NASA Technical Paper 3478

Navier-Stokes, Flight, and Wind Tunnel FlowAnalysis for the F/A-18 Aircraft

Farhad GhaffariLangley Research Center • Hampton, Virginia

National Aeronautics and Space AdministrationLangley Research Center • Hampton, Virginia 23681-0001

December 1994

https://ntrs.nasa.gov/search.jsp?R=19950012699 2020-02-07T02:01:48+00:00Z

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:

NASA Center for AeroSpace Information

800 Elkridge Landing Road

Linthicum Heights, MD 21090-2934

(301) 621-0390

National Technical Information Service (NTIS)

5285 Port Royal Road

Springfield, VA 22161-2171

(703) 487-4650

-1 :| i,

Contents

Summary .................................. i

Introduction ................................. 1

Symbols ................................... 2

Sources of Data ................................ 3

Flight Experiment ............................. 3

Computational Fluid Dynamics ........................ 4

Surface grid .............................. 4

Flow-ficld grid ............................. 5

Computational methodology ....................... 5Method performance and convergence ................... 6

Wind Tunnel Experiment .......................... 6

Results and Discussion ............................ 7

CFD Versus Flight Data ........................... 7General flow features .......................... 7

Surface pressures ............................ 9

Wind Tunnel Data ............................ 11

Effect of empennage ......................... 11Effect of inlet fairing ......................... 11

Effect of flap deflection ........................ 11

Flight Versus Wind Tunnel Data ...................... 12

CFD Versus Wind Tunnel Data ....................... 12

Reynolds number cffcct ........................ 12

Surface pressures ........................... 13

Effect of flap deflection ........................ 13Forces and moments ......................... 14

Concluding Remarks ............................ 15

Rcferenccs ................................. 16

Figures .................................. 18

o,o

III

1 1-1

Summary

Computational analysis of flow over the F/A-18

aircraft is presented along with complementary datafrom both flight and wind tunnel experiments. The

computational results are based on the threc-

(timensional thin-layer Navier-Stokcs formulationand arc obtained from an accurate numerical model

of the fuselage, the leading-edge extension (LEX),and the wing geometry. However, tile constraints

imposed by the flow solver and/or the complex-

ity associated with the flow-field grid generation

required certain geometrical approximations to beimplemented in the present numerical model. In par-

ticular, such constraints from tile flow solver inspired

the blocking (fairing) of the inlet face, wifich then

precluded the propulsion effects. The grid generationcomplexity required the removal of the empennage.

The results are computed for three different

free-stream flow conditions and compared with flight

test data for surface pressure coefficients, surface tuft

flow, and off-surface vortical flow characteristics that

included breakdown phenomena. Excellent surfacepressure correlations, both in terms of magnitude and

overall trend, are obtained on the forebody through-

out tile examined range of flow conditions. Reason-

able pressure agreement w_ found over the LEX; thegeneral correlation tends to improve at higher anglesof 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 theaerodynamic effects of existing configurational dif-

ferences between the flight vehicle and the numerical

model. This study revealed that in most cases, thegeometrical approximations made to the numerical

model had very little effect on overall aerodynamic

characteristics. Furthermore, to validate the latter

wind tunnel results at flight flow conditions, a com-putational study was conducted to determine the

aerodynamic influence of differcnces in the Reynolds

number. This computational study, which was per-

formed on exactly the same grid, showed that anorder-of-magnitude difference between the flight and

wind tunnel Reynolds numbcrs produces negligible

effects on the forcbody and the LEX surface pres-

sures as well as the longitudinal aerodynamic char-

acteristics. Vcry good surface pressure correlationbetween wind tunnel and flight data was obtained

on the LEX; however, the wind tunnel pressure data

appear to be slightly below thosc of the flight mca-

surement on the forebody, particularly at high anglesof attack.

Introduction

Combat aircraft arc often asscsscd on their high-

angle-of-attack aerodynamic performance for achiev-

ing superior levels of sustained maneuverability andagility. At high attitude, the flow characteristics over

these often geometrically complex aircraft configura-

tions generally become very complicated and (tiffi-cult to predict and control. One such flow charac-

teristic is the inevitable vortical flow precipitated by

flow separation that occurs when the aircraft oper-

ate at high angles of attack. In general, the pres-ence of such vortical flow over an aircraft surface can

be advantageous as long as it remains organized and

stable; this flow produces vortex lift, which can en-

hance maneuverability, ttowever, with increasing an-

gle of attack, such a coherent vortex system is sus-

ceptible to instabilities such as breakdown or flowasymmetry, which cause undesirable pitching, yaw-

ing, and/or rolling moment characteristics. As a

result, the flight-handling quality and controllability

of these aircraft are adversely affected by such flowphenomena; the ability of the aircraft, to maneuver

with high agility is often limited. The fundamental

understanding and predictability of flow phenomena

for a wide range of flow" conditions are of paramountinterest from both research and real aircraft design

perspectives.

The vortical flow formation or, in a more gen-eral form, the initial flow separation can bc classified

into two types. The first type is a flow separation

that primarily occurs over a smooth surface geome-

try because of an adverse pressure gradient interact-ing with the boumtary layer. Fhfid viscosity provides

the essential mechanism for this type of flow sepa-

ration to occur; this suggests that its formation is

highly sensitive to the local flow" Reynolds number.A typical flow separation of this type occurs over

a conic forebody at high angles of attack. Unlike

the first, the second type of flow separation occursat, and is primarily induced by, a surfaee discontinu-

ity such as the sharp leading edges of a delta wing.

Because of diminished sensitivity to the fluid viscos-

ity, the latter type of flow separation, fixed at thesurface discontinuity, is generally considered to be

insensitive to the local flow Reynolds number. In

this paper, the first and second type of flow sep-

aration is referred to as boun(tary layer and sharp

edge flow separation, respectively. In recent years,

various numerical (refs. 1 and 2) as well as experi-mental (refs. 3 6) research efforts have been made to

quantify the effects of Reynolds number on different

types of flow separation. The results from these stud-

ies are particularly important in providing insightinto the flow physics and the triggering mechanisms

responsiblefor thesubjectflowseparations.Specifi-cally,thebasicknowledgelearnedfromtheseinvesti-gationsmayleadto thedevelopmentof newcompu-tationaland/or experimentaltechniquesapplicableto a conventionalwind tunnelenvironmentfor sim-ulatingthehighReynoldsnumberflowencounteredbytheflight vehicle.

Uniqueresearchis presentlybeingconductedbyNASAwittlin tile High-Angle-of-AttackTechnologyProgram(HATP)(ref.7), whichhasasits majorob-jectivethe explorationof high-attitudeflowcharac-teristicsoveratypicalfighteraircraftduringmaneu-veringsituations.At theoutset,theF/A-18aircraftwaschosento bc the baselineconfigurationprimar-ily becauseof its high-angle-of-attack(i.e.,high-a)capability.Subsequently,anF/A-18 aircraft,desig-natedasttle High-AlphaResearchVehicle(HARV)(ref.8),washighlyinstrumentedforsurfacepressuremeasurementsaswellas in-fligtlt flowvisualization.(Seerefs.9 and 10.) On- andoff-surfaceflowvisu-alizationshavebeenconductedoil tile HARV;differ-ent techniqueshavebeenusedandincludeoneinno-vativeapproaci_to documentsurfaceflowpatterns.(Secref. 9.) In additionto the flight experimen-tation, HATP is utilizing ground-basedfacilitiestoacquireexperimentaldatafromvariouswindtunnelscalemodels(refs.11 15)aswellasa full-sizecon-figuration,whichtinsbeentestedin theAmes80-by120-FootWindTunnel.(Seercf. 16.)Theresultsoftheseexperimentshaveprovidedthedatabaseneededfor tim development and validation of the present

computational fluid dynamics (CFD) methodologies.

In the past few years, significant progress has

been made to fulfill the HATP objectives by pro-

viding detailed analyses of tfigh-angle-of-attack flow

over tim F/A-18 aircraft. In particular, the numer-ical analyses based on the thin-layer Navier-Stokes

formulation have made important contributions. Ini-

tial computational activities started in parallel at the

Langley and Ames Researctl Centers. Two different

approaches were taken to solve the flow over the iso-

lated F/A-18 aircraft forebody which included theleading-edge extension (LEX) geometry. Although

both approaches were based on multiblock structured

grid strategies, one method used a nonoverlapping

grid-block approach (refs. 17 and 18), and the othermett_od was based on an overset grid or Chimera

(ref. 19) approach. (See ref. 20.) Subsequently, both

computational methods were expanded to include

more of the F/A-18 aircraft geometrical components

such as the wing, aft portion of the fuselage, and em-

pennage (i.e., vertical and horizontal tails). Theseefforts were very successful and the computational

results with both the multiblock overset (refs. 21 23)

and nonoverlapping grids (ref. 24) are documented.

The primary objectives of the present investiga-

tion are summarized into the following five categories:

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

tions (ref. 24) for the F/A-18 aircraft configu-ration to include a wider range of flight flowconditions

2. Correlate computational results with flight testdata for both on- and off-surface flow character-

istics and surface pressure coefficients

3. Evaluate aerodynamic effects which result from

the geometrical differences between the flightvehicle and the numerical model through wind

tunnel experimentation

4. Correlate wind tunnel data with flight test re-

sults to assess the scale model simulation of high

Reynolds number flow; in addition, numericallyassess the effects of Reynolds number on the aero-

dynamic characteristics of the configuration

5. Correlate computational results with appropriate

wind tunnel data obtained on a configuration that

is more representative of the numerical model

Symbols

bref

CD

G

c.g,

Eo

F,G,H

J

l

reference wing span, 37.42 ft

Drag

drag coefficient, q_cSre f

Lift

lift coefficient, qocSref

pitching moment coefficient

Pitching momentreferenced to 0.25_,

q_cSrefC

pressure coefficient, p - P_qoc

wing mean aerodynamic chord,11.52 ft

center of gravity

total energy per unit volume

flux vectors

net flux

Jacobian of coordinate transformation

configuration body length measuredfrom nose to exhaust nozzle exit

plane, full scale, 54.4 ft

Mach number

free-stream Mach number

| |11

P

po

Po,I

Po,l/Po

Poc

Q

qoc

Re

Srd

8

t

"_l, 73, "03

V*

X

x/1

Y

y/s

y+

O_

5f

11

Y

0

(,71, ¢

P

T_ L,

Abbreviations:

CAD

CFD

C-O

static pressure

free-stream total pressure

local total pressure

normalized total pressure

free-stream static pressure

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

free-stream dynamic pressure, lb/ft 2

Reynolds number based on e

reference area of wing planform,

400 ft 2

LEX local-exposed semispan, in.

time, see

body-axis Cartesian velocity

components, ft/sec

wall friction velocity, V_P" ft/sec

axial distance from nose apex, in.

fraction of configuration body

length

distance along LEX local-exposed

scmispan, in.

fraction of LEX-exposed scmispan

inner-law variable, _v

angle of attack, deg

flap deflection angle, deg

viscosity, lb-sec/ff 2

kinematic viscosity, _, ft2/sec

azimuthal angle measured clockwise

viewed from front, dcg (0 ° located

at bottom dead center)

body-fitted coordinates

density, slug/ft 3

wall shear stress, lb/ft 2

computer-aided design

computational fluid dynamics

grid topology, C streamwise and Ocircumferential

FS fuselage station, full-scale, in.

HARV High-Alpha Research Vehicle

HATP High-Angle-of-Attack Technology

Program

H-H grid topology, H streamwisc andcircumferential

H-O grid topology, H streamwise and Ocircumferential

HST High-Speed Tunnel

IGES Initial Graphics Exchange

Specification

LEX leading-edge extension

MUSCL monotone upstream-centeredscheme for conservation laws

NAS numerical aerodynamic simulation

NASA National Aeronautics and SpaceAdministration

WT wind tunnel

A caret (^) over a symbol indicates scalingwith respect to the Jacobian J of the coordinatetransformation.

Sources of Data

Flight Experiment

The F/A-18 aircraft was chosen as a baseline con-

figuration for the HATP, primarily because of its

high-a capabilities. The aircraft (fig. 1), designatedtile High-Alpha Research Vehicle (HARV), is instru-

mented to measure surface static pressures over the

forebody and on both starboard and port sides of theLEX. The three views of the F/A-18 aircraft config-

uration are shown in figure 2. The figure also pro-vides the full-scale reference dimensions in feet. Fig-

ure 3 presents the plalfform of an F/A-18 aircraft

configuration and the corresponding cross-sectionalgeometry of the forebody and LEX fuselage stations

(FS) where tile surface pressures are measured. The

forebody surface pressure orifices were distributed as

a function of 0 at longitudinal FS 70, 85, 107, 142,and 184. These fifll-scale dimensions are in inches; for

reference, the nose apex starts at FS _ 60. When the

cross section is viewed by looking aft, tile azinmthal

angle 0 is measured clockwise from the windward

plane of symmetry of 0 = 0°; the LEX surface pres-sures were measured as a function of LEX-exposed

semispan y/s on both the upper and lower surfaces

at FS 253, 296, and 357. The semispan parameter

y/s = 0 corresponds to the LEX-fuselage juncture

and y/s = 1 corresponds to the LEX leading edge.

Negativevaluesof theparametery/s correspond tothe starboard side and positive values to the port side

of the configuration.

The aircraft is also equipped with a smoke-

generating system (refs. 9 and 10) designed to emit

smoke at appropriate locations along the forcbodyand the LEX apex for visualization of the off-surfacevortical flows as well as their interactions with one

another and/or with the neighboring aerodynamicsurfaces. In conjunction with the off-surface flow

visualization, surface tufts are used on the wing,

LEX, fuselage, and tails to assist in correlation ofthe off-surface flow interactions with the surface flow.

These in-flight flow visualization images arc recorded

by a camera located either onboard the aircraft or achase plane. Furthermore, a unique approach has

also been successfully used to capture the in-flight

surface flow pattern on the forcbody and the LEX.

