numerical Simulation of tIigh- 7. Gilbert, William P.; and Gatlin, Donald H.: Review of Incidence Flow Over the F-18 Fuselage Forebody. AIAA- the NASA tIigh-Alpha Technology Program. High-Angle- 89-0339, Jan. 1989.

of-Attack Technology, Volume I, Joseph R. Chambers, 21. Cummings, Russell M.; Rizk, Yehia M.; Schiff, Lewis B.; William P. Gilbert, and Luat T. Nguycn, cds., NASA and Chaderjian, Neal M.: Navier-Stokes Predictions of the CP-3H9, Part 1, 1992, pp. 23 59.

Flowfield Aroumt the F-18 (HARV) Wing and Fuselage at 8. Regenie, Victoria; Gatlin: Donald; Kempel, Robert; and Large Incidence. AIAA-90-0099, Jan. 1990.

Mathcny, Neff: The F-18 High Alpha Research Vehi- 22. Rizk, Yehia M.; Sehiff, Lewis B.; and Gee, Ken: Numer- cle: A High Angle-of-Attack Testbed Aircraft. NASA ical Sinmlation of the Viscous Flow Around a Simplified TM-104253, 1992.

F/A-18 at High Angles of Attack. AIAA-90-2999, 1990.

9. Fisher, David F.; Del Frate, John tt.; and Richwine, David 23. Rizk, Yehia M.; and Gee, Ken: Numerical Prediction of M.: [a-Flight Flow Visualization Characteristics of the the Unsteady Flowfield Around the F-18 Aircraft at Large NASA F-18 High Alpha Research _'_hiclc at High Angles Incidence. AIAA-91-0020, Jan. 1991.

of Attack. NASA TM-4193, 1990.

24. Ghaffari, Farhad; Luckring, James M.; Thomas, James 10. Fisher, David F.; Del Frate, John tI.; and Zuniga, Fanny L.; Bates, Brent L.; and Biedron, Robert T.: Multiblock A.: Summarg of In-Flight Flow _Zisualization Obtained Navier-Stokes Solutions About the F/A-18 Wing-LEX- From the NASA thigh Alpha Research _'_hicle. NASA Fuselage Configuration. J. Aircr., vol. 30, no. 3, May TM-10173,1, 1991.

June 1993, pp. 293 303.

11. Erickson, Gary E.: Water Tunnel Flow Visualization 25. Smith, B. M.; Brauner, K. M.; Kennicott, P. R.; aud _tSud Tunnel Data Analysis of the F/A-18. NASA Liewald, M.; and Wellington, J.: Initial Graphics Ex- CR-165859, 1982.

change Specification (IGES), Version 2.0. NBSIR-82- 12. Erickson, Gary E.: Wind Tunnel Investigation of Vortex 2631-AF, Feb. 1983. (Available from NTIS as PB 83- Flows on F/A-18 Configuration at Subsonic Throu9h 137448.)

Transonic Speeds. NASA TP-3111, 1991.

26. Brewer, J. A.; and Anderson, D. C.: Visual Interaction With Overhauser Curves and Surfaces. Proceedings of 13. Banks, Daniel W.: Wind-'Ikmnel Investigation of thc Fore- the 4th Annual Conference on Computer Graphics and body Aerodynamics of a Vortex-Lift Fighter Configura- Iuteractive Techniques, ASME, July 1977, pp. 132 137.

tion at High Angles of Attack. Advanced Aerospace Aerodynamics, SP-757, SAE, Oct. 1988, pp. 101 123.

27. Ghaffari, Farhad; Luckring, James M.; Thomas, Jamcs L.; (Availahle as SAE 881419.)

and Bates, Brent L.: Transonic Navier-Stokes Solutions About a Generic Hypersonic Configuration. J. Aircr., 14. Hebbar, Sheshagiri K.; Platzer, Max F.; and Cavazos, vol. 28, no. 6, June 1991, pp. 381 388.

Odilon V.: A Water Tunncl Investigation of the Effects '_Vll i 37.

28. Luckring, James M.; Ghaffari, Farhad; and Bates, Brent Fox, Charles H., Jr.: Real Time Data Reduction Capa- bilities at the Langley 7- by IO-Foot High-Speed Tunnel.

L.: Status of Navier-Stokes Computations About tile NASA TM-78801, 1980.

F/A-18 With Structured Grids. High-Angle-of-Attack Technology, Volume 1, Joseph R. Chambers, William P.

38.

BraMow, A. L.; and Knox, E. C.: Simplified Method for Gilbert, and Luat T. Nguyen, eds., NASA CP-31-19, Determination of Critical Height of Distributed Roughness Part 2, 1992, pp. 703 722.

Particles for Boundary-Layer Transition at Mach Num- 29. Thomas, J. L.; Taylor, S. L.; and Anderson, \V. K.: bers From 0 to 5. NACA TN 4363, 1958.

Navier-Stokes Computations of Vortical Flows Over Low Aspect Ratio \Vings. AIAA-87-0207, Jan. 1987.

39.

Iterriot, John G.: Blockage Corrections for Three- Dimcnsional-FIow Closed-Throat Wind Tunnels, With 30. Vatsa, V. N.; Thomas, J. L.; and tVedan, B. W.: Navier- Consideration of the Eff_et of Compressibility. NACA Stokes Computations of Prolate Spheroids at Angle of At- Rep. 995, 1950. (Supersedes NACA RM A7B28.)

tack. Technical Papers AIAA Atmospheric Flight Me- chanics Conference, Aug. 1987, pp. 488 506. (Available 40.

Gillis, Clarence L.; Polhamus, Edward C.; and Gary, as AIAA-87-2627.)

Joseph L., Jr.: Charts for Determining Jet-Boundary 31. Biedron, R. T.; and Thomas, J. L.: A Gcncralized Corrections for Complete Models in 7- by IO-Foot Clo.sed Patched-Grid Algorithm With Application to the F-18 Rectangular BSnd Tunnels. NACA WR L-123, 19,15.

Forebody With Actuated Control Strake. Comput. Sgs.

(Formerly NACA ARR L5G31.)

Eng., vol. 1, nos. 2 4, 1990, pp. 563 576.

41.

Sellers, William L., III; anti Kjelgaard, Scott O.: The 32. Baldwin, Barrett; and Lomax, Harvard: Thin-Layer Ap- Basic Aerodynamics Research Tunnel A Facility Dedi- proximation and Algebraic Model for Separated Turbulent cated to Code Validation. AIAA-88-1997, May 1988.

Flows. AIAA-78-257, Jan. 1978.

33. Degani, David; and Schiff, Lewis B.: Computation of Su- 42.

Fisher, David F.; and Meyer, Robert R., ,Jr.: Flow personic Viscous Flows Around Pointed Bodies at Large tSsualization Techniques for Flight Research. NASA Incidence. AIAA-83-0034, Jan. 1983.

TM-100455, 1988.

34. Roc, P. L.: Characteristic-Based Schemes for the Eu- 43. Hartwich, Peter-M.; IIsu, C.-H.; Luckring, James M.; ler Equations. Annual Review of Fluid Meehanics, Vol- and Liu, C. It.: Numerical Study of the Vortex Bmst ume 18, Milton van Dyke, J. V. Wehausen, and John L.

Phenomenon for Delta \\rings. AIAA-88-0505, Jan. 1988.