(See ref. 9.) The data gathered from the flight exper-iments have been instrumental in helping researchers

understand the subject flow phenomena. All the

HARV flight experiments were conducted by NASAat the Dryden Flight Research Center. The flight

data arc obtained for a wide range of angle of attack,

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

Computational Fluid Dynamics

The primary objective of the present compu-

tational analysis is to expand the earlier compu-

tations (ref. 24) to include a wider range of flowconditions and a more comprehensive flow analy-

sis. Various HARV flight test results (e.g., on- andoff-surface flow visualization photographs, surface

pressure data) were initially examined to identifythose flow conditions which exhibited the most chal-

lenging flow characteristics to be simulated numeri-

cally. Several flout conditions (table I) werc identified

which clearly demonstrate the complexity associatedwith the overall flow structures (e.g., LEX vortices

with subsequent breakdown at high-_ conditions,

forebody vortices, forebody-LEX vortex interactions,

and/or stalled flow over the wing).

Table I. Selected Flow Conditions

! 034 / la5 ×10°5.s I / 10.s

10.2

_], dog2534

34

Surface grid. The surface patch definition of the

complete F/A-18 aircraft was obtained from a de-

4

tailed computer-aided design (CAD) description in aformat known as Initial Graphics Exchange Specifi-

cation (IGES). (See ref. 25.) These data were thenused to extract a high-density surface grid definition

in thc form of cross sections. The CFD database grid

consisted of approximately 30 000 points on the fuse-

lage defined at 60 cross sections and 16 000 points on

the wing defined at 20 streamwisc cuts. Althoughnot used in the present computations, accurate sur-face definitions of both horizontal and vertical tailswere also included in this database.

This database was subsequently used to generate

a suitable surface grid for Navier-Stokcs computa-

tions by using an Ovcrhauser function (ref. 26) for

the interpolation. This function has been shown toalleviate the oscillatory behavior inherent in other

widely used functions (e.g., splines) in the region

where grid point distributions are not uniform. The

final computational surface grid was composed of ap-proximately 18 000 points. The geometrical simplifi-

cations made to the configuration included the fair-

ing over of the inlet, splitter plate, diverter, and the

LEX slot. Figure 4(a) shows a body cross sectionat FS 401 and illustrates the simplifications made

to the splitter plate and the diverter cavity region.

Similarly, figure 4(b) shows a body cross section at

FS 441 and illustrates the closing of the LEX slot aswell as the fairing of the cavity region between the

inlet and the LEX lower surface. The latter simpli-

fication is made to a limited region to facilitate the

flow-field grid generation in that area. Except forthe simplified regions, the two typical cross sections

shown in figures 4(a) and 4(b) demonstrate the accu-

racy with which the computational grid (_100 grid

points/station) represents the surface geometry ofthe much finer initial CAD cross-sectional defini-

tion (_500 grid points/station) despite the use of

only about a fifth as many grid points. Two ortho-

graphic views of the final F/A-18 aircraft CFD sur-

face grid definition are shown in figures 5(a) and 5(b)to illustrate the overall grid resolution. Further-

more, the wing geometry is modeled with two dif-

ferent leading-edge flaps: the undeflccted and the

blended flap. The latter will be discussed in the next

paragraph.

The F/A-18 aircraft wing leading-edge flap de-

flection angle varies as a function of angle of attackand Mach number. For free-stream subsonic Mach

numbers (Moc _< 0.76), the aircraft control systemis programmed to vary the flap deflection angle lin-

early as a function of angle of attack according to

the relationship _1- = 34_/25.6. The maximum flapdeflection angle of 34 ° is reached when c_ = 25.6 °

and the flap angle remains constant for a _> 25.6 °.

7 :| i

As a result, the flap deflectionangles,whichcorrespondto the flow conditions(tableI) of thisinvestigationareasfollow: @ = 25° for c_ = 19 °

and af = 34 ° for c_ = 25.8 ° and 30.3 ° . To sim-plify the flow-field grid generation, the surface ge-

ometry of the wing deflected leading-edge flap was

approximated in this computation and is designatedas a blended flap. The principle behind the deflected

flap geometry modification was the preservation of

the wing-body intersection with the undeflected flap,

which permitted the same overall blocking strategyto be used for both undeflected and blended flap con-

figurations. This modification smoothly blended theinboard 15-percent semispan of the flap between the

deflected flap and the undeflected flap wing-body in-

tersect.ion. A nose-down front view of the F/A-18aircraft CFD surface grid definition with both the

blended flap (starboard) and undeflected flap (port)

is shown in figure 6. In addition, a closeup view of the

CFD surface definition is shown in figure 7 with shad-

ing to highlight tile surfaces of the baseline F/A-18aircraft that are simplified, blended, and/or modified

(e.g., inlet, diverter, splitter plate, and the inboard

section of the deflected wing leading-edge flap).

Flow-field 9rid. Tire selection of the flow-field

grid strategy is primarily dictated by the two dis-

tinct types of aerodynamic shapes of the F/A-18

aircraft configuration: a slender type, which con-sists of the front forebody-LEX geometry, and a

high-aspect-ratio type, which contains tile wing com-

ponent. An H-O grid topology is selected for the

slender part, whereas a C-O grid is chosen for

the high-aspect-ratio wing configuration. A unique

global grid strategy is then devised which appro-priately links various grid topologies while main-

taining tile grid quality. To illustrate the selected

global grid strategy, isometric far-field (fig. 8(a)) andnear-field (fig. 8(b)) views of the shaded F/A-18 air-

craft, surface are shown for the configuration maxi-

mum half-breadth plane along with the field grids in

the plane of symmetry. For clarity, the grid densityshown in tile figures has been reduced in both longi-

tudinal and radial directions. The flow-field domain,

which consists of about 1.24 million grid points, isdivided into five regions with each composed of one

or more topologically similar blocks. Ill figures 8(a)

and 8(b), the region boundary edges arc highlighted

with thick, solid lines and the corresponding blockinterfaces within each region are denoted by thick,dashed lines.

A side view of the flow-field grid for selected sur-

faces is shown in figure 9 to illustrate the various

regions from a different perspective. The figure de-picts tile overall three-dimensional far-field bound-

aries with the wing wedge-shaped region sectioned

out of the field domain. Again, the boundaries of

the various regions and the corresponding block in-

terfaces are highlighted with thick, solid and dashedlines, respectively. Three factors contributed to the

selection of the grid topology for each region: the

consideration of local geometry, the proper resolu-

tion of the expected flow structure, and the ap-

propriate grid connectivity between regions. The

selected topologies should provide good resolution

(refs. 17 and 18) of all edge flows (e.g., LEX andwing leading edges and wingtip) and juncture flows

(e.g., LEX-body, wing-body, and canopy-body).

The volume grid is generated with established

transfinite interpolation techniques (refs. 17, 24,

and 27) with sufficient normal clustering near the

surface to adequately resolve the laminar sublayer ofthe turbulent-boundary-layer flow for a typical flight

free-stream condition of c_ = 19°, Moc = 0.34, and

Rc = 13.5 x 10G. This grid produced an averagenormal cell size next to the wall of approximately

7.2 × 10-Gc, which corresponds to 9 + _. 3 for turbu-

lent computations; a laminar sublayer generally ex-

tends out to y+ .-_ 8.5. The radial far-field boundary

extends to about 7.6c. A downstream grid extension

is created by repeating the grids generated about thebase cross section aft to about 4.7c. No grid is gener-

ated for the face of the base geometry (i.e., open), noris the flow simulated between the interior surfaces of

the model geometry and the exterior flow-field do-main. Note that the flow-field grid structure is gen-

erally designed to be consistent with those of previ-

ous computational studies (refs. 17 and 28) on the

isolated F/A-18 aircraft, forebody-LEX configurationwhere tile structures had been found to have ade-

quate cell size next to the wall, radial grid stretching,

circumferential grid resolution, and far-field bomM-

ary locations.

Computational methodology. The computational

results have been obtained from an algorithm that

has been successfully applied to a variety of aero-

dynamic problems with both simple and complexconfigurations for a wide range of flow conditions.

(See refs. 2, 17, 18, 24, and 27.) The algo-

rithm (refs. 1, 17, 18, and 29 31) is based on tirecompressible, time-dependent, Reynolds-averaged,

Navier-Stokes equations, which are written in a

curvilinear coordinate system. The equations are

solved with a finite volume approach and are com-

posed in a conservative form as

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

The subscripts with a comma denote partial

differentiation, the subscript v identifies the viscous

terms,and the caret (") overthe vectorsindicatesscalingwith respectto theJacobianJ of the coor-dinate transformation. Details of these terms are in-

eluded in reference 17. In addition to the ideal gas

assumption in the present study, tile thin-layer ap-proximation of the governing equations is invoked

(i.e., Fv = (_v = 0) and thus accounts for viscous

flux terms only in the direction _ normal to the body.

Turbulence effects are accounted for by the notion

of eddy viscosity and conductivity. The algebraic

turbulence model developed by Baldwin and Lomax

(ref. 32) is used to evaluate the required turbulencequantities. For separated vortical flow regions, the

method introduced by Degani and Schiff (ref. 33)

is used to ensure that the proper turbulence lengthscales are used.

The integral form of the conservation equations

is represented by

N

where the time rate of change Ot of the state vec-

tor I_ within a cell volume dV is balanced by

the net flux f across the cell surface doe with the

unit normal ft. The convective and pressure flux

quantities are represented by the upwind-biased,

flux-difference-splitting approach of Roe (ref. 34),whereas tile shear stress and heat transfer terms

are centrally differenced. Tile monotone upstream-centered scheme for conservation laws (MUSCL) of

Vail Leer (ref. 35) is used to interpolate state vari-ables at the cell interfaces. A detailed discussion

on the algorithm development for interpolating the

mass, momentum, and energy across tile various

planar and nonplanar interfaces that separate the

grid blocks is presented by Biedron and Thomas inreference 31.

Method performance and convergence. All com-

putations are performed on the numerical aero-

dynamic simulation (NAS) Cray-2* computer, lo-cated at Ames Research Center. On this machine,

the algorithm requires approximately 20 #see periteration per grid point and about 100 million words

of memory. Starting from the free-stream flow condi-

tion, a typical solution converged in about 3000 itera-tions, which consumed about 20 hr of computer time.The 3000 iterations were sufficient to reduce the

residuals by a little more than 2 orders of magnitude

*Trademark of Cray Research, Inc., Minneapolis, MN

55402.

6

and to reduce oscillations in CL to a negligible level

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

Subsequent solutions for different angles of attack

were obtained by starting the computations from an

existing solution, which then generally reduced thecomputational time for a converged result by as much

as a half. Similar convergence rates are achieved for

the computations at higher angles of attack despite

the presence of vortex breakdown in the solutions.

The computations are performed without the use ofmesh sequencing or multigrid iteration. (See ref. 29.)

Wind Tunnel Experiment

Tile wind tunnel experiment was conducted in the

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

(See refs. 36 and 37.) This is a closed-circuit,

continuous-flow, atmospheric tunnel with a solid walltest 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 about4 x 106 ft -].

Tile wind tunnel testing was conducted with a

0.06-scale model of the F/A-18 aircraft configuration.

This wind tunnel model had been used in a previous

experimental investigation, and the results arc pub-lished in reference 12. The model was instrumented

for surface static pressure measurements at four sta-

tions on the forebody and three stations on the LEX

upper surface; LEX lower surface measurcments weretaken only at the last station FS 357. The surface

pressures were mcasured on both starboard and portsides of the aircraft to assess flow asymmetry. The

fuselage stations on the model at which the surfacepressures arc measured are identical to those of the

flight vehicle with the exception of the first forebody

station FS 70 where no model data were acquired.

The sting-mounted wind tunnel model (fig. 11) was

equipped with an internally mounted strain gaugebalance to me_ure the six component forces and mo-

ments. Furthermore, the forebody of the model wasequipped with two transition grit strips positioned

longitudinally across the windward plane of symme-try at 0 = 45 ° and 315 °. Based on the method of ref-

erence 38, the No. 180 grit was found to be adequate

for tripping the laminar boundary layer to a turbu-lent flow to simulate the flight Reynolds number flowcharacteristics at the conditions listed in table I.

Two different configurations of the model were

tested: the first was the baseline F/A-18 aircraft andthe second incorporated modifications representative

of the numerical model. The second configuration is

referred to as the "CFD wind tunnel model" (CFD

WT) from here on. The data obtained from the

:1 Ili

CFD wind tunnel model are used to assess the aero-

dynamic effects of empennage removal, inlet fairing,

and wing leading-edge flap deflection. Modifications

of the baseline wind tunnel model were patterned af-

ter the numerical representation. Dental plaster was

used for fairing over regions of the model such asthe inlet and diverter. Figures 12(a) and 12(5) show

the CFD wind tunnel model from two perspectivesand illustrate the various modifications of the base-

line F/A-18 aircraft wind tunnel model such as fair-

ing over the inlet, splitter plate, and diverter; closingof the gap between the deflected flap and fuselage;

and the removal of the empennage.

The CFD wind tunnel model was a good represcn-

tation of the numerical model except in the regionsof the blended flap and the splitter plate. Unlike the

numerical model, where surface modifications were

made to the inboard 15-percent scmispan of the flap

by a smooth blending of the deflected flap geometry

to the undeflected flap wing-body intersection, theCFD wind tunnel modcl did not incorporate flap sur-

face blending. However, the gap between the inboard

face of the deflected flap and the fuselage was closed

with a metal sheet. (fig. 12(b)) that vertically joinedthe two edges. This gap renmined closed for the de-

flected flap configurations of the present CFD wind

tunnel model. Furthermore, dental plaster was also

used to fair over the cavity region between the splitterplate and the fuselage but in a slightly different nmn-

ner than the numerical model approach (fig. 4(a)) in

which the lower part of the splitter plate was trun-cated. This difference between the numerical model

and the CFD wind tunnel model in the geometric

representation of the splitter plate and diverter cav-

ity regions is illustrated for a typical cross section at.FS 401 in the lower right corner of figure 12(b).