Lumley, eds., Annual Reviews Inc., 1986, pp. 337 365.

35. Van Leer, Brain: Upwind-Difference Methods for Aero- d,L Vatsa, Veer N.; Sanetrik, Mark D.; aim Parlette, Edward dynamic Problems Governed by the Euler Equations.

B.: Development of a Flexit)le and Efficient Multigrid- La_yc-Scale Computations in Fluid Mechanics, Bjorn E.

Based Multibloek Flow Solver. AIAA-93-0677, Jail 1993.

Engquist, Stanley Osher, and Richard C. J. Somerville, eds., American Math. Soe., 1985, pp. 327 336.

45.

Degani, David; Schiff, Lewis B.; and Levy, Yuval: Phys- ical Considerations Governing Computation of Turbulent 36. Fox, Charles It., Jr.; and Huffman, Jarrett K.: Calibration Flows Over Bodies at Large Incidence. AIAA-90-0096, and Test Capabilities of the Langley 7- by lO-Foot High- Jan. 1990.

Spet:d Tunnel. NASA TM X-74027, 1977.

i L-91-65 Figure 1. The F/A-18 High-Alpha Research Vehicle (HARV).

Reference dimensions Sre f = 400 f12 bre f = 37.42 ft _:= i 1.52 ft c.g. = 0.25?=

I L --56"0O

Figure 2. Tile F/A-18 aircraft geometry. All linear dimensions are in feet.

-y -----.j_ y 18o° _ e ,, 90 ° 270 FS 357 0° _' ......

180 ° _, ....

.V _y

270 ° FS 296 " 180 ° 90°@ 270° _'_Y '_- _ _Y FS 107 FS 253 0 o 180 ° 90 ° 270 ° 90 ° 270 °

'8°o

FS 85 FS 70 O° 0 o Figure 3 Planform of F/A-18 aircraft with forebody and LEX-fuselage cross-sectional pressure measurement stations.

Representation

CAD

..... CFD

Faired-over

splitter plate

Splitter plate

(a) FS 401.

tat ion ..... CFD X slot _"_/_ Faired-over cavity (b) FS 441.

Figure 4. Typical CAD and CFD cross-sectional grids.

(a) Oblique top view.

(b) Obliquebottomview.

Figure5. TheF/A-18aircraftCFD surface gridrepresentation.

Figure6. TheF/A-18aircraft CFDsurface grid with undeflected (port) andblended (starboard) flaps.

Deflected wing

leading-edge

Blended flap

re_"

splitter plate

Figure7. Close-up of F/A-18aircraftCFD surface gridandhighlighted surface modifications.

\\\\,\\\\_

,,,,\ \_,_,\\\

...... -- i!i [I I H-( : i (a) Far field.

\ \ \

\

(b) Near field.

Figure 8. The F/A-18 aircraft flow-field blocking strategy.

Figure9. Far-field sideview of F/A-18 aircraftflow-field blockingstrategy.

! |1i

0 m -,5 _ -i.0 -I.5 _ao o .a -2.0 -2.5

-30 I I I I I I

0 I 2 3x 10 3 Iterations (a) Residual convergence.

,.--1 r,.)

0--

I i I i I

-2 [ I 2 3x 10 3 Iterations (b) CL convergence.

Figure 10. Typical convergence characteristics, ct = 19°; Moo = 0.34; Re = 13.5 x 106; 5]" = 0%

L-92-01827

Figure11. Sting-mounted 0.06-scale F/A-18aircraftwindtunnelmodel.

| _|Ii L-92-2805 (a) Oblique rear view.

Metal sheet to close gap betw deflected fla and fuselage leading-edge flap Representation -- CAD ..... CFD CFD WT Splitter plate FS 401 L-92-2801 (b) Close-up of geometrical modifications.

Figure 12. Sting-mounted 0.06-scale F/A-18 aircraft CFD wind tunnel model.

(a) a = 19°; M_o = 0.34; Re = 13.5 × 106;_f = 25 ° .

(b) c, = 25.8 ; Mac = 0.25; Rc 10.8 x 106; _f = 25 °.

Figure 13. Cross-flow normalized total pressure contours with vortex core particle traces.

(c) a = 30.3°;/I,I_= 0.24; Re = 10.2 x 106; _f = 25 °.

Figure 13. Concluded.

LEXprimary

\,'OlleX core.

Disorganized / tUlls (a) a _ 20°; M_c _ 0.3; Re. _ 10 × 106; 6f _ 25 °.

t}rimary .... YOI'|CX core (b) c, _ 25°; Mac _ 0.3; Rc _ 10 x 106; 5/,-_ 34 °.

Figure 14. Tile HARV in-flight surface and off-surface flow visualization.

3O Illi vorlex c_re (c) c_ _ 30°; Mac _ 0.3; Rc -._ 10 × 106; 5f _ 34 °.

Figure 14. Conchided.

(a) c, = 19°; Mac = 0.34; Re = 13.5 × 106; 5f = 25 °.

Figure 15. Unrestricted surface flow pattern with vortex core particle traces.

Forebody primary

vortex core

streamlines

(b) cr= 25.8°; A[_o = 0.25;/_a. = 10.8 × 106; 6f = 25 °.

(c) cr = 30.3°; M_ = 0.24; Re = 10.2 × 106; 6f -- 25 °.

Figure 15. Concluded.

O Flight, smoke, Moo = 0.3, RF = 10 x 106 (ref. 42)

[] Flight, natural condensation, Moo = 0.3, R_- = 10 x 106 (ref. 42) Wind tunnel, vapor screen, M,_ = 0.4, RF = 2 x 106 (ref. 12) • Wind tunnel, smoke, with empennage, M = 0.1, RF = 1 × 106 (ref. 15) • Wind tunnel, smoke, without empennage, Moo = 0.1, R_: = 1 x 106 _,ref. 15) A Present Navier-Stokes predictions, Moo = 0.3, R? = 10 x 106 5O 4O

%

3O O_A • A 25 4, ii • A O I I I I I I I I I I 0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 x// Figure 16. LEX primary vortex-breakdown correlations between flight, wind tunnel, and computational results.

! I I (a) a _ = 19°; Mcc = 0.34; Re = 13.5 x 106; _f = 25 °.

(b) c_ = 25.8°; 2_I_ = 0.25; Re = 10.8 x 106; 6f = 25 °.

Figure 17. Computed surface pressure coefficients.

(c) c_ = 30.3°; 2riot= 0.24; Re = 10.2 × 106; _f = 25 °.

Figure 17. Concluded.

-1.0 FS 85 Data ix, deg M_ R e xl0 _' _f, deg Tails Inlet -.5 CFD 30.3 0.243 10.20 25 Off Faired CFD 25.8 0.253 10.80 25 Off Faired Cp CFD 19.0 0.340 13.50 25 Off Faired .5 1.0 0 45 90 135 180 225 270 315 360 0, dcg -3 -1.0 - FS 107 -2 -.5 FS 253 Cp 0 Cp - I .5 0 I , I _ I , k J_ £, I t I _ _1___ 1.0 = • 1 , I , I , I , I , I , I , I -I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0 45 90 135 180 225 270 315 360 0, deg y/s -3 1.0 - FS 296 FS 142 -2 -.5 Cp -I Cp .5 - t _I , I , I ,_1__ i 1.0 _____I , 1 t I _ 1 _t_ 1 I _ I _ -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 45 90 135 180 225 270 315 360 0 y/s 0, deg -I.0 - FS 184 -3 _ FS 357 -.5 Cp Cp 0 .5 lI , I , I , A_, I , 1 , I , I , 1 1.0 1 , I z I .._._1 , I _ I , I , I -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 45 90 135 180 225 270 315 360 0, deg y/s (a) Forebody. (b) LEX.