The measured wind tunnel data are corrected

for the effects of angle of attack, wall interference,

and model base. drag. The model support systemincorporated an a.ccelerometer to measure the an-

gle of attack and is subsequently corrected to ac-

count for tile balance and sting deflection under load.The wall interference effects are accounted for by

the principles of 51ockage (ref. 39) and jet bound-ary (ref. 40) corrections. The model base pressures

are measured and subsequently integrated to obtain

tile resulting force acting normal to the base planeof the model. This normal force is then subtracted

from the total axial force component measured by

the internally mounted strain gauge balance to ex-

clude the pressure drag caused by the local wake

flow on the base of the model. As a result, the

model base pressure drag contril)ution to the configu-ration total forces and moments is adjusted to corm-

spond to the free-stream static pressure Poc. Surface

pressure measurements are obtained with electroni-

cally scanned pressure (ESP) transducers; the over-

all accuracy of this system is about +0.1 percent ofthe full-load range, which is approximately equal to

4-0.03 lb/in 2.

Results and Discussion

CFD Versus Flight Data

General flow features. The normalized total pres-

sure Po,1/Po contours in various cross-flow planes aswell as the LEX and forebody vortex core stream-

lines (where applicable) computed at the selected

flow conditions (table I) are shown in figures 13(a)

13(c). The magnitude associated with each normal-ized total pressure contour is displayed with the ap-

propriate color bar. The normalized total pressure

function is used to highlight the viscous losses within

a separated flow structure such as a vortex. For tilesame purpose, this function has also been success-

fully used in an experimental investigation reported

in reference 41. The results shown in figures 13(a )

13(c) are all obtained with a fully turbulent bound-

ary layer model at flight flow conditions with the

blende(t flap configuration. The computations arcperformed for half the configuration, but the results

are presented for the fifll configuration by using the

mirror-image principle of symmetry. Although both

Mach and Reynolds numbers vary slightly, the com-

putational results presented in figures 13(a) 13(c) re-veal tile effects of angle of attack on the flow charac-

teristics. These figures, discussed in the subsequent

paragraphs, highlight the following three general flowfeatures and their interactions: LEX w)rtex system,

wing flow field, and forebody riot" field.

The normalized total pressure contours in fig-

ures 13(a) 13(c) clearly indicate the presence of awell-organized LEX vortex riot" structure up to theLEX-wing leading-edge juncture. Over this longitu-

dinal extent, the overall LEX vortical riot, structure

generally remains similar even as the angle of attack

is increased. At c_ = 19 °, the LEX primary vor-tex system appears to remain coherent and maintain

its tight core structure over the entire configuration

body length. However, with increasing angle of at-

tack, figures 13(b) and 13(c) illustrate that the LEX

vortex core region (highlighted by" the lower levels ofthe normalized total pressure, which signify higher

levels of viscous loss) expands dramatically aft of

about the wing root midchord. This sudden core

expansion in the LEX vortical flow system is gen-

erally associated with a phenomenon referred to asvortex burst or breakdown. The vortex breakdown is

7

oftencharacterizedby anabruptreductionin veloc-ity (particularlytheaxialcomponent)andtile lossofcohesivenesswithin thevorticalflowstructure.ThelattereffectisclearlydemonstratedbytheLEXvor-tex corestreamlinesat c, = 25.8 ° and 30.3 °. The

predicted location of the LEX vortex breakdown is

discussed later in conjunction with flight and windtunnel test results.

The normalized total pressure contours at

c_ = 19 ° clearly illustrate the separated flow region

over the wing upper surface. This massive flow sep-aration over tile wing is essentially a confined regionof retarded airflow with a chaotic behavior which is

discussed in reference 24. When the angles of at-

tack are increased to 25.8 ° and 30.3 °, the separated

flow region over the wing appears to move outboardand exhibits lower levels of viscous toss as indicated

by the higher levels of the normalized total pressure.One contributor to this flow change is tile expansion

of the LEX vortex flow which extends spanwise onto

tile wing with increasing angle of attack.

With increasing angle of attack, the flow within

the boundary layer over the smooth leeward side of

the forebody separates and leads to a well-organizedvortical flow structure as shown in figures 13(b)

and 13(e). At about the middle (fig. 13(b)) or im-

mediately aft (fig. 13(c)) of the canopy, the forebody

vortex migrates downstream into a region where its

trajectory becomes affected by the much stronger

neighboring LEX vortex system. The forebody vor-tex flow is initially drawn into the LEX-fuselage

juncture from which it is entrained outboard by the

LEX vortex system. Note that just aft of the wingLEX leading-edge juncture, the streamlines originat-

ing from the forebody vortex core split; some wraparound the LEX vortical flow and the rest interact

with the wing flow field.

The HAI/V in-flight photographs (ref. 9) of thetufts as well as the LEX primary vortex core smoke

visualization are presented in figures 14(a), 14(b),

and 14(e) for c_ _ 20 °, 25 °, and 30 °, respectively.Tile photographs clearly illustrate the LEX vortex

breakdown just ahead of the vertical tail at ct20 ° and its upstream progression with increasing

angle of attack. In addition, the tufts show thesurface flow patterns over the wing, LEX, fuselage

aft of the canopy, and the vertical tail. In general,

for this range of angle of attack, the tufts reveal

a fairly orderly flow pattern over the LEX up to

the LEX-wing leading-edge juncture. However, thetufts clearly indicate a chaotic flow pattern over the

wing and the vertical tail with some tufts standing

up off the surface; these disordered flow structures

are directly attributed to stalled flow over the wing

because of a massive flow separation and the LEX

vortex breakdown, respectively.

Computational results for LEX vortex core

streamlines superimposed on surface tuft flow pat-terns are presented in figures 15(a) 15(e) for the flow

conditions listed in table I. For the higher angles of

attack (i.e., a = 25.8 ° and 30.3°), the forebody vor-

tex core streamlines are also shown to highlight their

paths and influence on the overall flow structuresboth on and off the surface. Surface tuft flow pat-

terns are simulated computationally with tile method

of unrestricted streamline tracing introduced and dis-cussed in detail in reference 24. Unlike tile conven-

tional method of tracing the experimental surface oil

flows and tuft patterns, this new approach does not

impose the restriction that tile streamline calcula-

tions lie in a particular grid plane near the surface.

The method (ref. 24) has demonstrated the capability

of simulating surface flow patterns in regions of at-tached as well as separated flows and is particularly

applicable to a stalled flow environment. Becauseof the stalled flow characteristics over the wing, the

method applied here initiated tile particle tracing at

a grid plane slightly off the surface (_0.02 in. full

scale, which is _0.00014_:) where the flow velocity

magnitudes become sufficiently large to produce avisible tuft flow pattern within a reasonable number

of time steps. Note that the number of time steps

used in computing tile unrestricted streamline tracesis constant, which results in variable length traces be-

cause of the nommiform distribution of the velocity

magnitudes in a given flow region. Tile tuft flow pat-

terns of the model shown in figures 15(a) 15(c) quali-

tatively simulate those patterns including the stalledflow region over the wing observed on the flightaircraft.

The simulated tuft flow patterns appear to have

been influenced by the off-surface flow structures

such as the LEX vortices. This effect is particularly

evident over the wing in tile aft inboard region wheretile tufts indicate a spanwise flow pattern caused by"

the flow expansion around the LEX vortex break-down at o_ = 25.8 ° and 30.3 °. This flow expansionaround the LEX vortex breakdown and the result-

ing interaction with the wing flow field is also evi-dent in the computational results shown earlier in fig-

ures 13(b) and 13(c). The accuracy of the predictedlongitudinal location of the LEX vortex breakdown

as a function of angle of attack is discussed in the

next two paragraphs.

Several approaches, publieally available in the sci-entific literature, have been devised to locate the

vortex breakdown in a given flow structure. One

such method that has been widely investigated and

::1 |1

is adoptedheredefinesthe onsetof vortexbreak-downat a point in the corewherethe axial veloc-ity componentbecomeszero(i.e.,u = 0) orreversesdirection(i.e., u _<0) fromthat of the free-streamcomponent.(Seeref. 29.)Thepresentnumericalso-lutions are examined one cross-flow plane at a timeto determine the magnitude of u within the LEX

primary vortex core. By this analysis, no evidence

exists that a vortex breakdown occurs at (_ = 19°;

however, at a = 25.8 ° and 30.3 °, the LEX vortex

breakdown develops longitudinally at x/l _ 0.72 and

0.65, respectively.

Tile predicted LEX primary vortex breakdown lo-cations arc presented in figure 16 along with those ob-

tained from different flight (ref. 42) and wind tunnel

(refs. 12 and 15) experiments at subsonic conditionsfor various angles of attack. Tile experimental inves-

tigation of reference 15 was conducted in a low-speedtunnel with a 7- by 9-ft test section on a 1/9-scale

model of the F/A-18 aircraft at Re _ 1 × 106. The

wind tunnel data for the longitudinal location of theLEX vortex breakdown are presented over a range

of a = 21.5 ° to 29 ° for the configuration with and

without the empennage (i.e., vertical and horizontal

tails). The data for the baseline configuration (i.e.,

with empennage) indicate that the vortex break-down location moves upstream with increasing angle

of attack and that the overall characteristics gener-

ally correlate well with the data gathered from other

sources. However, the data (rcf. 15) presented in

figure 16 clearly indicate that the LEX vortex break-down moves further aft without the empennage par-

ticularly for the lower range of angle of attack (i.e.,

21.5 ° < a _< 24°). As expected, the empennage

and, in particular, the vertical tails, which arc lo-cated downstream in the path of the LEX vortical

flow, induce a pressure-fieht disturbance that prop-

agates upstream and precipitates vortex breakdown.The vortex breakdown location is predicted farther

aft than those obtained experimentally at c_ _ 26 °

and 30 °. Although no data are presented in refer-cnce 15 for the LEX vortex breakdown location at

c_ = 19 °, which corresponds to the angle of attack

of interest in the present computation, the trend of

the data reported for the tailless (i.e., without em-

pennage) configuration indicates a strong possibility

of a coherent vortex system over the entire lengthof the configuration. The absence of a LEX vortex

breakdown at c_ = 19 ° is consistent with the present

computational prediction as discussed in the previous

paragraph.

Surface pressures. The static surface pressure co-

efficients computed for the F/A-18 aircraft configu-

ration at the selected flow conditions (table I) with

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

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

attack. (See the color bar.) At high angles of attack,the suction peaks in two small regions of the LEX

apex and over the blended flap exceed the lower con-

toured limit of -3.0; the pressure coefficients in these

two regions are represented by solid white. Limita-tion of the pressure coefficient contours to a narrower

range would allow more color variation, which wouht

accentuate the pressure gradients. The following sta-

tions at which both flight and wind tunnel data havebeen measured are highlighted in white: FS 85, 107,

142, 184, 253, 296, and 357.

The effects of angle of attack on the com-puted surface pressure coefficients as presented in

figures 17(a) 17(c) appear to be most pronouncedin two regions. These regions over the LEX and

the wing upper surface are directly influenced by

the neighboring off-surface LEX vortex system andstalled flow over the wing, respectively. With increas-

ing angle of attack, the LEX vortical flow" appears to

accumulate more strength as evidenced by the highersuction-peak footprint. At high angles of attack (i.e.,

_> 25.8°), the increase in the LEX vortex strengthnot only affects the aerodynamic loads on the LEX

surface, but it also has significant influence on the

adjacent surfaces. For example, figure 17(c) clearly

illustrates regions of low pressures acting on the fuse-lage aft of the canopy; these pressure levels are com-

parable in magnitude to those computed on the LEX

upper surface.

At a = 19 °, figure 17(a) indicates that a major

area of the wing upper surface, aft of the wing-flap

hinge line, has pressure coefficient levels of about

-0.7 _< Cp _< 0.4. As discussed earlier in conjunc-tion with the tuft patterns (figs. 14(a) 14(c) for flight

tests and figs. 15(a) 15(c) for numerical simulation),this portion of the wing exhibited chaotic flow char-

acteristies attributed to stalled flow. With increasing

angle of attack, the surface pressure coefficients com-

puted on the wing upper surface show an extended

region of lower Cp levels (i.e., Cp <_ -0.7) becauseof the localized flow expansion. Also evident was a

suction-peak footprint associated with a leading-edge

vortex flow, which developed over the blended flap

region at o_ = 19 ° and intensified significantly with

increasing angle of attack. At c_ = 19°, the surfacepressure coefficients indicate a small suction-peak

footprint associated with the wingtip vortical flow,

which does not appear in the solutions at higher an-

gles of attack.

9

The surfacepressurecoefficientscomputedfortheF/A-18 aircraftconfigurationat all threeanglesof attackarepresentedin figures18(a)and 18(b).Tile pressurecoefficientsat the forcbodystationsare shownin figure 18(a) and are plotted as afunctionof 0. (See fig. 3.) In general, the forebody

surface pressure distribution shows an increasingsuction-peak level with increasing angle of attack.

The computed surface pressures suggest an incipi-

ent flow separation at 0 _ 150 ° (starboard) and 210 °

(port side) between FS 142 and FS 184 for c_ = 25.8 °and FS 107 and FS 142 for a = 30.a °. These flow sep-

arations would subsequently form the leeward fore-

body vortices with clearly defined suction-peak foot-

prints (i.e., 0 ,,_ 155 ° and 205 °) at FS 184 for both

c_ = 25.8 ° and ao.3 °. Because the forebody geom-

etry is composed of a smooth curved surface withno discontinuities (or limited to within the numeri-

cal discretization error), the triggering mechanism for

the resulting flow separation is an adverse pressure

gradient within tile boundary layer.

The computed LEX surface pressure coefficients

are plotted in figure 18(b) as a function of LEX-

exposed semispan y/s. The LEX pressure distribu-tions are presented for the same range as in the pre-

vious color contour figures 17(a) 17(c). In general,

the LEX upper surface pressure distribution can becharacterized by a large suction-peak footprint, at

y/s _ -t-0.50 associated with the primary vortex sys-

tem. At y/s ,_ -t-0.80 just outboard of this largesuction-peak footprint at FS 296 and FS 357, areas

of smaller suction-peak footprints exist that corre-

spond to the LEX secondary vortex system. Note

that the sharp spikes in the LEX upper surface pres-

sure distribution just inboard of the leading edge

(i.e., y/s _ ±1) result from numerical artifacts andhave occurred previously in numerous computational

studies of vortical flow separations from sharp-edged

configurations. (See refs. 17, 20 24, and 43 44.) Aswith tile forebody, the LEX pressure distributions

also indicate higher suction peaks with increasing

angle of attack, except at FS 357, where the lack

of increase in the LEX primary suction peak for> 25.8 ° can be attributed to the influence of vortex

breakdown. IIowever, this effect does not appear toimpact the secondary vortex suction peak as evident

from its consistent increase with increasing angle of

attack. Note that the computed surface pressures at

c_ = 30.3 ° clearly indicate a small low-pressure re-

gion over the upper surface LEX-fuselage juncture

at FS 357 (y/s _ 4-0.1). This low-pressure region,which has also been seen both in wind tunnel and

flight data, is chiefly attributed to the entrainment of

the forebody vortices at the LEX-fuselage juncture.

10

(See fig. 15(c).) The LEX lower surface pressure dis-

tribution shows increased compression at the higherangles of attack.

The correlations of the computed surface pressure

coefficients with the flight data for the forebody and

the LEX are presented in figures 19, 20, and 21 for

= 19 °, 25.8 °, and 30.3 °, respectively. Note that the

flight data shown in figures 19(a) and 19(b) are ob-tained at slightly different flow conditions and with

some geometrical differences between the numerical

model and the flight vehicle. Experimental aero-

dynamic effects from the latter geometrical differ-ences have been found to bc small and arc discussed

in detail in the following section.

Computational and flight pressure coefficientdata for the entire forebody length are in excellent

agreement throughout the examined range of flow

conditions. The computational results not only pre-

dict the overall pressures as well as the trends butaccurately simulate tile pressure distributions that

correspond to small flow features such as the leeward

forebody vortices. The pressure data disagreements

at FS 142 (0 _ 90 ° and 270 °) are caused primarily byan antenna fairing on the HARV that was not mod-

eled numerically. This antenna fairing can clearly be

seen in the flight photograph of the ttARV (fig. 14(a))

just ahead of FS 142 (highlighted in white). Notethat for c_ = 30.3 °, the suction peak associated with

tile primary vortex flow at FS 142 (0 _ 158 ° and

202 °) is slightly underpredicted.

The computed upper and lower surface pressure

coefficients for the LEX are generally in good agree-ment with the flight data at all three angles of attack.

However, a more detail assessment of the pressure

correlations reveals some differences and the possible

causes. In general, the correlations tend to degrade inthe outboard region, which is essentially dominated

by the LEX secondary vortex flow. The LEX pri-

mary vortex suction peak is predicted to be slightly

outboard at FS 253 and the magnitude is under-

estimated at FS 296 throughout the examined rangeof c_. However, at the last LEX station FS 357, the

magnitude of the primary vortex suction peak is pre-

dicted very well at the higher angles of attack of 25.8 °and 30.3 ° but is underestimated at 19 ° .

Finally, the complete computational results are

correlated with the corresponding flight data for the

forebody and the LEX in figures 22(a) and 22(b).The results clearly demonstrate the accuracy with

which tile theoretical solutions predict the sensitiv-

ity of the surface pressures to changes in angle of

attack. The incremental changes and trends of thecomputed surface pressure distributions as a function

":! |li

of angle of attack appear to agree well with the cor-responding flight data. In particular, note the fairly

good prediction for the LEX primary vortex suctionpeak at FS 357, which reveals the upstream influ-

ence of the blockage precipitated by tile vortex break-down. A favorable correlation is also presented for

the LEX lower surface pressure distributions, which

clearly demonstrates pressur9 sensitivity to the angle

of attack. As expected, the computed lower surface

pressure distributions at FS 357 exhibit excess com-

pression caused by fairing and closing off the inletface and are discussed in the next section.

The present computational results are encourag-

ing for simulating the overall flow features and the

pressure distributions for the forebo(ty and LEX con-

figuration at various flight conditions. Nonetheless,a wind tunnel experiment was initiated to evaluate

the aerodynamic effects of various configurational dif-

ferences between the flight vehicle and the numer-

ical model, such as inlet flow simulation, empen-

nage, and the deflected flap geometry. As mentionedearlier, these simplifications of the numerical model

were incorporated because of the limitations imposed

by either the flow solver and/or the grid generation

complexity.

Wind Tunnel Data

As discussed earlier, the experimental data pre-

sented here were obtained with a 0.06-scale F/A-18aircraft model which was tested in the Langley 7-

by 10-Foot High-Spee<t Tunnel. The primary objec-tive of the test was to validate the present compu-

tational results by providing experimental data on a

configuration that was more representative of tile nu-merical model. The experimental data analysis was

conducted to isolate tile acrodynamic effects of the

empennage (vertical and horizontal tails), inlet, and

various flap deflection settings with and without tile

empennage by evaluating the surface pressure coeffi-cients measured on the forebody and the LEX.

Effect of empennage. The forebody and LEXpressure coefficients measured on the CFD wind tun-

nel model, with and without empennage, are pre-

sented in figures 23(a) and 23(b) for 5f = 0 ° atthree angles of attack. In general, tile removal of

tim empennage has minimal effect on measured pres-sure coefficients of tile forebody as well as the LEX;

at a = 19 ° tile difference is almost indistinguishable.

Also, note that at FS 357, the removal of the tails

causes a very small increase in the suction level ata = 25.8 °.

These tail effects on the forebody and LEX pres-

sures arc presented in figures 24 and 25 for the same

range of flow conditions at 5y = 25 ° and 34 °, respec-tively. Generally, the forebody pressures remain in-sensitive to the empennage presence regardless of the

flap deflection angle. However, the LEX pressures

begin to be influenced by the presence of the tails,

particularly in the aft LEX region. In general, the

experimental data indicate that the mlgmentationsof the vertical-horizontal tails result in tile following:

1. Insignificant cffcct on the forcbody pressures

throughout the examined ranges of 5f and (_

2. Negligible effect oil the LEX pressures measured

at FS 253 and FS 296 throughout the examined

ranges of 5.[ and ct

3. A slight decrease in the LEX vortex suction peak

at FS 357; the effect is greater with increasing 5f'

particularly for c_ _> 25.8 °

Effect of inlet fairing. The effect of the in-

let fairing on tile measured forebody and LEX sur-

face pressure coefficients is presented in figures 26(a)

and 26(b), respectively, for the flow conditions ofinterest. These aerodynamic data were obtained on

the configuration that included the empennage and

5f = 0°. TILe results presented in the figure clearlydemonstrate that fairing over the inlet face has very

little effect on the measured forebody surface pres-

sure coefficients throughout the examined range of a _.

However, the fairing over the inlet appears to slightly

decrease (i.e., more negative) the measured pressurecoefficients associated with the LEX primary and the

secondary vortex suction peak, particularly at the ad-

jacent aft stations. Note that tile latter effects seem

to diminish at higher angles of attack. As expected,

the fairing over the inlet causes only slight flow com-

pression under the LEX as reflected in tile lower sur-face pressure coefficient measurements at FS 357.

Effect of flap deflection. The effect of the wing

leading-edge flap deflection on tile forebody and LEXpressure coefficients measured on tile CFD wind tun-

nel model is presented ill figures 27, 28, and 29 for

a _ 19 °, 26 ° , and 30 ° , respectively. The results

clearly indicate that tile flap deflection angle has neg-

ligible influence on tile forebody and the LEX pres-sure coefficients throughout the examined range of a

except at FS 357 for a _ 30 °. (See fig. 29(b).) At thisLEX station, tile data reveal only a slight increase in

both tile primary as well as tile secondary vortex suc-

tion peaks (i.e., more negative) with increasing flapdeflection. Because of these small effects, the ap-

proximation (figs. 20 and 21) made in computing the

flow over the configuration with @ = 25 ° (instead of

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

One of the objectives of this study was to ascer-

tain whether the tails of tile F/A-18 aircraft CFD

11

wind tunnelmodelwouldalter tile previousconch>sionwith regardto theeffectofflapdeflectionontheforebodyandLEX pressures. Figures 3O 32 show

the experimental data obtained from tile F/A-18 air-craft CFD wind tunnel model with the empennage.

The aerodynamic effect of wing leading-edge flap de-flection on tile forebody and LEX pressure measure-

ments appears to be insignificant over the examined

range of c_ except at FS 357. The pressure distribu-tion at the last LEX station shows a small sensitiv-

ity to the flap deflection at all three angles of attack.

Unlike the results (fig. 29(b) for FS 357) discussed inthe previous paragraph, tile slight increase in both

tile primary and the secondary vortex suction peaks

at FS 357 for o_ _ 30 ° is no longer achieved with theempennage installed. Actually, at this angle of at-

tack, the pressure distribution over the last LEX sta-

tion (fig. 31(b)) indicates a slight drop in the primary

and secondary vortex suction peak when increasing

5f from 25 ° to 34 °. Tile reduction in the LEX vor-

tical flow suction peak at a _ 30 ° with 5f = 34 °can be attributed to the nearby LEX primary vor-

tex breakdown at x/l _ 0.45 (fig. 16) precipitated

by the empennage. In general, the effect of wingleading-edge flap deflection on the forebody and LEX

surface pressure distribution over the F/A-18 aircraftCFD wind tunnel model with and without the em-

pennage is small.

Flight Versus Wind Tunnel Data

As discussed in tile previous section, the wind

tunnel data are primarily used to determine the aero-

dynamic effects of the various geometrical differencesbetween the numerical model and the HARV. Be-

cause the ultimate objective is to validate tile com-

putational results with the flight data, the accuracy

with which the wind tunnel data simulated the flightReynolds number flow characteristics is important.

As mentioned earlier, the forebody of the scale model

had grit strips that were positioned longitudinally

across the windward plane of symmetry at azimuthalangles 0 = 45 ° and 315 ° to trip tile expected laminar

boundary layer to a turbulent flow and thus simulate

tile assumed flight flow characteristics.

The forebody and LEX surface pressure coeffi-cients measured oil the CFD wind tunnel model with

empennage are presented in figures 33, 34, and 35,for a _ 19 °, 26 °, and 30 °, respectively. As discussed

earlier (figs. 26(a) and 26(5)), the aerodynamic ef-

fects on tile forebody and LEX pressures from the

fairing of the inlet face when compared with the

flow-through inlet were experimentally very small,confined only to the last LEX station, and diminished

quickly at higher angles of attack. The wind tunnel

12

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

positive) than the flight data with the correlation for

the aft stations FS 142 and FS 184 degrading with

increasing angles of attack (c_ = 25.8 ° and 30.3°).The degradation in the surface pressure correlations

is also apparent in the suction-peak regions of lee-

ward forebody vortices at 0 _ 158 ° and 202 ° . As

compared earlier (figs. 19 21), the pressure disagree-

ments at FS 142 (0 _ 90 ° and 270 °) are primarilycaused by an antenna fairing on the HARV that was

not incorporated on the CFD wind tunnel model.

Figures 33 35 clearly indicate excellent correla-

tion between the wind tunnel and flight data for allLEX stations in terms of both magnitude and general

trends throughout the examined range of c_. Thesefavorable correlations are attributed to the inviseid

flow characteristic of the LEX primary vortex sep-

aration line (i.e., fixed at the sharp leading edges),which leads to the development of the leeward vorti-

cal flows. As a result, the pressure distribution of the

LEX primary vortex flow indicates only a small sen-sitivity to the difference between the flight and wind

tunnel Reynolds numbers. Tile excellent correlation

also extends to the LEX outboard region where theflow separation that leads to the formation of tile sec-

ondary vortex structure is generally considered to be

a boundary-layer phenomenon. Note that the slight

pressure data disagreement on the lower surface ofthe last LEX station is primarily from additional

compression caused by the fairing of the inlet face

of the experimental wind tunnel model.

CFD Versus Wind Tunnel Data

The computational solutions were all obtained at

flight Reynolds numbers which were generally about

an order of magnitude higher than those achieved

experimentally in the wind tunnel investigation. As

a result, a computational study was performed toexamine the effect of Reynolds number on the solu-

tions. After assessing the Reynolds number effect,

the measured surface pressure coefficients are corre-

lated with the computational results for the forebodyand the LEX configuration. In addition, the aero-

dynamic effects of the wing leading-edge flap deflec-

tions are evaluated experimentally as well as compu-

tationally. Finally, the measured forces and momentsare correlated with the computational results.

Reynolds number effect. The primary objective ofthis section is to determine if the results of a typical

existing solution computed at flight flow conditions

are sensitive to an order-of-magnitude reduction in

Reynolds number. A new solution with a fully tur-bulent flow assumption was computed with the same

I | 1 =

flow-field grid by continuing thc solutions from the

results that had been obtained earlier at flight flow

conditions. As expected, the new converged solutionsat the wind tunnel Reynolds numbers indicate that

the flow-field grid provides finer resolution of the

boundary layer than that obtained earlier at flight

Reynolds numbers. This finer grid resolution of theturbulcnt boundary layer flow is naturally extended

onto the laminar sublayer region where it was rc-

solved with y+ .._ 1 instead of y+ _ 3 for tile car-

lier computations at flight Reynolds number. Onthe basis of the prior solutions (rcf. 17) obtained on

thc isolated F/A-18 aircraft forebody-LEX configu-ration, this order of grid refinement is not expected

to have any significant effect oil the prescnt compu-tational results.

The effect of Reynolds number on the computed

forcbody and tile LEX surface pressure coefficients is

shown in figures 36(a) and 36(b). The results clearlyindicate that thc computed surface pressure coeffi-

cients arc insensitive to the change in Reynolds num-ber at c_ = 19° and Mzo = 0.34. At this flow con-

dition, the Reynolds number effect on the computed

forces and pitching moment was also small and is dis-

cussed later in conjunction with the measured windtunnel data. To assess thc sensitivity of the com-

putational results to changes in Reynolds number at

the higher angles of attack, a similar computational

study was performed at c_ = 30.3 ° and Moc = 0.24.At this flow condition, the results also indicated that

the surface pressure cocfficients, forces, and pitching

moment were generally insensitive to the change in

Reynolds number. Note that the small sensitivityof the surface pressure distribution to the change in

Reynolds number for a comparable range and magni-

tude has also been reported in reference 45 for a tan-

gent ogivc configuration at c_ = 30 ° and M_c = 0.2.

These findings justify the surface pressure correla-tions between the present wind tunnel data and the

computational rcsults that have been obtained at

flight Reynolds number flow conditions.