Figure 18. Effect of angle of attack on computed surface pressure coefficients.

1.0 FS 85 Data or, dog M_ R e xI0 + 81 , deg Tail,, Inlet -.5 CFD 19.0 0.340 13.50 25 Off Faired Flight 19.1 0.300 I 1.50 25 On Open Cp .5 1.0 l [ , I t t , I , I , I , I , I 0 45 90 135 t 80 225 270 315 360 0, dog -I .0 3 -- FS 253 FS 107 -2 -.5 Cp -I Cp .5 1.0 , 1 , I , I , I 1 I , I , I J I 1 , I , I , I _ I i I , I , I , I 45 90 135 180 225 270 315 360 -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, dog y/s -I .0 FS 142 -3 _- FS 296 -.5 -2 Cp .5 1.0 , I _ I _ I , I , I , I _ I L I ,_5 315 360 -t.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0 45 90 135 180 "_'_ 270 0, dog y/s - 1.0 -- FS 184 -3 _- FS 357 / -.5 -2 F ooo oO o , Cp Cp .5 1.0 , I , I , I _ I i I I _ t , I I , I , I _I , I , I , I z I J I I 0 45 90 135 180 225 270 315 360 - 1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, deg y/s (a) Forebody. (b) LEX.

Figure 19. Correlation of computed surface pressure coefficients with flight data at a _ 19 °.

III -1.0 FS 85 Data I_, deg M_ R e xl0 6 15f, deg Tails Inlet -,5 CFD 25.8 0.253 10.80 25 Off Faired Flight 25.8 0.253 10.80 34 On Open Cp .5 1.0 , I , I , I L I , I , I L L.L_!

0 45 90 135 180 225 270 315 360 0, deg -I.0 -3 FS 107 FS 253 -2 -.5 0 -1 Cp Cp .5 1 I _ I .t_._l , I , I _ I t. I _ _ 1.0 , 1 , 1 , I __.___ I , I t I , I t_J 0 .25 .50 .75 1,00 0 45 90 135 180 225 270 315 360 -I.00 -.75 -.50 -.25 0, dcg y/s -1.0 -3 F FS 296 II f FS 142 -2 / D D DD -.5 [] Cp Cp .5 ___9_-_ 4::_ .-- _ _-_ I _a_.5_.___-t. , I , 1 L I , 1 _ 1,0 0 45 90 135 180 225 270 315 360 -I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, deg y/s -3 -I.0 - FS 184 _- FS 357 -.5 Cp -I Cp .5 0 _"_-..--...._D_ l_ _ _ D D i_..___._../_ I , t , I_ 1 _, L t A , I , I _ I 1.0 , I , I , I ,_..I _ I _ I i _.[_J.2 -I.00 -.75 -.50 -.25 0 .25 .50 ,75 1.00 0 45 90 135 180 225 270 315 360 O, deg y/s (a) Forebody. (b) LEX.

Figure 20. Correlation of computed surface pressure coefficients with flight data at o_ = 25.8 °.

-I.0 FS 85 Data _x. deg M_ R e xl0 6 5f, deg Tails Inlet -.5 CFD 30.3 0.243 10.20 25 Off Faired Flighl 30.3 0.243 10.20 34 On Open Cp 0 .5 1.0 ____ • 1 _ A_ .L_A _ I , I , I 45 90 135 180 225 270 315 360 0, deg -3 -1.0 FS 107 -2 -.5 Cp -1 Cp 0 (

o

.5 1 I , 1 _ ] , I • I , I _ I J I L 1.0 -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 45 90 135 180 225 270 315 360 O, deg y/s -1.0 -3 FS 296 FS 142 ¢_00 000 O o 0 0 -2 -.5 Cp 0 Cp -1 .5 1.0 _h t I I , I _ L i I i I _ J i l 45 90 135 180 __5 270 315 360 -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, deg y/s -3 - -1.0 FS 357 FS 184 -2 -.5 Cp 0 .5 1.0 1 j____ , i , I , I , I , I 1 1 J ] -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0 45 90 135 180 225 270 315 360 0, dog y/s (a) Forebody. (b) LEX.

Figure 21. Correlation of computed surface pressure coefficients with flight data at c_ = 30.3 °.

:t II1

-1.0-

FS 85 Data o_, deg M_ R e xl0 _ 8 t , deg Tails Inlet

-.5

CFD 30.3 0.243 10.20 25 Off Faired CFD 25.8 0.253 10.80 25 Off Faired

Cp

CFD 19.0 0.340 13.50 25 Off Faired o Flight 30.3 0.243 10.20 34 On Open

.5

[] Flight 25.8 0.253 10.80 34 On Open O Flight 19.I 0.300 I 1.50 25 On Open I _ 1.___ 1 • t _ 1 , I , I 1.0 0 45 90 135 180 225 270 315 360 0, deg -3 -I .0 /_o FS 253 o_ -2 -.5_ Cp -I

Cp

o r

5L

I 1.0 ___L I , I __ _k_t._. I , I 0 .25 .50 .75 1.00 -1.00 -.75 -.50 -.25 0 45 90 135 180 225 270 315 360 y/s 0, dcg -3 - -1.0 FS 296 FS 142 O0 000 O (D -.5 _._.,___ _ -._-._.o _." ,.,o_'__ ...,..- - Cp -[

Cp

.5 1.0 __L., J L .I , I , • .___1 0 .25 .50 .75 .00 -I.00 -.75 -.50 -.25 0 45 90 135 180 225 270 315 360 y/s 0, dog -3 - -1.0 FS 357 f FS 184 -.5 o Cp -1

Cp

.5 L__ _ _____1_ ___ I 1.0 0 .25 .50 .75 1.00 -I.00 -.75 -.50 -.25 0 45 90 135 180 225 270 315 360 y/s 0, deg (a) Forebody. (b) LEX.

Figure 22. Effect of angle of attack on computed and measured surface pressure coefficients.

-1.0 - FS 85 Dala or, deg M_ Re xl0 b 131.,deg Tails Inlet -.5 o• • o _ _',,! !:.,.

• Tunnel 30.5 0.251 1.21 0 On Faired

-" i •

• Tunnel 25.9 0.250 1.21 0 On Faired Cp @ Tunnel 19.2 0.342 1.61 0 On Faired O Tunnel 30.6 0.253 1.11 0 Off Faired l

t

.5 !

D Tunnel 25.8 0.252 1.12 0 Off Faired O Tunnel 19.1 0.343 1.46 0 Off Faired 1.0 _ • J I , I , I • l 1 .L x .1 a.___J 0 45 90 135 180 225 270 315 360 0, dog -I.0 -3-- FS 107 FS 253 • _ • • • • o_:O . , itl_o -2 -.. = • ,."