Surface pressures. Tile computed forebody and

the LEX surface pressure coefficients are compared

with the experimental data obtained on the CFD

wind tunnel model (figs. 37 39) for the range of c_ ofinterest. Tile computed results are the same as those

correlated earlicr with the flight test data. Also,

note tile consistency of tile configuration geometrical

representation used for both sets of data such as tileflap deflection angle, empennage, and inlet.

The pressure coefficients for the forebody indicatethat the wind tmmel data measurements are slightly

higher (i.e., more positive) than the computationalpredictions, which wcrc shown earlier to be in excel-

lent agreement with the flight data throughout tile

examined range of (_. (Sec figs. 19 21.) In addition,measured surface prcssure distributions at FS 184 do

not indicate the expected suction-pcak footprints as-

sociated with the presence of the vortical flows at

higher angles of attack (i.e., c_ _> 25.8°).

The primary and secondary vortex suction-peak

footprints in the measured LEX upper surface pres-sure distributions indicate that the expected overall

flow physics of the LEX configuration has been ex-perimentally simulated. The computed pressure co-

efficients on the LEX upper surface appear to be in

reasonable agreement with the corresponding windtunnel data for the cxamincd range of c_. As cx-

petted, the subject correlations rcvcal some exist-

ing differenccs that are generally very similar both

in trcnd and magnitude to those discussed earlier in

conjunction with tile computational-flight data com-

parison. (Sec figs. 19 21.) Tile computed lower sur-face pressure distribution at FS 357 is clearly in good

agreement with the measured wind tunnel data at all

three angles of attack. This favorable correlation onthe LEX lower surface can be attributed mainly to

thc similarity of the geometry representation for the

inlet fairing in both the numerical and the CFD windtunnel models.

Finally, the complete computational results arccorrelated with the corresponding wind tunnel data

over the forebody and tile LEX in figures 40(a) and

40(b), respectively. The results clearly show the sen-

sitivity of tile forcbody and the LEX surface pres-sure distribution to the changes in angle of attack

for both the computed and the measured wind tun-

nel data. Similar to the earlier comparisons between

the computed and the flight test results (figs. 22(a)

and 22(b)), the present wind tunnel data correlate

reasonably welt with the computational results intrends and incremental changes of the surface pres-

sures as a flmction of angle of attack except for thevortex flow simulation at the aft forebody stations

for the range of higher c_. As discussed earlier, thediscrepancies of surface pressures for the forel)ody,

which had a smooth surface geometry with no dis-continuities, are attributed chiefly to the lack of scale

simulation of the high Reynolds number flow, partic-

ularly in the separated flow regions where the viscous

effects dominate tile ensuing flow characteristics.

Effect of flap deflection. The primary objec-

tive in this section is to investigate the capabil-

ity of the present computational method to predict

the aerodynamic effect resulting from different wingleading-edge flap deflections. The computational re-

sults as well as the experimentally measured surface

pressure coefficients for the forebody and the LEX

13

configurationarepresentedfor both_ = 0° and25°in figures41(a)and 41(b). Theexperimentaldatawereobtainedat flowconditionsthat wereverycloseto thoseof the computationswith the exceptionoftile Reynoldsnumber.However,at theseflowcon-ditionsdiscussedearlier(figs.36(a)and36(b)),thecomputedsurfacepressurecoefficientsfor the fore-bodyandtile LEX configurationwereinsensitivetotheReynoldsnumber.

Theaerodynamiceffectof wingleading-edgeflapdeflectionon tile computedsurfacepressurecoeffi-cientsoil theforebodyappearsto beverysmallandnearlyconstant.Similarly,this insensitivityof theforebodypressuresto the flapdeflectionis alsoev-ident from the experimentalwindtunneldatapre-sentedill figure41(a). However,the aerodynamicinfluenceof flapdeflectionon thecomputedsurfacepressurecoefficientsfor the LEX configurationap-pearsto be slightly morepronounced,particularlyat the aft stationswherethey becomephysicallycloserto theflapconfiguration.In general,tile com-putedresultsshowthat theflapdeflectioncausesanincreasein the LEX suction-peaklevelin a regionwhichessentiallyliesbelowtheprimaryvortexflow.However,theexperimentalwindtunneldataindicateonly a minimalchangein the LEX-measurcdprcs-suredistributionresultingfrom the flap deflection.Theeffectsof flapdeflectionon theLEX lowersur-facepressuresalsoappearto bc small,as indicatedt)5'both thecomputationalresultsandthemeasureddata.

Forces and moments. In this section, the com-

puted longitudinal aerodynamic characteristics are

correlated with those experimentally measurcd onthe CFD wind tunnel model. Because all the com-

putational results were obtained initially at flight

Reynolds numbers, which were gcncrally about an

order of magnitude greater than those achieved inthe experiment, tile effect of Reynolds number on

the computed forces and moments for a typical casc

is cvaluated. Note that a similar analysis on tile

surface pressure distributions, discussed earlier, in-

dicated that an order-of-magnitudc reduction in theReynolds number had negligible effects on the com-

puted surface pressures on the forcbody and tile LEX

configuration.

The longitudinal aerodynamic characteristics,computed at both the flight and the wind tunnel

Reynolds numbers, are presented in figure 42 for

a = 19 °. Figure 42 also includes the correspond-

ing experimental data point obtained for the wind

tunnel model that matched the geometry of the nu-merical representation. To be consistent with the

subsequent data analysis, plotting scales are selected

for a range that bounds the overall available longi-

tudinal aerodynamic characteristics. Among others,

two specific conclusions can be drawn from the re-

sults with respect to the effects of Reynolds numberon the computed forces and moments and the correla-

tion betwccn predicted and measured data. Similar

to the earlier findings in conjunction with tim sur-

face pressures, the cffect of Reynolds number on the

computed total forces and moments also appears to

be very small. However, note the slight increase inthe drag coefficient, which is attributed directly to

the reduction of the Reynolds number to match that

achieved in the wind tunnel expcrimcnt. Further-

more, the computed results presented in figure 42agree reasonably well with the measured wind tun-

nel data except for the total lift coefficient, which

appears to have been slightly underpredicted. Thclift underprediction at a = 19 ° is not surprising be-

cause as shown earlier with regard to the surface

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

results also underpredicted the measured LEX pri-mary vortex suction peak in the aft stations. As a

result, the LEX vortex lift contribution to the total

lift has probably been compromised.

Experimental and computational longitudinalaerodynamic characteristics for the entire range of

flow conditions arc presented in figure 43. Thc ex-perimental data arc presented for the CFD wind

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

configuration without geometrical alteration, which

providcd the necessary datum for the force and

moment data analysis. Similarly, figure 43 showscorresponding computational results that have been

obtained with the numerical model. Although thelatter two sets of data are consistent with one an-

other as a function of Mach numbcr (i.e., Mo _ 0.34,0.25, and 0.24 for c_ _ 19° , 25.8 °, and 30.3 ° ,

respectively), they differ slightly from the constant

M_ = 0.30 at which the data for the baseline F/A-18

aircraft configuration were obtained. However, pre-

vious experimental data (ref. 12) obtained from the

same 0.06-scalc F/A-18 aircraft wind tunncl model

clearly indicate that small variations (i.e., +0.05)in Maeh number, particularly in the low subsonic

range, do not have significant influence on overall

longitudinal aerodynamic characteristics. The lattercffeet as well as the carlicr finding of the influence

of Reynolds number on the computed longitudinal

aerodynamic characteristics validates the data cor-

relations presented in figure 43 despite the inconsis-

tencies in Mach and Reynolds numbers. The dataanalysis of the forces and moments is presented in

the following three categories: the experimental aero-

dynamic characteristics for thc baseline configuration,

14

I II_

the experimental aerodynamic characteristics for the

CFD wind tunnel model, and the correlations be-

tween the computational results and the correspond-

ing wind tunnel data.

The experimentally measured lift, drag, andpitching moment coefficients for the baseline F/A-18

aircraft configuration indicate essentially stable aero-

dynamic characteristics throughout the examined

range of o_. However, some degradation in aero-

dynamic characteristics is apparent, particularly inthe reduced rate of increase in lift coefficient with

increased angle of attack beyond _18 °, which can

be attributed to tile LEX vortex breakdown (fig. 16)

and/or the stalled flow over the wing. (See figs. 14(a)

an(] 15(a).) The pitching moment characteristics forthe baseline configuration remain fairly stable (i.e.,

dCm/da < 0) even at the range of higher c_ despite

tile loss of lift caused by the LEX vortex breakdown

ami stalled flow over the wing.

The experimentally determined effects of configu-

ration geometrical simplifications on the longitudinalaerodynamic characteristics are evident ill figure 43

by tile difference between the data presented with

open and solid circular symbols. Tile fairing of the

inlet face and the removal of the empennage cause a

slight decrease in lift and an increase in drag coeffi-cients only at the higher angles of attack of 25.8 ° and

30.3 °. These geometrical simplifications also result in

a reversal of tile pitching moment characteristics (i.e.,

dCm/dc_ > 0). Note that the latter change in thepitching moment characteristics can be attributed di-

rectly to the absence of the horizontal tail.

The computed longitudinal aerodynamic charac-

teristics (i.e., solid square symbol) for the geomet-rically simplified F/A-18 aircraft configuration com-

pare favorably with the corresponding wind tunnel

data (i.e., solid circular symbol). Tile computed lift,drag, and pitching moment coefficients correlate rea-

sonably well with experimental data throughout the

examined range of ct.

Concluding Remarks

Flow analyses of results from a variety of flight

tests, wind tunnel experiments, and thin-layerNavier-Stokes flow simulations are presented for

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 ofattack. In general, the computational results cor-

rectly predict the major flow characteristics such as

the forebody flow structures, the wing leading-edge

extension (LEX) vortex system, and the subsequent

vortex breakdown with increasing angle of attack,

forebody and LEX vortex interactions, and deflectedflap leading-edge flow separation leading to a stalled

flow over the wing upper surface.

A wind tunnel experiment was conducted with

a 0.06-scale F/A-18 aircraft model to ascertain theaerodynamic effects of the geometrical differences be-

tween the flight aircraft and the numerical model.

The wind tunnel data revealed isolated aerodynamic

effects of the empennage, fairing of the inlet face, and

wing leading-edge flap deflection angles. In general,

analyses indicated that the geometrical differenceshave only minimal influence on the surface pressure

distributions on the forebody throughout the exam-

ined angle of attack range. However, the LEX sur-face pressures are affected slightly by the geometrical

changes, particularly in the aft region and with in-

creasing angle of attack.

The experimental wind tunnel data are also com-pared with the flight test results to determine the

capability of the ground-based facility to simulate

the flight Reynolds number flow characteristics. This

study revealed that the LEX surface pressure coeffi-cients measured on the wind tunnel scale model cor-

relate very well with the flight test data. However,

data analysis of the surface pressures on the fore-

body indicates some disagreement between the flightand wind tunnel data, particularly in the aft stations

where the flow is separating at tile higher angles ofattack.

The wind tunnel data are presented for the longi-

tudinal aerodynamic characteristics measured on the

baseline F/A-18 aircraft configuration as well as theCFD wind tunnel model. These data reveal that the

fairing of the inlet face and the removal of the empen-nage cause a slight decrease in lift and an increase in

drag coefficients only at higher angles of attack. As

expected, the experimental wind tunnel data also in-dicate that these geometrical simplifications resulted

in a pitch-up moment characteristic. Furthermore,

the computed longitudinal aerodynamic character-

istics correlate reasonably welt with the experimen-tal measurements obtained on the CFD wind tunnel

model.

NASA Langley Research CenterHampton, VA 23681-0001

August 30, 199.1

15

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10. Fisher, David F.; Del Frate, John tI.; and Zuniga, Fanny

A.: Summarg of In-Flight Flow _Zisualization Obtained

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14. Hebbar, Sheshagiri K.; Platzer, Max F.; and Cavazos,

Odilon V.: A Water Tunncl Investigation of the Effects

of Pitch Rate and Yaw on LEX Generated Vortices of an

F/A-18 Fighter Model. AIAA-91-0280, Jail 1991.

15. Martin, C. A.; and Thompson, D. H.: Scale Model

Measurements of Fin Buffet Due to Vortex Bursting on

F/A-18. Maneuvering Aerodynamics, AGARD-CP-497,

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16. Meyn, Larry A.; Lanser, Wendy R.; and James, Kevin

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18. Thomas, James L.; _,Valters, Robert W.; Reu, Taekyu;

Ghaffari, Farh_ut; Weston, Robert P.; and Luckring,

James M.: A Patched-Grid Algorithm for Complex Con-

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89-0121, Jan. 1989.

19. Benek, J. A.; Steger, J. L.; Dougherty, F. C.; and

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89-0339, Jan. 1989.

21. Cummings, Russell M.; Rizk, Yehia M.; Schiff, Lewis B.;

and Chaderjian, Neal M.: Navier-Stokes Predictions of the

Flowfield Aroumt the F-18 (HARV) Wing and Fuselage at

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22. Rizk, Yehia M.; Sehiff, Lewis B.; and Gee, Ken: Numer-

ical Sinmlation of the Viscous Flow Around a Simplified

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2631-AF, Feb. 1983. (Available from NTIS as PB 83-

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and Bates, Brent L.: Transonic Navier-Stokes Solutions

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vol. 28, no. 6, June 1991, pp. 381 388.

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'_Vll i

28. Luckring, James M.; Ghaffari, Farhad; and Bates, Brent

L.: Status of Navier-Stokes Computations About tile

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Technology, Volume 1, Joseph R. Chambers, William P.

Gilbert, and Luat T. Nguyen, eds., NASA CP-31-19,

Part 2, 1992, pp. 703 722.

29. Thomas, J. L.; Taylor, S. L.; and Anderson, \V. K.:

Navier-Stokes Computations of Vortical Flows Over Low

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30. Vatsa, V. N.; Thomas, J. L.; and tVedan, B. W.: Navier-

Stokes Computations of Prolate Spheroids at Angle of At-

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as AIAA-87-2627.)

31. Biedron, R. T.; and Thomas, J. L.: A Gcncralized

Patched-Grid Algorithm With Application to the F-18

Forebody With Actuated Control Strake. Comput. Sgs.

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

32. Baldwin, Barrett; and Lomax, Harvard: Thin-Layer Ap-

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

personic Viscous Flows Around Pointed Bodies at Large

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34. Roc, P. L.: Characteristic-Based Schemes for the Eu-

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ume 18, Milton van Dyke, J. V. Wehausen, and John L.

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

35. Van Leer, Brain: Upwind-Difference Methods for Aero-

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La_yc-Scale Computations in Fluid Mechanics, Bjorn E.

Engquist, Stanley Osher, and Richard C. J. Somerville,

eds., American Math. Soe., 1985, pp. 327 336.

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38.

39.

40.

41.

42.

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d,L

45.

Fox, Charles H., Jr.: Real Time Data Reduction Capa-

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17

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"0OFigure 2. Tile F/A-18 aircraft geometry. All linear dimensions are in feet.

18

18o° _ e ,,90 ° 270

0° _' ......180 °

_, ....

270 °

" 180 °

90°@ 270°

FS 1070o

.V

180 °

'8°o 90 ° 270 ° 90 ° 270 °

FS 85 FS 70 O°0o

-y -----.j_ y

FS 357

_y

FS 296

_'_Y '_- _ _YFS 253

Figure 3 Planform of F/A-18 aircraft with forebody and LEX-fuselage cross-sectional pressure measurementstations.

19

RepresentationCAD

..... CFD

Faired-oversplitterplate

Splitterplate

(a) FS 401.

tat ion

..... CFD

X slot

_"_/_ Faired-over

cavity

20

(b) FS 441.

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

(a) Obliquetop view.

(b) Obliquebottomview.

Figure5. TheF/A-18aircraftCFDsurfacegridrepresentation.

21

Figure6. TheF/A-18aircraftCFDsurfacegrid withundeflected(port) andblended(starboard)flaps.

Deflectedwingleading-edge Blendedflap

re_"

splitterplate

Figure7. Close-upof F/A-18aircraftCFDsurfacegridandhighlightedsurfacemodifications.

22

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

i!i [I...... --

I

H-(

: i

(a) Far field.

(b) Near field.

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

\

\

\

\

23

Figure9. Far-fieldsideviewof F/A-18 aircraftflow-fieldblockingstrategy.

24

! |1i

0 m

-i.0

-I.5

_aoo

.a-2.0

-2.5

-,5 _

-30 I I I I I I0 I 2 3x 103

Iterations

(a) Residual convergence.

,.--1r,.)

8

6

4

2

0--

-2 [0

I i I i II 2 3x 103

Iterations

(b) CL convergence.

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

25

Figure11.Sting-mounted0.06-scaleF/A-18aircraftwindtunnelmodel.L-92-01827

26

| _|Ii

leading-edge flap

(a) Oblique rear view.

Metal sheet toclose gap betw

deflected flaand fuselage

L-92-2805

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.

27

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

28

(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.

29

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.

31

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

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

32

Forebodyprimaryvortexcorestreamlines

(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.

33

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

45

4O

35

3O

25

20

15

10

%

O_A • A

4, ii • A

O

I I I I I I I I I I0 .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.

34

! 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.

35

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

Figure 17. Concluded.

36

Cp

Cp

Cp

Cp

-1.0

-.5

0

.5

1.0

0

-1.0 -

FS 85

45 90 135 180 225 270 315 360

0, dcg

Data ix, deg M_ R e xl0 _' _f, deg Tails Inlet

CFD 30.3 0.243 10.20 25 Off Faired

CFD 25.8 0.253 10.80 25 Off Faired

CFD 19.0 0.340 13.50 25 Off Faired

-.5

0

.5

1.0

0

FS 107

= • 1 , I , I , I , I , I , I , I

45 90 135 180 225 270 315 360

0, deg

-3

-2

Cp - I

0

I , I _ I , k J_ £, I t I _ _1___

-I.00 -.75 -.50 -.25

FS 253

0 .25 .50 .75 1.00

y/s

1.0 -

-.5

i

45 90 135 180 225 270 315 360

FS 142

.5 -

1.0 _____I , 1 t I _ 1 _t_ 1 I _ I _

0

0, deg

-3

-2

Cp -I

0

FS 296

t _I , I , I ,_1__

-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

-I.0 -

-.5

0

.5

1.0

FS 184

1 , I z I .._._1 , I _ I , I , I

45 90 135 180 225 270 315 360

Cp

-3 _ FS 357

0

lI , I , I , A_, I , 1 , I , I , 1

-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

0, deg y/s

(a) Forebody. (b) LEX.

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

37

Cp

Cp

Cp

Cp

1.0

-.5

0

.5

1.0

0

-I .0

-.5

0

.5

1.0

0

-I .0

-.5

0

.5

1.0

0

- 1.0 --

FS 85

Data or, dog M_ R e xI0 + 81 , deg Tail,, Inlet

CFD 19.0 0.340 13.50 25 Off Faired

Flight 19.1 0.300 I 1.50 25 On Open

l [ , I t t , I , I , I , I , I

45 90 135 t 80 225 270 315 360

0, dog

FS 107

, 1 , I , I , I 1 I , I , I J I

45 90 135 180 225 270 315 360

0, dog

3 --

-2

Cp -I

0

1

-1.00 -.75 -.50 -.25

FS 253

, I , I , I _ I i I , I , I , I

0 .25 .50 .75 1.00

y/s

FS 142

, I _ I _ I , I , I , I _ I L I

,_5 315 36045 90 135 180 "_'_ 270

0, dog

-3 _- FS 296

-2

-t.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

-.5

0

.5

1.0 I

0 45 90 135 180 225 270 315 360

FS 184

, I , I , I _ I i I I _ t , I

0, deg

Cp

-3 _- FS 357

/-2 F ooo oO o ,

0

I , I , I _I , I , I , I z I J I

- 1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

(a) Forebody. (b) LEX.

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

38

III

Cp

Cp

Cp

Cp

-1.0

-,5

0

.5

1.0

0

-I.0

-.5

0

.5

1.0

0

FS 85

Data I_, deg M_ R e xl0 6 15f, deg Tails Inlet

CFD 25.8 0.253 10.80 25 Off Faired

Flight 25.8 0.253 10.80 34 On Open

, I , I , I L I , I , I L L.L_!

45 90 135 180 225 270 315 360

0, deg

FS 107

, 1 , 1 , I __.___ I , I t I , I t_J

45 90 135 180 225 270 315 360

0, dcg

Cp

-3

-2

-1

FS 253

0

1

-I.00 -.75 -.50 -.25

I _ I .t_._l , I , I _ I t. I _ _

0 .25 .50 .75 1,00

y/s

-1.0

f FS 142-.5 []

.5

1,0

0 45 90 135 180 225 270 315 360

0, deg

-I.0 -

-.5

0

.5

1.0

0

FS 184

, I , I , I ,_..I _ I _ I i _.[_J.2

45 90 135 180 225 270 315 360

O, deg

Cp

-3 F FS 296

II

-2 / D D DD

0

___9_-_ 4::_ .-- _ _-_

I _a_.5_.___-t. , I , 1 L I , 1 _

-I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

-3

_- FS 357

Cp -I

0 _"_-..--...._D_ l_ _ _ D D i_..___._../_

I , t , I_ 1 _, L t A , I , I _ I

-I.00 -.75 -.50 -.25 0 .25 .50 ,75 1.00

y/s

(a) Forebody. (b) LEX.

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

39

-I.0

-.5

Cp 0

.5

1.0

-1.0

-.5

Cp 0

(.5

1.0

-1.0

-.5

Cp 0

.5

1.0

-1.0

-.5

Cp 0

.5

1.0

FS 85

Data _x. deg M_ R e xl0 6 5f, deg Tails Inlet

CFD 30.3 0.243 10.20 25 Off Faired

Flighl 30.3 0.243 10.20 34 On Open

____ • 1 _ A_ .L_A _ I , I , I45 90 135 180 225 270 315 360

0, deg

FS 107

45 90 135 180 225 270 315 360

O, deg

-3

-2

Cp -1

0 o1 I , 1 _ ] , I • I , I _ I J I L

-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

FS 142

O o

_h t I I , I _ L i I i I _ J i l

45 90 135 180 __5 270 315 360

0, deg

-3

-2

Cp -1

0

FS 296

¢_00 000

0 0

-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

FS 184

-2

0

1j____ , i , I , I , I , I 1 1 J ]

0 45 90 135 180 225 270 315 360

0, dog

-3 -FS 357

-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

(a) Forebody. (b) LEX.

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

40

:t II1

Cp

Cp

Cp

Cp

-1.0-

-.5

.5

1.00

FS 85

I _ 1.___ 1 • t _ 1 , I , I

45 90 135 180 225 270 315 360

0, deg

o

[]

O

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

CFD 30.3 0.243 10.20 25 Off Faired

CFD 25.8 0.253 10.80 25 Off Faired

CFD 19.0 0.340 13.50 25 Off Faired

Flight 30.3 0.243 10.20 34 On Open

Flight 25.8 0.253 10.80 34 On Open

Flight 19.I 0.300 I 1.50 25 On Open

-I .0

-.5_

o r

5L1.0 ___L I , I __ _k_t._. I , I

0 45 90 135 180 225 270 315 360

0, dcg

-3

-2

Cp -I

0

I

-1.00 -.75 -.50 -.25

/_o FS 253 o_

0 .25 .50 .75 1.00

y/s

-1.0

-.5

.5

1.0

FS 142

O (D

__L., J L .I , I , • .___1

0 45 90 135 180 225 270 315 360

0, dog

-3 -FS 296

O0 000

_._.,___ _ -._-._.o _." ,.,o_'__ ...,..--

0 .25 .50 .75

Cp -[

0

1

-I.00 -.75 -.50 -.25

y/s

.00

-1.0

-.5f FS 184

.5

1.0

0 45 90 135 180 225 270 315 360

0, deg

-3 -FS 357

o

L__ _ _____1_ ___

0 .25 .50 .75 1.00

Cp -1

0

I-I.00 -.75 -.50 -.25

y/s

(a) Forebody. (b) LEX.

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

41

-1.0 -FS 85

Cp

-.5

.5

o• • o

_ _',,! !:.,.-" i •

t l!

1.0 _ • J I , I , I • l 1 .L x .1 a.___J

0 45 90 135 180 225 270 315 360

0, dog

Dala or, deg M_

• Tunnel 30.5 0.251

• Tunnel 25.9 0.250

@ Tunnel 19.2 0.342

O Tunnel 30.6 0.253

D Tunnel 25.8 0.252

O Tunnel 19.1 0.343

Re xl0 b

1.21

1.21

1.61

1.11

1.12

1.46

131.,deg Tails Inlet

0 On Faired

0 On Faired

0 On Faired

0 Off Faired

0 Off Faired

0 Off Faired

-I.0

.5

1.0

-3--FS 107

o_:O . , itl_o -2

|.il*'"lillooiil. if..i-- -nn • @I_41 Cp -1

I t OlI0

__ , 1 _ 1 t t • t _ I _ I I

45

FS 253 •_ •

• • •

-.. = • ,."-n i i 4 •• • • llm •

@4 @ @ @ 4 4 4 14

90 135 180 225 270 315 360

0, dcg

-I.00 -.75 -,50 -.25 0 .25 .50 .75 1.00

y/s

°1 F,.. "FS 142 FS 296

• •llii

-.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

-,5

Cp 0

FS 184

-2

Cp -1

..;illli;llllilli;.,,_:.- tti0

1.0 _ _ • 1 , 1 L 1 , 1 _L__I__

0 45 90

[] FS 357iii

i | joile i tl

ill " in i" iLl":l..!i

i

_ t It t t i• • 1 _l • I _ I , I

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°.

42

:1 111

Cp

-1.0 -FS 85

OO O O

5 ,_,'"! ,.,,:o **** III ****

5 !t I:

1.0 _ ___1 J l _1 _ [ •_ I_1

0 45 90 135 180 225 270 315 360

0, deg

Cp 0

-1,0 --

FS 107

.5 o,_,_,. • -",o.._u m'- ui • g_U

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

Io Oll

Cp

Cp

Data or, deg M_

• Tunnel 30.5 0.250

• Tunnel 25.8 0.249

• Tunnel 19.1 0.342

0 Tunnel 30.5 0.251

El Tunnel 25.8 0.251

O Tunnel 19,1 0.341

-3

R(. xl0 '_ 8 t , deg Tails Inlet

1.15 25 On Faired

1.15 25 On Faired

1.54 25 On Faired

1.15 25 Off Faired

1.15 25 Off Faired

1.51 25 Off Faired

-2

Cp -I

.5 0

1.0 ____ 1 * [ _ 1 t I , I , 1 _ I _t/ l

0 45 90 135 180 225 270 315 360 -I.13

0, deg

• FS 253 •

• • []• i •

O0 [] [] i i O o

Ui ( • • • ) / i• • • •

@ • @ @

___1_, I l J • _ 1 _1 , 1

-.75 -.513 -.25 0 .25 .50 .75 1.00

y/s

-1,0 f -3 L

FS 142 FS 296 •• •

Oi i • • [] Oo •-.5 i Io_ _Leo -2 oil iioooo_/ Oli_ __dA_i i --m Lh_i i il/m TM

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

0 Cp -1 li_ i i • i• i

.5 0

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 --

-.5

FS 184 i FS 357

-2 @ i i

• II I,*ub_l_ go.a_ _ _ •

"'tl illllili '" .i ,.•. .. c,ot .1;;i i"

00

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°'

43

-I.0 -FS 85

Cp

-.5

.5

1.0

0

-.! ,,,..