-n i i 4 •• • • llm • |.il*'"lillooiil. if..

i-- -nn • @I_41 Cp -1 @4 @ @ @ 4 4 4 14

I t Ol I

.5 0 1.0 __ , 1 _ 1 t t • t _ I _ I I 45 90 135 180 225 270 315 360 -I.00 -.75 -,50 -.25 0 .25 .50 .75 1.00 0, dcg y/s FS 142 FS 296 • • llii

°1 F,.. "

-.5 oli_ _llio -2 Oi • I ii •001_ r mtmi_mL .... ,l_imm I •i • i + mill- _, .sLoiOiO..._...Ool I co-,L"+'o .. n. "'....

1.0 • J _J _ 1 ,._t _ J J I J I 1 1 _ t .J J _ I J __ _ 0 45 90 135 180 225 270 315 360 -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, deg y/s -I.0 -3 FS 184 [] FS 357 iii -,5 -2 i | joil e i tl

ill " in i" iLl

Cp 0 Cp -1

":l..!i

..;illli;llllilli;.,,_:.- tt

i

i _ t It t t i 1.0 _ _ • 1 , 1 L 1 , 1 _L__I__ • • 1 _l • I _ I , I 0 45 90 135 180 225 270 315 360 -1.0t -.75 -.50 -.25 0 .25 .50 .75 1.00 0, deg y/s (a) Forebody. (b) LEX.

Figure 23. Effect of empennage on measured forebody and LEX surface pressure coefficients at various values of a with 6f = 0 °.

:1 111 -1.0 - FS 85 Data or, deg M_ R(. xl0 '_ 8 t , deg Tails Inlet OO O O

5 ,_,'"! ,.,,: • Tunnel 30.5 0.250 1.15 25 On Faired

o **** III ****

• Tunnel 25.8 0.249 1.15 25 On Faired Cp • Tunnel 19.1 0.342 1.54 25 On Faired 0 Tunnel 30.5 0.251 1.15 25 Off Faired 5 !t I: El Tunnel 25.8 0.251 1.15 25 Off Faired O Tunnel 19,1 0.341 1.51 25 Off Faired 1.0 _ ___1 J l _1 _ [ •_ I_1 0 45 90 135 180 225 270 315 360 0, deg -1,0 -- -3 FS 107 • FS 253 • • • [] • i • -2

.5 o,_,_,. • -",o.._

O0 [] [] i i O o u m'- ui • g_U Ui ( • • • ) / i • • • •

t.,,****II0w001* *_"i

Cp -I Cp 0 @ • @ @

I o Oll

.5 0 1.0 ____ 1 * [ _ 1 t I , I , 1 _ I _t/ l ___1_, I l J • _ 1 _1 , 1 -.75 -.513 -.25 0 .25 .50 .75 1.00 0 45 90 135 180 225 270 315 360 -I.13 0, deg y/s FS 142 FS 296 • • • -1,0 f -3 L Oi i • • [] Oo • -.5 i Io_ _Leo -2 oil iioooo_ / Oli_ __dA_i i -- m Lh_i i il/m TM 0 Cp -1 li_ i i • i Cp • i .5 0

II 11;' " II "" " " " " " " ""_

0 45 90 135 180 225 270 315 360 - 1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 O, deg y/s -1.0 -- FS 184 i FS 357 -.5 -2 @ i i • II I,*ub_l_ go.a_ _ _ • "'tl illllili '" .i ,.

•. .. c, Cp

ot .1;;i i"

I t t t t t

, I ,__I _ I J I _ t , I _ [ U l , I , I ..... I _ 1 i I _ , I , I 45 90 135 180 225 270 315 360 -I.00 -.75 -,513 -.25 0 .25 .50 .75 1.00 0, deg y/s (a) Forebody, (b) LEX.

Figure 24. Effect of empennage on measured forebody and LEX surface pressure coefficients at various values of _ with _I = 25°' -I.0 - FS 85 Data a, deg M_ Re×l 0 o 8j, deg Tails Inlet -.5 • Tunnel 30,6 0.251 1.18 34 On Faired -.! ,,,..

• Tunnel 25.8 0.250 1.17 34 On Faired Cp • Tunnel 19.2 0.341 1.54 34 On Faired O Tunnel 30.5 0.250 1,21 34 Off Faired |:

!

.5 [] Tunnel 25.8 0.251 1.22 34 Off Faired O Tunnel 19.1 0.340 1.59 34 Off Faired 1.0 __ I _ I K l J [ , 1 L 1 , I , I 0 45 90 135 180 225 270 315 360 0, deg -3 -l,0 - FS 107 • FS 253 • • • [] [] :: .,:..

[] I • t i• I|r_i||l __|i |_JJ

I._ -.wJe,ee+• ._%l I . ": +[]

Cp -I 4) 4, • • • • • • O0 i o ot.

.5 1 _ J_ ± I l I _ I J 1 _ J t 1 t _1 1.0 45 90 135 180 225 270 315 360 - 1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, deg y/s _ 3 FS 296

[

- 1.0 _ FS 142 -2 _._J i_ i_ _[][]_

| ,,.| _[]i-. • []

-5 [ _.+_o)

1 _o,e ,'e't • [] | • Cp -I

co OItol'** ***el| |

°i

I ) I + I J 1 , I 1.0 t 1 l __. _1 _ I , t _ _. __ -1.00 -,75 -,50 -.25 0 .25 .50 .75 1.00 0 45 90 135 180 225 270 315 360 0, deg y/s -3 " -I.0 FS 184 -2

-.5 8W

-1 Cp Cp 0

! o!

.5

t t t t l t e

I t l L J _ ! , I J I + l ,___ 1.0 -I.00 -.75 -,50 -.25 0 ,25 ,50 .75 1.00 45 90 135 180 225 270 315 360 y/s 0, deg (a) Porebody. (b) LEX.

Figure 25. Effect of empennage on measured forebody and LEX surface pressure coefficients at various values of c_ with _j- = 34 ° 1 I i -1.0 - FS 85 Data ft, deg M_ R_ xlO_ 81 , deg Tails Inlet -.5 _oo Q_ • Tunnel 30.6 0.250 1.II 0 On Open

" |

• Tunnel 25.9 0.253 1.13 0 On Open O @ Cp • Tunnel 19.1 0.341 1.47 0 On Open 1.21 0 Oi1 Faired O Tunnel 30.5 0.25 I • | .5 1.21 0 On Faired

| 121 Tunnel 25,9 0.250

C' Tunnel 19.2 0.342 1.61 0 On Faired 1.0 0 45 90 135 180 225 270 315 360 0, dcg -3- -1.0 - FS 253 • FS 107 • J -2 -O,_ i • • mo° • • @ • il • • •

|•,,,••,,,no•el,•

Cp -I -@• •

i'

°t!

.5 I __1 t_ _ [ _ J , _± _l L_._ 1.0 0 .25 .50 .75 1.00 -1.00 -.75 -.50 -,25 45 90 135 180 225 270 315 360 0, dcg y/s -1.0 -- FS 142 FS 296 - •

3 L • • [] °o

-.5 - ••-. _,,._•Oo -2 w_ j i • _ Ji•o,_ |" _ _mb .... o..m_tdi_nlin mn n I _iw'_ • ui + im_mm • * Cp __ l__ + * | I **_ Cp 0 .5 I tl..._ "11|| o[ * * * * 1.0 0 45 90 135 180 225 270 315 360 -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, dcg y/s -1.0 - -3 [- FS 357 FS 184 i []w -.5

I i

k

*|_| m@ Cp co -' t i

! 0!