! |:

__ I _ I K l J [ , 1 L 1 , I , I

45 90 135 180 225 270 315 360

0, deg

Data a, deg M_ Re×l 0 o 8j, deg Tails Inlet

• Tunnel 30,6 0.251 1.18 34 On Faired

• Tunnel 25.8 0.250 1.17 34 On Faired

• Tunnel 19.2 0.341 1.54 34 On Faired

O Tunnel 30.5 0.250 1,21 34 Off Faired

[] Tunnel 25.8 0.251 1.22 34 Off Faired

O Tunnel 19.1 0.340 1.59 34 Off Faired

-l,0 -

.5

1.0

FS 107-3

I|r_i||l __|i |_JJ

I._ -.wJe,ee+• ._%lCp -I

i o ot.0

45

• FS 253 •

• • [][]

:: .,:..[] I • t i•

I . ": +[]4) 4, • • • • • • O0

1 _ J_ ± I l I _ I J 1 _ J t 1 t _1

90 135 180 225 270 315 360 - 1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

0, deg y/s

- 1.0 _ FS 142

-5 [ _.+_o)| ,,.|

co OItol'** ***el| |

1.0 t 1 l __. _1 _ I , t _ _. __

0 45 90 135 180 225 270 315 360

0, deg

Cp -I

_ 3 [

-2 _._J i_ i_

_[]i-. • []1_o,e ,'e't •

°i-1.00 -,75 -,50 -.25

FS 296

_[][]_

[] | •

I ) I + I J 1 , I

0 .25 .50 .75 1.00

y/s

" -I.0

-.5

Cp 0

.5

1.0

0

FS 184

! o!

45 90 135 180 225 270 315 360

0, deg

Cp

-3

-2

-1

0

I-I.00 -.75 -,50 -.25

8W

t t t t l t et l L J _ ! , I J I + l ,___

0 ,25 ,50 .75 1.00

y/s

(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 °

44

1 I i

-1.0 -FS 85

Cp

-.5

0

.5

1.0

0

_oo Q_

" |O @

| •|

45 90 135 180 225 270 315 360

0, dcg

Data ft, deg M_

• Tunnel 30.6 0.250

• Tunnel 25.9 0.253

• Tunnel 19.1 0.341

O Tunnel 30.5 0.25 I

121 Tunnel 25,9 0.250

C' Tunnel 19.2 0.342

R_ xlO_ 81 , deg Tails Inlet

1.II 0 On Open

1.13 0 On Open

1.47 0 On Open

1.21 0 Oi1 Faired

1.21 0 On Faired

1.61 0 On Faired

-1.0 -

.5

1.0

FS 107

|•,,,••,,,no•el,•

i' °t!

45 90 135 180 225 270 315 360

Cp

-3-

-2 -O,_ i

-I -@• •

0

I

-1.00 -.75 -.50 -,25

FS 253 •

• J

• • mo°• • @ • il• • •

__1 t_ _ [ _ J , _± _l L_._

0 .25 .50 .75 1.00

0, dcg y/s

-1.0 --

-.5

Cp 0

.5

1.0

3LFS 142 FS 296- •

• • [] °o

- ••-. _,,._•Oo -2 w_ j i • _ Ji•o,_|" __mb .... o..m_tdi_nlin mn n I _iw'_ • ui + im_mm

• * Cp __l__ + * | I **_

I tl..._ "11|| o[ * * * *

0 45 90 135 180 225 270 315 360 -1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

0, dcg y/s

Cp

-1.0 -

-.5

.5

1.0

0

FS 184

! 0!

45 90 135 180 225 270 315 360

0,deg

-3 [- FS 357

k i []w

I i

*|_| m@co -' t i

I __ _ _a L _ L ,, J -,_

-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

(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 °.

45

Cp

-1.0

-.5

0

.5

1.0

0

FS 85

O

O00000000000Q

O0

I , l _ A_, J , I _ t _ I _ I

45 90 135 180 225 270 315 360

O, deg

Data 0_,deg M_

O Tunnel 19.1 0.343

[3 Tunnel 19.1 0.34I

<) Tunnel 19.1 0.340

R e ×10 '_ 81 , deg Tails Inlet

1.46 0 Off Faired

1.51 25 Off Faired

1.59 34 Off Faired

Cp

-1.0 -

-.5

0

10 0

.5 -

1.0

0

FS 107

OQoI_O0000000000 O01_O 0

O0 o

_, I _ I J t _ 1 _J 1 ._ t_ ..L._

45 90 135 180 225 270 315 360

0, deg

-3 -

-2

Cp -I -(30

0

FS 253

0 o 0 0

0 0 OO0 0 0 0

1 _ LJ , I , I , I , I , I _ I

-I .00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

Cp

-I.0 _ FS 142

-.5

u 0000

1.0 1 _-, 1 _ 1 __A__L J. _ .1_ x_J

0 45 90 135 180 225 270 315 360

0, deg

-3

-2

Cp -1

FS 296

_I_00 oo 0 0 00_0 0

0 0

I _.t _ [ _A _ , I , I

-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

Cp

-1.0

-.5

.5

1.0

FS 184

00_00O0

l t J 1 t J l 1 _ I , I [ I 1 J

45 90 135 180 225 270 315 360

0, deg

-3--

-2

FS 357

o_0 o

o @cp -I _ ;_;_0_ 0 0 °°0_

0 00 0 0 0 0 0

1 , I , I _ ! , I , I , I , 1 __3

- 1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

0ooOo °

(a) Forebody. (b) LEX.

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

46

! 81:

Cp

Cp

-I.0-

-.5

FS85

O0 0 oOOO0 Oooo 0

0 0 o.5

1.0 _, 1 , I , I _±__1 _ t •

0 45 90 135 180 225 270 315 360

0, deg

Data (x, deg M_

0 Tunnel 25.8 0..5.

121 Tunnel 25.8 0.251

0 Tunnel 25.8 0.25t

Re xl0 _ _1, deg Tails Inlet

1.12 0 Off Faired

I. 15 25 Off Faired

1.22 34 Off Faired

-I ,0 --

-.5

0

_o o

.5

1.0

0 45

FS 107

oOCg°°OOooooooO_ O_bo °

O0 o

[__l_ 1 , 1 , 1 L [ , I , I

90 135 180 225 270 315 360

0, dcg

-3

-2

Cp -1

FS 253

00

0 O 0 0

D O O0

0 0 0 0

I _, J . 1 [ • _t • , I , I

-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

Cp

-1.0 --

-.5

0

t

.5

1.0

FS 142

O00O_OOoo

00 O00

45 90 135 180 225 270 315 360

0, deg

-3 -

O O-2 -

Cp -1

0

FS 296

O

a o °°ooc_

0 0

1 _ 1 L 1 J t 1 1 _ L t_ [ 'J

-I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

Cp

-I .0

-.5

0

.5

1.0

0

FS 184

0 o

.___L ____._l , 1 , I J [ • d45 90 135 180 225 270 315 360

0, deg

-3 -

0 FS 357OQ

O

QOQO0

-2

6c_c_ ° {}Cp -I - 0

0 O

0 0 0 0 0 0

-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

(a) Forebody. (b) LEX.

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

47

Cp

-1.0

-.5

.5

1.0

0

FS 85

000 OOQ

0 00000 Q

-0 00

J 1 J 1 , [ , I _ 1 , I , I , J45 90 135 180 225 270 315 360

0, deg

Data a, deg M_

O Tunnel 30.6 0.253

E3 Tunnel 30.5 0.251

O Tunnel 30.5 0.250

Re x I0 " 5f, deg Tails. Inlet

1.11 0 Off Faired

1.15 25 Off Faired

t .21 34 Off Faired

Cp

-1.0

-.5

FSI07

o_% o o_ °_bo0 000000 0

0

0{ }0 000

.5

1.0 ___ _ 1 , I , I L I _, l _

0 45 90 135 180 225 270 315 360

0, deg

-3

-2

Cp -I

0 FS 253 0

O OO O

DO 0 0 O00

0

0

1 _ I , I , t , [ ____1 t_ , _

-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

Cp

-1.0

-.5

.5

1.0

FS 142-3

-2

0 0- C o -1

0 0

0 0 0

, [ , I , I , 1 L I , ] , I , I

45 90 135 180 225 270 315 360

8, deg

FS 296

O O O

0 0

1 , I , 1 , _1 J 1 , I _ I L _1

-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

Cp

-I.0

-.5

0

.5 '_

1.0

0

FS 184

O 0

, I _ 1 • t US_ , 1 , I , I , 1

45 90 135 180 225 270 315 360

O, dog

3f-2

Cp -10 I O

] , I ,

-1.00 -.75 -.50 -.25

FS 357

@_000_

0

0 0 0 0 0

0 .25 .50 .75 [ .00

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 °.

48

| !!11

Cp

Cp

Cp

Cp

-1.0

-.5

0

.5

1.0

0

FS 85

0000000000000

0 O0

45 90 135 180 225 270 315 360

0, deg

Data _,deg M_

0 Tunnel 19.2 0.341

Tunnel 19,I 0.342

O Tunnel 19.2 0.342

Re xl0 _

1.54

1.54

1.61

8f, deg Tails Inlet

0 On Faired

25 On Faired

34 On Faired

-I .0 -

-.5

0

.5

1.0

0

FS 107

ooOWOOOooooooOO °_bo o

}00 OO o

45 90 135 180 225 270 315 360

0, deg

-3 -

-2

Cp -1

FS 253

O OO

<30 0 0

0 0 0 0

0

I

-1.00 -.75 -.50 -.25

OO

1 l t_ _ _ a 1 _ L _ 1 I _ J0 .25 .50 .75 1,00

y/s

-1.0 F-

t FS 142_ _ 5 F

000_000 o

0 foOO Oo O

IlIFI I. _J ± _ i l t__L _ l_

0 "45 90 135 180 225 270 3t5 360

0, deg

-3

f FS 296-2

iot:t_o o o o oooc _o

0 0

0

-I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

Cp -1

-I .0

-.5

.5

1.0

0

FS 184

Oo_oO

}0 OO

_ I , 1 , i _ L c 1 L 1 _

45 90 135 180 225 270 315 360

O, deg

-3-

-2

Cp -1

0

1

- 1.0_

FS 357

0 88 o

0000_0

00 0 0 0 0 0

-.75 -.50 -,25 0 .25 ,50 .75 1.00

y/s

(a) Forebody. (b) LEX.

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

49

Cp

-1.0

-.5

.5_

1.0

FS 85

_00 O00 00000 O0

0 0 o

• [ , A_ 1__1 I , I , r

45 90 135 180 225 270 3 [ 5 360

0, dcg

Data or. deg M_

O Tunnel 25.8 0.250

[3 Tunnel 25.8 0.249

© Tunnel 25.9 0.250

R,2 ×10 6 8f, deg Tails lnlel

1.17 0 On Faired

1.15 25 On Faired

1.21 34 On Faired

Cp

-1.0

-.5

0

.5

1.0

0

FS 107

oOdm°OOooooooO_ °_bo °

0 000

, 1 _ I , 1 , 1 _ 1 , I , t _ ]

45 90 135 180 225 270 315 360

0, dcg

-3

-2O

OO

Cp -1

0

1

-I.00 -.75 -.50 -.25

FS 253

0O

0 0 0O0

0 0 C_ 0

J I , I , I , [ ___L_ _ I , I J I

0 .25 .50 .75 1.00

y/s

Cp

-I.0

-.5

0

I

.5

FS 142

0 000

1.0 , 1 , I , I , I _ !.=L._ t l

0 45 90 135 180 225 270 315 360

O. deg

-3 -

-2 O _'

Cp -I -

0

FS 296

o

o o oo_

0 0

1 ___ L_ I , I , 1 t I , I , 1

-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

yls

Cp

-1.0 -

-.5

0

I

.5

1.0

0

FS 184

}O O0

_1 , I , I , I , 1 _ 1 , I , I

45 90 135 180 225 270 315 360

0, dog

-3 -

-2

Cp -1

8FS 357

6 8 _

O066_)6

0

0 0 0 0 0 0

1 _J I , I , t , I , i _ I __

-I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

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 °.

50

:1 lli

Cp

Cp

Cp

Cp

-1.0

-.5

.5

1.0

0

FS 85

000 000

0 0000 ° 0

0

45 90 135 180 225 270 3t5 360

0, deg

Data ¢x, deg M_

o Tunnel 30.6 0.251

D Tunnel 30.5 0.250

O Tunnel 30.5 0.251

R_..× 10 _ 8f, deg Tails Inlet

1.18 0 On Faired

1.15 25 On Faired

1.21 34 On Faired

-I .0

-.5

0

"5_3 O 0

FS 107

o_P°_o o °_b°

0 0000000 0

O0 o

1.0 _ _.1 t 1 t _l t_ _ 1 _1

0 45 90 135 180 225 270 315 360

0, dog

-3

-2

Cp -1

0

0 FS 253 O

O O

O O

_o 0 0 O0o

0

1 , 1 , 1 , I , I 1 _L _ _J

-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

-1.0 -

-.5

0

.5

1.0

FS 142

0 _

O0

45 90 135 180 225 270 315 360

0, dog

-3 -

0 0

-2 -0_00

Cp -I

FS 296

0

0 o O000¢KP

0 0

0

I _ _l_. 1 , 1 1 1L 1 _ _1__± I _ • t _1

-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

-I.0

-.5

0

.5

1.0

0

FS 184

Oo o oO

Oo

45 90 135 180 225 270 315 360

0, deg

-3 -

-2

Cp -i _0

0

I

-1.00 -.75 -.50 -.25

FS 357

9_eeoeo ° °oo¢_

0

- O iO O O O O

0 .25 .50 .75 1.00

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 °.

51

Cp

-1.0

-.5

t

.5

1.0

0

FS 85

o+ _ .°++ o.oi+g,o*g,

_i_°Ooq__o o°

_L_I , 1 , I + I , I _ I , I___

45 90 135 180 225 270 315 360

0, dcg

Dala _, deg M_ R e xlO ++

Flight 19.1 0.300 I 1.50

Tunnel 19.1 0.342 1.54

_t , deg Tails Inlet

25 On Open

25 On Faired

Cp

Cp

-1.0 _- FS 107

-.5

0Ill.5

1.0 t .J l + I , l , I t I t L_I

0 45 90 135 180 225 270 315 360

0, deg

-! .0

-.5

.5

1.0

0

FS 142

O O

_ _ _ _ +"_i_m_+__ _ i_,o v i_i _

----L_--_- _--L--L_ I , l +--_

45 90 135 180 225 270 315 360

0, deg

-3

-2

Cp -I

_,°o°_o

FS 253

o_°%

00 0 0 0 0

I , I L I t I L t _ l , I , I , l

-1.00 -.75 +,50 -.25 0 .25 .50 .75 1,00

y/s

-3

FS 296

0 ° 0 o

ii_ _ i:° v i_°_ "_-qm_0

• 0 (J_©

0

OO 0 0 0 O0 0

l,l,l,l,=_l,l,l,l,l

-I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

Cp

-I.0

-.5

,5

1.0

0

FS 184

,-"_,_i_o• 0 o

+_ 1_ + I , I J+ I , ! , [ _ I

45 90 135 180 225 270 315 360

0, deg

-3

FS 357

Cp -I

-2 _oo

0 ©

I ___ I , I , I , I , I , I , I , I

- 1.00 -.75 -.50 -.25 fl 2_ .5Cl .75 ! .00

y/s

_o

e

%0+,+

o o0o %0 _ ,

(a) Forebody. (b) LEX.

Figure 33. Correlation of surface pressure coefficients from flight and CFD wind tunnel model tests with

empennage at c_ _ 19 °.

52

! ;!!_

Cp

Cp

Cp

-I.0 -

-.5

FS 85

oo • _oo_o,o,, _o

o oo o

,_,o %_

45 90 135 180 225 270 315 360

O, deg

Data ct, deg M_ RexlO 6

Flight 25,8 0.253 [0.80

Tunnel 25.8 0.250 I.I7

_1, deg Tails Inlet

34 On Open

34 On Faired

-1.0

-.5

.5

1.0

0

FS 107

O O

45 90 135 180 225 270 315 360

0, deg

-3

-2

Cp -1

0

1

FS 253

o_o °_oo

oo._o O _o •

-I.00 -.75 -.50 -.25

o o o o o o o o

___L_ _ _t 1 , I I I

0 .25 .50 .75 1.00

y/s

-1.0 -

-.5

FS 142

O

o

,5 -

1,0 A L_I _ I i F _ i ] i 1 k_

0 45 90 135 180 225 270 315 360

0, dcg

-2

Cp -I

0

I

o

-I.00 -.75 -.50 -.25

FS 296

©, t_ O

_o °o

o o o o o o o o

0 .25 .50 .75 1.00

y/s

Cp

-1.0 -

-.5

0

.5

1.0

FS 184

o_°__o

00

__ A___J _ L _ 1_ i I , I _ l __.1

45 90 135 180 225 270 315 360

O, deg

-3

-2

Cp -I

FS 357

_o eoo • _

o i_o o o o o o •• • •- •

I __[ _ _- _LI _ • t_ • _1__-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

(a) Forebody. (b) LEX.

Figure 34. Correlation of surface pressure coefficients from flight and CFD wind tunnel model tests with

empennage at a _ 26 °.

53

Cp

-1.0 -

-.5

0

.5

1.00

FS 85

0%.0 • 10 00 O

o, i_%o,looi_ "o

o o

o o

o_ °Ib 0

__L L .L_[_ J t ± 1 -L L_45 90 135 180 225 270 315 360

0, dog

Data o< deg M_ Rcx I 0 _

Flight 30.3 0.243 10.20

Tunnel 30.6 0.251 1.18

8f, deg Tails Inlet

34 On Open

34 On Faired

Cp

-I .0 -

-.5

.5

1.0

FS 107

o o

o • v% •

I , 1 L 1 K 1 _ I , I , I , I

45 90 135 180 225 270 315 360

0, dcg

-3

-2

Cp -I

o_O FS 253

o• _o

• 8 °

o •o o

{OlD

o _

0 0 0 0 0 00 0

1 __I t _ _ l_ _ 1 _ I , ! __ _ _J

-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

Cp

- 1.0

-,5

i

.5

1.0

0

FS 142

o OO c_

.

, I t I , I , I t I , I _ I J I

45 90 135 180 ,_._" 270 315 360

0, dog

Cp

o°il o-2 •

-1

O

FS 296

8 o @o

_ooo o_

0

0 0 0 0 0 0 0 0

I _ 1 _ 1 _ t _____ , 1 _ I

-I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

Cp

-1.0 -

-.5

.5'

1.0 _j I L I , 1 _ I , I t I , I , I

0 45 90 135 180 225 270 315 360

0, deg

3[_ FS 184 FS 357

_ '_o 0 _ o °

• O• -1(}_i' 0o o• o °

[ _ 1 n I_ ___1 _ J , I , 1..._

-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

(a) Forebody. (b) LEX.

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

54

II I1

Cp

Cp

Cp

Cp

-I.0 -

-.5

.5

1.0

0

FS 85

45 90 135 180 225 270 315 360

0, deg

Data oc, deg M_ R e ×10 _' ,fir, deg Tails Inlet

CFD 19.0 0.340 13.50 0 Off Faired

CFD 19.0 0.340 1.45 0 Off Faired

-1.0FS 107

-.5

.5

1.0 _ , l , [ t _ , I • 1 _l

0 45 90 135 180 225 270 315 360

0, deg

-3 _ FS 253

-2

0I t 1 t_

-I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

-I.0 --

-.5

.5

1.0

0

FS 142

45 90 135 180 225 270 315 360

0, dcg

-3 -

-2

Cp -1

0

1

-1.00 -.75 -.50 -.25

FS 296

J [ _ I J 1 , l L J _, ,_ l0 .25 .50 .75 .00

y/s

-1.0

-.5

0

.5

1.0 _ _1 _± A _ 1 __1 _ [ _ J

0 45 90 135 180 225 270 315 360

0, deg

-3 -

-2

Cp -I

0

I

-1.00 -.75 -.50 -.25

FS 357

0 .25 .50 .75 1.00

y/s

(a) Forebody. (b) LEX.

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

55

Cp

-1.o

-.5

I

.5

1.0

0

FS 85

45 90 135 180 225 270 315 360

0, deg

Data _, deg M_ R_ x l0 6 _1, deg Tails Inlet

CFD 19.0 0.340 13.50 25 Off Faired

Tunnel 19. I 0.341 1.51 25 Off Faired

Cp

-1.0

-.5

FS 107

.5

1.0

0

___A _ I a I ___L , l _ I__

45 90 135 180 225 270 315 360

0, deg

Cp

-3

I

-1.00 -.75 -.50 -.25

[-

FS 253

_2 P

_,o

0

0 .25 .50 .75 1.00

y/s

Cp

-1.0 F

FS 142

I _ 5 V

.5

1.0 J._._ _ 1 , I

0 45 90 135 180 225 270 315 360

0, dcg

-3

-2

Cp -I

1

-I.00 -.75 -.50 -.25

f FS 296

f

0 .25 .50 .75 .00

y/s

Cp

- 1.0 -FS 184

.5

1.0 __J , 1 J 1 _ L _ .l • l_J_._

0 45 90 135 18(1 225 270 315 360

0, deg

Cp

-3

-2

-1

0

I

-I.00 -.75 -.50 -.25

FS 357

ooOo

0 .25 .50 .75 1.00

y/s

(a) Forebody. (b) LEX.

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

56

"I • I

Cp

Cp

Cp

Cp

-1.0 -

-.5

.5

1.0

FS 85

45 90 135 180 225 270 315 360

0, deg

Data _, deg M_ Re × I 0 _' 5f, deg Tails Inlet

CFD 25.8 0.253 10.80 25 Off Faired

Tunnel 25.8 0.251 I. 15 25 Off Faired

-1.0

-.5

.5

1.0

0

FS 107

, _ L.__ I _L [ , • L [

45 90 135 180 225 270 315 360

0, deg

Cp

-3-FS 253

© ©

-1

I

-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

-I .0 -

-.5

FS 142

(

.5 --

1.0 ____ • 1_.1 • _ I _ I _ • , 1 ,1

0 45 90 135 180 225 270 315 360

0, deg

-3

-2

Cp -1

0

FS 296

-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

-1.0 -- -3

-.5

0

C

.5

FS 184

-2

Cp -I

0

1.0 ___ _ 1 __.1 t 1 , I , I I0 45 90 135 180 225 270 315 360 - 1.00 -.75 -.50 -.25

0, deg

FS 357O

©O

o° o

0 .25 .50 .75 1.00

yls

(a) Forebody. (b) LEX.

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

57

Cp

-I.0 -FS 85

-.50

.5

1.0 J 1 , 1 ± • , I , I , .1___1

0 45 90 135 180 225 270 315 360

0, deg

Data ct, deg M_ R e × I0 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

-1.0 -

-.5

0

.5 /

1.0

0

FS 107

I , I , 1. _ I , I , I L 1 _ I

45 90 135 180 225 270 315 360

0, deg

-3

/_ FS 253 /_

1

-1.00 -.75 -.50 -.25

1 , I , L_,_ 1 , i , 1 , t ,

0 .25 .50 .75 1.00

y/s

Cp

-1.0FS 142

-.50(

.5

1.0 _ I , I , I l I , I , I , I , I

0 45 90 135 180 225 270 315 360

0, deg

-3 -

o ©

-2 ___

Cp -I

FS 296©

1 _t L , 1 _ I _ I , I , I _ I

-I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

Cp

-I.0 -

-.5

.5

1.0

0

FS 184

I , / l l , I _ I , I , I , I

45 90 135 180 225 270 315 360

0, dog

-3

-2

Cp -1

FS 357

©

© ©o

0

o- o-I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

(a) Forebody. (b) LEX.

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

58

:I II

-1.0 -

-.5

Cp 0

.5

1.0

FS 85

45 90 135 180 225 270 315 360

O, deg

0

[]

0

Data _, deg M_ R e ×10 a tSf, deg Tails Inlet

CFD 30.3 0,243 10.20 25 Off Faired

CFD 25.8 0.253 10.80 25 Off Faired

CFD 19.0 0.340 13.50 25 Off Faired

Tunnel 30,5 0.251 1.15 25 Off Faired

Tunnel 25,8 0.251 1.15 25 Off Faired

Tunnel 19.1 0.341 1.51 25 Off Faired

Cp

-1.0FS 107

-.50_

{.5

1.0 __ 1_ , • t [ k_ _ 1

0 45 90 135 180 225 270 315 360

0, deg

-3

-2

Cp -I

0

k _ FS 253

o 5o ,0oy/s

-1.o -FS 142

-.5

.5

1.0 __L__L__._ 1 __ ___ •_ • • J

0 45 90 135 180 225 270 315 360

0, deg

Cp 0 Cp -1

-3-FS 296 [

0 0 0 n

0 " "_ --- -0- _

1 __ _1 _ • _ [ _ I

-I.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

-1.0 _- FS 184

-.5

,5

1.0 _ A_ l _ 1_ [ _0 45 90 135 180 225 270 315 360

0, deg

Cp 0 Cp -I

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 °.

59

-1.0FS 85

Cp

-.5

0

i

.5

1.0 l • L____ 1___ A_ , 1 , 1

45 90 135 180 225 270 315 360

0, deg

Dala oL deg M_

CFD 19.0 0.340

CFD 19.0 0.340

Tunnel 19.1 0.343

Tunnel 19.I 0.341

Re x I[) " 81 , deg Tails lnlel

13.50 0 Off Faired

13.50 25 Off Faired

1.46 0 Off Faired

1.51 25 Off Faired

Cp

-1.0

-.5

1.0

FS 107

L I _k i 1 L_L.t_ 1 , 1 , 1_

45 90 135 180 225 270 315 360

0, deg

Cp

-3

-2

0

I

-I.00 -.75 -.50 -.25

FS 253

0 .25 .50 .75 1.00

y/s

Cp

-I .0

-.5

.5

1.0

FS 142

, _ I , 1 _ 1 , _, 1 , I E I

45 90 135 180 225 270 315 360

0, dcg

-3 _- FS 296

-2

Cp -1-.-4:3.__jd-J-

0

1 t 1__ I _ 1_ 1 _1 b

- 1.00 -.75 .50 -.25 0 .25 .50 .75 1.00

y/s

Cp

-I .0

-.5

.5

1.0

FS 184

_x I _ J _ t_ l t _ 1 _ I

45 90 135 180 225 270 315 360

0, dcg

-3 _ FS 357

-2 t- [] ¢]n n

-1.00 -.75 -.50 -.25 0 .25 .50 .75 1.00

y/s

(a) Forebody. (b) LEX.

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

60

:_! I I

Cm

.4

.3

.2

.1

0

-.1 , I

Data M_ R e ×10 6 5f, deg Tails Inlet

O Tunnel 0.34 1.46 0 Off Faired

• CFD 0.34 1.45 0 Off Faired

• CFD 0.34 13.50 0 Off Faircd

I I _ 1 _ I , I

CL

2.0 --

1.5

1.0

.5

0

-,5 I

-5

I .1_ I _ I t I _ I _ I J_ I _ t , I 1 1

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 correlationwith measurements on CFD wind tunnel model at a _ 19°.

61

C m

.4-

.3

.2

.!

0

-.!

o o

__ I

M O =

O O

0.34

0|

0

0.25 0.24

l'

0OOOO

°°(_Oo

1 i 1 i..: I i 1 1 J

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

o Tunnel 0.30 --1 25 On Open

• Tunnel Mo =1 25 Off Faired

• CFD Mo =10 25 Off Faired

CL

2.0 B

1.5

1.0

.5

0O

Ol

O

O

oO(_ °°

0110o10

0

0

0

o¢ii°o| °

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.

62

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REPORT DOCUMENTATION PAGE OM8 No. 0704-0188

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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

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

NASA Langley Research Center

Hampton, VA 23681-0001

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

"Washington, DC 20546-0001

B. PERFORMING ORGANIZATION

REPORT NUMBER

L-17336

10. SPONSORING/MONITORINGAGENCY REPORT NUMBER

NASA TP-3478

11. SUPPLEMENTARY NOTES

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified Unlimited

Subject Category 02

Availability: NASA CASI (301) 621-0390

12b. DISTRIBUTION CODE

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-layerNavier-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 approximationsto 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 differentfree-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 structurescompared 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 approximationsmade to the numerical model had very little effect on overall aerodynamic characteristics.

14. SUBJECT TERMS

F/A-18 aircraft; Computational fluid dynamics; Navier-Stokes analysis; Flight data;

Wind tunnel data; High angles of attack; Vortical flows; Vortex breakdown

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

NSN 7540-01-280-5500

18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION

OF THIS PAGE OF ABSTRACT

Unclassified Unclassified

15. NUMBER OF PAGES

65

16. PRICE CODE

A04

20. LIMITATION

OF ABSTRACT

;tandard Form 298(Rev. 2-89)Prescribed by ANSI Std Z39-18298-102

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