.5 I __ _ _a L _ L ,, J -,_ 1.0 -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 45 90 135 180 225 270 315 360 y/s 0, deg (a) Forebody. (b) LEX.

Figure 26. Effect of inlet fairing on measured forebody and LEX surface pressure coefficients at various values of a with _Sf = 0 °.

-1.0 FS 85 Data 0_,deg M_ R e ×10 '_ 81 , deg Tails Inlet -.5 O Tunnel 19.1 0.343 1.46 0 Off Faired 0 O00000000000Q [3 Tunnel 19.1 0.34I 1.51 25 Off Faired Cp <) Tunnel 19.1 0.340 1.59 34 Off Faired O O0 .5 1.0 I , l _ A_, J , I _ t _ I _ I 0 45 90 135 180 225 270 315 360 O, deg -1.0 - -3 - FS 107 FS 253 -.5 -2 0 o 0 0 OQoI_O0000000000 O01_O 0 Cp Cp -I -(30 0 0 OO 10 0 O0 o 0 0 0 0 .5 - 0 1.0 _, I _ I J t _ 1 _J 1 ._ t_ ..L._ 1 _ LJ , I , I , I , I , I _ I 0 45 90 135 180 225 270 315 360 -I .00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, deg y/s -3 FS 296 -I.0 _ FS 142 -.5 -2 _I_00 oo 0 0 00_ 0 0 Cp Cp -1 u 0000 0 0 1.0 1 _-, 1 _ 1 __A__L J. _ .1_ x_J I _.t _ [ _A _ , I , I 0 45 90 135 180 225 270 315 360 -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, deg y/s -1.0 -3-- FS 184 FS 357 o -.5 -2 _0 o o @ 0 0 °°0_ cp -I _ ;_;_0_ 00_00 Cp O0 0ooOo ° .5 0 0 0 0 0 0 0 0 1.0 l t J 1 t J l 1 _ I , I [ I 1 J 1 , I , I _ ! , I , I , I , 1 __3 45 90 135 180 225 270 315 360 - 1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, deg y/s (a) Forebody. (b) LEX.

Figure 27. Effect of wing leading-edge flap deflection on surface pressure coefficients measured on F/A-18 aircraft CFD wind tunnel model at a _ 19 °.

! 81:

-I.0-

FS 85

Data (x, deg M_ Re xl0 _ _1, deg Tails Inlet

-.5

O0 0 oOOO 1.12 0 Off Faired 0 Tunnel 25.8 0..5.

0 Oooo 0 121 Tunnel 25.8 0.251 I. 15 25 Off Faired

Cp

1.22 34 Off Faired 0 Tunnel 25.8 0.25t 0 0 o .5 1.0 _, 1 , I , I _±__1 _ t • 0 45 90 135 180 225 270 315 360 0, deg -3 -I ,0 -- FS 253 FS 107 -2 -.5 0 O 0 0 D O O0 oOCg°°OOooooooO_ O_bo ° 0 0 0 0

Cp Cp -1

O0 o _o o .5 I _, J . 1 [ • _t • , I , I 1.0 [__l_ 1 , 1 , 1 L [ , I , I -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0 45 90 135 180 225 270 315 360 0, dcg y/s -3 - -1.0 -- FS 142 FS 296 O O O -2 - -.5 a o °°ooc_ O00O_OOoo 0 0 Cp -1 Cp 0 O00 t .5 1 _ 1 L 1 J t 1 1 _ L t_ [ 'J 1.0 -I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 45 90 135 180 225 270 315 360 0, deg y/s -I .0 -3 - FS 184 0 FS 357 OQ O -2 -.5 6c_c_ ° {} Cp Cp -I - 0 QOQO0 0 o .5 0 O 0 0 0 0 0 0 1.0 .___L ____._l , 1 , I J [ • d 45 90 135 180 225 270 315 360 -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 y/s 0, deg (a) Forebody. (b) LEX.

Figure 28. Effect of wing leading-edge flap deflection on surface pressure coefficients measured on F/A-18 aircraft CFD wind tunnel model at a _ 26 °.

-1.0 FS 85 Data a, deg M_ Re x I0 " 5f, deg Tails. Inlet -.5 000 OOQ O Tunnel 30.6 0.253 1.11 0 Off Faired 0 00000 Q 1.15 25 Off Faired E3 Tunnel 30.5 0.251 Cp t .21 34 Off Faired O Tunnel 30.5 0.250 .5 -0 1.0 J 1 J 1 , [ , I _ 1 , I , I , J 45 90 135 180 225 270 315 360 0, deg -1.0 -3 FSI07 0 FS 253 0 O O O O -2 -.5 o_% o o_ °_bo DO 0 0 O0 0 000000 0 Cp Cp -I { }0 000 .5 1.0 ___ _ 1 , I , I L I _, l _ 1 _ I , I , t , [ ____1 t_ , _ 0 45 90 135 180 225 270 315 360 -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, deg y/s -3 -1.0 FS 296 FS 142 O O O -2 -.5 0 0 - C o -1 Cp 0 0 0 0 0 .5 , [ , I , I , 1 L I , ] , I , I 1 , I , 1 , _1 J 1 , I _ I L _1 1.0 -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 45 90 135 180 225 270 315 360 8, deg y/s -I.0 FS 184 FS 357

3f

-2 -.5 Cp @_000_ O 0 .5 '_ Cp -10 I O 0 0 0 0 0 , I _ 1 • t US_ , 1 , I , I , 1 ] , I , 1.0 -1.00 -.75 -.50 -.25 0 .25 .50 .75 [ .00 0 45 90 135 180 225 270 315 360 O, dog y/s (a) Porebody. (b) LEX.

Figure 29. Effect of wing leading-edge flap deflection on surface pressure coefficients measured on F/A-18 aircraft CFD wind tunnel model at c_ ._ 30 °.

| !!11 -1.0 FS 85 Data _,deg M_ Re xl0 _ 8f, deg Tails Inlet -.5 0 Tunnel 19.2 0.341 1.54 0 On Faired 0000000000000 25 On Faired 0 Tunnel 19,I 0.342 1.54 Cp 1.61 34 On Faired O Tunnel 19.2 0.342 0 O0 .5 1.0 0 45 90 135 180 225 270 315 360 0, deg -I .0 - -3 - FS 107 FS 253 -2 -.5 O O O ooOWOOOooooooOO °_bo o OO Cp Cp -1 <30 0 0 }00 OO o 0 0 0 0 .5 1.0 I 1 l t_ _ _ a 1 _ L _ 1 I _ J 0 .25 .50 .75 1,00 0 45 90 135 180 225 270 315 360 -1.00 -.75 -.50 -.25 0, deg y/s -3 -1.0 F- t FS 142 f FS 296 _ _ 5 -2

F

iot:t_o o o o oooc _ 000_000 o o Cp -1 Cp 0 0 0 foOO Oo O IlIFI I. _J ± _ i l t__L _ l_ -I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0 "45 90 135 180 225 270 3t5 360 0, deg y/s -I .0 -3- FS 357 FS 184 -2 -.5 0 88 o Oo_oO Cp Cp -1 }0 OO 0000_0 .5 0 0 0 0 0 0 1.0 _ I , 1 , i _ L c 1 L 1 _ 1 -.75 -.50 -,25 0 .25 ,50 .75 1.00 0 45 90 135 180 225 270 315 360 - 1.0_ O, deg y/s (a) Forebody. (b) LEX.

Figure 30. Effect of wing leading-edge flap deflection on surface pressure coefficients measured on F/A-18 aircraft CFD wind tunnel model with empennage at a _ 19 °.

-1.0 FS 85 Data or. deg M_ R,2 ×10 6 8f, deg Tails lnlel -.5 _00 O0 O Tunnel 25.8 0.250 1.17 0 On Faired 0 00000 O0 [3 Tunnel 25.8 0.249 1.15 25 On Faired Cp © Tunnel 25.9 0.250 1.21 34 On Faired 0 0 o .5 _ 1.0 • [ , A_ 1__1 I , I , r 45 90 135 180 225 270 3 [ 5 360 0, dcg -1.0 -3 FS 107 FS 253 O -2 -.5 O 0 0 0 O0 OO oOdm°OOooooooO_ °_bo ° 0 0 C_ 0 Cp -1 Cp 0 000 .5 0 1.0 , 1 _ I , 1 , 1 _ 1 , I , t _ ] 1 J I , I , I , [ ___L_ _ I , I J I 0 45 90 135 180 225 270 315 360 -I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, dcg y/s -3 - -I.0 FS 142 FS 296 o -2 O _' -.5 o o oo_ 0 0 Cp -I - Cp 0 000 I .5 0 1.0 , 1 , I , I , I _ !.=L._ t l 1 ___ L_ I , I , 1 t I , I , 1 0 45 90 135 180 225 270 315 360 -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 O. deg yls -1.0 - -3 - FS 184 FS 357 -2 -.5 6 8 _ Cp -1 Cp O066_)6 }O O0 I .5 0 0 0 0 0 0 1 _J I , I , t , I , i _ I __ 1.0 _1 , I , I , I , 1 _ 1 , I , I -I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0 45 90 135 180 225 270 315 360 0, dog y/s (a) Forebody. (b) LEX.

Figure 31. Effect of wing leading-edge flap deflection on surface pressure coefficients measured on F/A-18 aircraft CFD wind tunnel model with empennage at (_ _ 26 °.

:1 lli -1.0 FS 85 R_.. × 10 _ 8f, deg Tails Inlet Data ¢x, deg M_ -.5 000 000 0 0000 ° 0 o Tunnel 30.6 0.251 1.18 0 On Faired D Tunnel 30.5 0.250 1.15 25 On Faired Cp O Tunnel 30.5 0.251 1.21 34 On Faired .5 1.0 0 45 90 135 180 225 270 3t5 360 0, deg -3 -I .0 FS 107 0 FS 253 O O O O O -.5 -2 o_P°_o o °_b° _o 0 0 O0 o 0 0000000 0 Cp -1 Cp O0 o "5_3 O 0 1 , 1 , 1 , I , I 1 _L _ _J 1.0 _ _.1 t 1 t _l t_ _ 1 _1 -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0 45 90 135 180 225 270 315 360 0, dog y/s -1.0 - -3 - FS 142 FS 296 0 0 -.5 0 o O000¢KP -2 -0_00 0 0 Cp -I Cp 0 _ O0 .5 I _ _l_. 1 , 1 1 1L 1 _ _1__± I _ • t _1 1.0 45 90 135 180 225 270 315 360 -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, dog y/s -3 - -I.0 FS 357 FS 184 -.5 -2

Oo o oO

Cp -i _0 Cp 9_eeoeo ° °oo¢_ O o .5 - O iO O O O O 1.0 I 0 .25 .50 .75 1.00 45 90 135 180 225 270 315 360 -1.00 -.75 -.50 -.25 0, deg y/s (a) Forebody. (b) LEX.

Figure 32. Effect of wing leading-edge flap deflection on surface pressure coefficients measured on F/A-18 aircraft CFD wind tunnel model with empennage at a _ 30 °.

-1.0 FS 85 -.5 Dala _, deg M_ R e xlO ++ _t , deg Tails Inlet Flight 19.1 0.300 I 1.50 25 On Open o+ _ .°++ o.oi+g,o*g, Cp Tunnel 19.1 0.342 1.54 25 On Faired _i_°Ooq_ _o o° t .5 1.0 _L_I , 1 , I + I , I _ I , I___ 45 90 135 180 225 270 315 360 0, dcg -3 FS 253 -1.0 _- FS 107 -.5 o_°% _,°o° Cp _o .5

0Ill

0 0 0 0 0 1.0 t .J l + I , l , I t I t L_I I , I L I t I L t _ l , I , I , l 0 45 90 135 180 225 270 315 360 -1.00 -.75 +,50 -.25 0 .25 .50 .75 1,00 0, deg y/s -! .0 -3 FS 142 FS 296 -.5 -2 O O 0 ° 0 o _ _ _ _ +"_i_m_+__ _ i_ ii_ _ i:° v i_ °_ "_-qm_ Cp Cp -I ,o v i_i _ • 0 (J_© .5 OO 0 0 0 O0 0 1.0 ----L_--_- _--L--L_ I , l +--_ l,l,l,l,=_l,l,l,l,l 45 90 135 180 225 270 315 360 -I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, deg y/s -I.0 -3 FS 184 FS 357 -.5 -2 _o _o o e ,-"_,_i_o Cp

Cp -I %0+,+

• 0 o ,5 0 © o o0o %0 _ , 1.0 +_ 1_ + I , I J+ I , ! , [ _ I I ___ I , I , I , I , I , I , I , I 45 90 135 180 225 270 315 360 - 1.00 -.75 -.50 -.25 fl 2_ .5Cl .75 ! .00 0, deg y/s (a) Forebody. (b) LEX.

Figure 33. Correlation of surface pressure coefficients from flight and CFD wind tunnel model tests with empennage at c_ _ 19 °.

! ;!!_ -I.0 - FS 85 Data ct, deg M_ RexlO 6 _1, deg Tails Inlet -.5 34 On Open Flight 25,8 0.253 [0.80 oo • _oo_o,o,, _o 34 On Faired Tunnel 25.8 0.250 I.I7 Cp o o o o ,_,o %_ 45 90 135 180 225 270 315 360 O, deg -3 -1.0 FS 253 FS 107 o_o °_o -2 o -.5 oo.

_o O _o • Cp -1 Cp O O .5 o o o o o o o o ___L_ _ _t 1 , I I I 1.0 -I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0 45 90 135 180 225 270 315 360 y/s 0, deg -1.0 - FS 296 FS 142 _o ° -2 -.5 O o o ©, t_ O o Cp -I Cp ,5 - o o o o o o o o I 1,0 A L_I _ I i F _ i ] i 1 k_ 0 .25 .50 .75 1.00 -I.00 -.75 -.50 -.25 0 45 90 135 180 225 270 315 360 0, dcg y/s -3 -1.0 - FS 357 FS 184 _o eo -2 -.5 o • _ o_°__o 0 Cp -I Cp .5 o i_ o o o o o o • • • • - • I __[ _ _- _LI _ • t_ • _1__ 1.0 __ A___J _ L _ 1_ i I , I _ l __.1 -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 45 90 135 180 225 270 315 360 y/s O, deg (a) Forebody. (b) LEX.

Figure 34. Correlation of surface pressure coefficients from flight and CFD wind tunnel model tests with empennage at a _ 26 °.

-1.0 - FS 85 Data o< deg M_ Rcx I 0 _ 8f, deg Tails Inlet -.5 0%.

0 • 10 00 O o, i_%o,looi_ "o Flight 30.3 0.243 10.20 34 On Open 0 Tunnel 30.6 0.251 1.18 34 On Faired Cp o o o o o .5 _ °Ib 0 1.0 __L L .L_[_ J t ± 1 -L L_ 0 45 90 135 180 225 270 315 360 0, dog -I .0 - -3 o • FS 107 o o o_O FS 253 {OlD -.5 -2

o

o _ • _o • 8 ° Cp Cp -I o o o • v% • .5 0 0 0 0 0 0 0 0 1.0 I , 1 L 1 K 1 _ I , I , I , I 1 __I t _ _ l_ _ 1 _ I , ! __ _ _J 45 90 135 180 225 270 315 360 -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, dcg y/s - 1.0 FS 142 FS 296 _oo o O o°il o o o_ -2 • -,5 O c_

. O

8 o @o -1 Cp Cp i .5 0 0 0 0 0 0 0 0 0 1.0 , I t I , I , I t I , I _ I J I I _ 1 _ 1 _ t _____ , 1 _ I 0 45 90 135 180 ,_._" 270 315 360 -I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, dog y/s -1.0 - _ FS 184 FS 357 -.5

3[

_ '_o 0 _ o ° Cp .5' • O• -1(} _i' 0 o o• o ° 1.0 _j I L I , 1 _ I , I t I , I , I [ _ 1 n I_ ___1 _ J , I , 1..._ 0 45 90 135 180 225 270 315 360 -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, deg y/s (a) Forebody. (b) LEX.

Figure 35. Correlation of sm'face pressure coefficients from flight and CFD wind tmmel model tests with empennage at ct _ 30 °.

II I1 -I.0 - FS 85 Data oc, deg M_ R e ×10 _' ,fir, deg Tails Inlet -.5 CFD 19.0 0.340 13.50 0 Off Faired CFD 19.0 0.340 1.45 0 Off Faired Cp .5 1.0 0 45 90 135 180 225 270 315 360 0, deg -1.0 FS 107 -3 _ FS 253 -.5 -2 Cp .5 I t 1 t_ 1.0 _ , l , [ t _ , I • 1 _l -I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0 45 90 135 180 225 270 315 360 0, deg y/s -I.0 -- -3 - FS 142 FS 296 -2 -.5 Cp -1 Cp .5 1 J [ _ I J 1 , l L J _, ,_ l 1.0 0 .25 .50 .75 .00 45 90 135 180 225 270 315 360 -1.00 -.75 -.50 -.25 0, dcg y/s -3 - -1.0 FS 357 -.5 -2 0 Cp -I Cp .5 I 1.0 _ _1 _± A _ 1 __1 _ [ _ J 0 .25 .50 .75 1.00 0 45 90 135 180 225 270 315 360 -1.00 -.75 -.50 -.25 0, deg y/s (a) Forebody. (b) LEX.

Figure 36. Effect of Reynolds number on computed surface pressure coefficients at a _ 19 °.

-1.o FS 85 Data _, deg M_ R_ x l0 6 _1, deg Tails Inlet -.5 CFD 19.0 0.340 13.50 25 Off Faired Tunnel 19. I 0.341 1.51 25 Off Faired Cp I .5 1.0 45 90 135 180 225 270 315 360 0, deg -3 -1.0 [- FS 107 FS 253 _ 2 -.5

P

Cp Cp o .5 _, I 1.0 ___A _ I a I ___L , l _ I__ -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0 45 90 135 180 225 270 315 360 0, deg y/s -3 -1.0 F FS 142 f FS 296 -2 I _ 5

V

Cp -I Cp .5 f 1.0 J._._ _ 1 , I 1 0 45 90 135 180 225 270 315 360 -I.00 -.75 -.50 -.25 0 .25 .50 .75 .00 0, dcg y/s -3 - 1.0 - FS 357 FS 184 o -2 oOo -1 Cp Cp .5 0 1.0 __J , 1 J 1 _ L _ .l • l_J_._ I 0 45 90 135 18(1 225 270 315 360 -I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, deg y/s (a) Forebody. (b) LEX.

Figure 37. Correlation of surface pressure coefficients from computed and CFD wind tunnel model tests at a _ 19 °.

"I • I -1.0 - FS 85 Data _, deg M_ Re × I 0 _' 5f, deg Tails Inlet -.5 CFD 25.8 0.253 10.80 25 Off Faired Tunnel 25.8 0.251 I. 15 25 Off Faired Cp .5 1.0 45 90 135 180 225 270 315 360 0, deg -1.0 -3- FS 253 FS 107 -.5 © © -1 Cp Cp .5 1.0 , _ L.__ I _L [ , • L [ I 0 .25 .50 .75 1.00 45 90 135 180 225 270 315 360 -1.00 -.75 -.50 -.25 0, deg y/s -3 -I .0 - FS 296 FS 142 -.5 -2 Cp -1 Cp ( .5 -- 1.0 ____ • 1_.1 • _ I _ I _ • , 1 ,1 -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0 45 90 135 180 225 270 315 360 0, deg y/s -1.0 -- -3 FS 357 FS 184 O ©O -2 -.5 Cp -I Cp ° o C

o

.5 0 1.0 ___ _ 1 __.1 t 1 , I , I I 0 .25 .50 .75 1.00 0 45 90 135 180 225 270 315 360 - 1.00 -.75 -.50 -.25 0, deg yls (a) Forebody. (b) LEX.

Figure 38. Correlation of surface pressure coefficients from computed and CFD wind tunnel model tests at a _ 26 °.

-I.0 - FS 85 Data ct, deg M_ R e × I 0 6 _f, deg Tails Inlet CFD 30.3 0.243 10.20 25 Off Faired Tunnel 30.5 0.251 1.15 25 Off Faired Cp -.50 .5 1.0 J 1 , 1 ± • , I , I , .1___1 0 45 90 135 180 225 270 315 360 0, deg -1.0 - -3 FS 107 /_ FS 253 /_ -.5 Cp .5 / 1.0 I , I , 1. _ I , I , I L 1 _ I 1 1 , I , L_,_ 1 , i , 1 , t , 0 45 90 135 180 225 270 315 360 -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, deg y/s -3 - -1.0 FS 296 FS 142 © o © -2 ___ Cp -I Cp -.50( .5 1.0 _ I , I , I l I , I , I , I , I 1 _t L , 1 _ I _ I , I , I _ I 0 45 90 135 180 225 270 315 360 -I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0, deg y/s -I.0 - -3 FS 184 FS 357 © -2 -.5 Cp Cp -1 © ©o .5 0

o- o

1.0 I , / l l , I _ I , I , I , I -I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0 45 90 135 180 225 270 315 360 0, dog y/s (a) Forebody. (b) LEX.

Figure 39. Correlation of surface pressure coefficients from computed and CFD wind tunnel model tests at a _ 30 °.

:I II -1.0 - FS 85 Data _, deg M_ R e ×10 a tSf, deg Tails Inlet -.5 CFD 30.3 0,243 10.20 25 Off Faired CFD 25.8 0.253 10.80 25 Off Faired Cp 0 CFD 19.0 0.340 13.50 25 Off Faired 0 Tunnel 30,5 0.251 1.15 25 Off Faired .5 [] Tunnel 25,8 0.251 1.15 25 Off Faired 0 Tunnel 19.1 0.341 1.51 25 Off Faired 1.0 45 90 135 180 225 270 315 360 O, deg -1.0 -3 FS 107 k _ FS 253 -2 Cp -I Cp -.50_ { .5 0 1.0 __ 1_ , • t [ k_ _ 1

0 45 90 135 180 225 270 315 360 o 5o ,0o

0, deg y/s -3- -1.o - FS 142 FS 296 [ 0 0 0 n -.5 Cp -1 Cp 0 .5 0 " "_ --- -0- _ 1 __ _1 _ • _ [ _ I 1.0 __L__L__._ 1 __ ___ •_ • • J -I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 0 45 90 135 180 225 270 315 360 0, deg y/s -1.0 _- FS 184 -.5 Cp -I Cp 0 ,5 1.0 _ A_ l _ 1_ [ _ 0 45 90 135 180 225 270 315 360 0, deg y/s (a) Forebody. (b) LEX.

Figure 40. Effect of angle of attack on computed and measured surface pressure coefficients for CFD wind tunnel model with 6f = 25 °.

-1.0 FS 85 Dala oL deg M_ R e x I[) " 81 , deg Tails lnlel -.5 13.50 0 Off Faired CFD 19.0 0.340 13.50 25 Off Faired CFD 19.0 0.340 Cp 1.46 0 Off Faired Tunnel 19.1 0.343 i 1.51 25 Off Faired Tunnel 19.I 0.341 .5 1.0 l • L____ 1___ A_ , 1 , 1 45 90 135 180 225 270 315 360 0, deg -3 -1.0 FS 253 FS 107 -2 -.5 Cp Cp I L I _k i 1 L_L.t_ 1 , 1 , 1_ 1.0 -I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 45 90 135 180 225 270 315 360 0, deg y/s -I .0 FS 142 -3 _- FS 296 -2 -.5 Cp -1 Cp -.-4:3. __jd-J- .5 1 t 1__ I _ 1_ 1 _1 b 1.0 , _ I , 1 _ 1 , _, 1 , I E I - 1.00 -.75 .50 -.25 0 .25 .50 .75 1.00 45 90 135 180 225 270 315 360 0, dcg y/s -I .0 FS 184 -3 _ FS 357 -.5 -2 t- [] ¢]n n Cp .5 1.0 _x I _ J _ t_ l t _ 1 _ I -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00 45 90 135 180 225 270 315 360 y/s 0, dcg (a) Forebody. (b) LEX.

Figure 41. Effect of wing leading-edge flap deflection on computed and measured surface pressure coefficients for CFD wind tmmel model at c_ _ 19 °.

:_! I I .4 .3 Data M_ R e ×10 6 5f, deg Tails Inlet .2 O Tunnel 0.34 1.46 0 Off Faired • CFD 0.34 1.45 0 Off Faired

Cm

• CFD 0.34 13.50 0 Off Faircd .1 -.1 I I _ 1 _ I , I , I 2.0 -- 1.5 1.0 CL .5 I .1_ I _ I t I _ I _ I J_ I _ t , I 1 1 -,5 I -5 5 15 25 35 45 -.4 0 .4 .8 1.2 i .6 _, deg CD Figure 42. Effect of Reynolds number on computed longitudinal aerodynamic characteristics and correlation with measurements on CFD wind tunnel model at a _ 19°.

0.34 0.25 0.24 M O = .4-

l'

.3 Data M_ R e x 10 -_ 8f, deg Tails lnle!

o o .2 o Tunnel 0.30 --1 25 On Open

|

O O C m 0 • Tunnel M o =1 25 Off Faired • CFD M o =10 25 Off Faired .!

OOOO °°(_Oo -.! 1 i 1 i..: I i 1 1 J __ I 2.0 B oO(_ °° 1.5

o¢ii°

0110o1

o| °

Ol

1.0 O CL .5 O 0 O 0 0 0 0 0 -.5 , I i 1 i I i I , I i 1= I I E 1 i I ±=__1 -5 5 15 25 35 45 -.4 0 .4 .8 1.2 1.6 or, deg CD Figure 43. Measured and predicted longitudinal aerodynamic characteristics for baseline F/A-18 aircraft and CI?D wind tunnel model.

Form Approved REPORT DOCUMENTATION PAGE OM8 No. 0704-0188 Public reporting burden for this collection of information is estimated to average I 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 (0704-018B), Washington, DC 20503 1. AGENCY USE ONLY(Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED December 1994 Technical Paper 4. TITLE AND SUBTITLE S. FUNDING NUMBERS Navier-Stokes, Flight, and Wind T_ume] Flow Analysis for the F/A-18 Aircraft WU 505-68-30-03 6. AUTHOR(S) Farhad Ghaffari B. PERFORMING ORGANIZATION 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) REPORT NUMBER NASA Langley Research Center Hampton, VA 23681-0001 L-17336 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORING/MONITORING AGENCY REPORT NUMBER National Aeronautics and Space Administration NASA TP-3478 "Washington, DC 20546-0001 11. SUPPLEMENTARY NOTES 12b. DISTRIBUTION CODE 12a. DISTRIBUTION/AVAILABILITY STATEMENT Unclassified Unlimited Subject Category 02 Availability: NASA CASI (301) 621-0390 13. ABSTRACT (Maximum 200 words) Computational analysis of flow over the F/A-18 aircraft is presented along with complementary data from t)oth flight and wind tmme] experiments. Tile computational results are based on the three-dimensional thin-layer Navier-Stokes formulation and are obtained from an accurate surface representation of the filselage, leading-edge extension fLEX), and the wing geometry. However, the constraints imposed by either the flow solver and/or the complexity associated with tile flow-fieht grid generation required certain geometrical approximations to be implemented in the present numerical model. In particular, such constraints inspired the removal of the empennage and the blocking (fairing) of tile inlet face. The results arc computed for three different free-stream flow conditions and compared with flight test data of surface pressure coefficients, surface tuft flow, and off-surface vortical flow characteristics that included 1)reakdown phenomena. Excellent surface pressure coefficient correlations, both in terms of magnitude and overall trend, are obtained on the forebody throughout the range of flow conditions. Reasonable pressure agreement was obtained over tile LEX; the general correlation tends to improve at higher angles of attack. The surface tuft flow and the off-surface vortex flow structures compared qualitatively well with the flight test results. To evaluate the computational results, a wind tunnel investigation was conducted to determine the effects of existing eonfigurational differences between the flight vehicle and the numerical model on aerodynamic characteristics. In most ea._es, the geometrical approximations made to the numerical model had very little effect on overall aerodynamic characteristics.

14. SUBJECT TERMS 15. NUMBER OF PAGES F/A-18 aircraft; Computational fluid dynamics; Navier-Stokes analysis; Flight data; Wind tunnel data; High angles of attack; Vortical flows; Vortex breakdown 16. PRICE CODE A04 20. LIMITATION 17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION OF ABSTRACT OF REPORT OF THIS PAGE OF ABSTRACT Unclassified Unclassified Unclassified NSN 7540-01-280-5500 ;tandard Form 298(Rev. 2-89) Prescribed by ANSI Std Z39-18 298-102

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