NASA Technical Memorandum 4783 

The F/A-18 High-Angle-ofAttack Ground-to-Flight 

Correlation: Lessons Learned 

Daniel W. Banks, David F. Fisher, Robert M. Hall, 

Gary E. Erickson, Daniel G. Murri, Sue B. Grafton, 

and William G. Sewall 

January 1997 

-- II BOEING TECHNICAL LIBRARY 

I 

ST. LOUIS 

MAILCODE': S111-1025 

111111111 11111111 I 

LM203732E 

/ 
NASA Technical Memorandum 4783 

The F/A-18 High-Angle-ofAttack Ground-to-Flight 

Correlation: Lessons Learned 

Daniel W. Banks and David F. Fisher 

Dryden Flight Research Center 

Edwards, California 

Robert M. Hall, Gary E. Erickson, Daniel G. Murri, 

Sue B. Grafton, and William G. Sewall 

NASA Langley Research Center 

Hampton, Virginia 

National Aeronautics and 

Space Administration 

Office of Management 

Scientific and Technical 

Information Program 

1997 

TECHNIC L L ARY 
THE F/A-18 HIGH-ANGLE-OF-ATTACK GROUND-TO-FLIGHT 

CORRELATION: LESSONS LEARNED 

Daniel W. Banks and David F. Fisher 

NASA Dryden Flight Research Center 

Edwards, California 

Robert M. Hall, Gary E. Erickson, Daniel G. Murri, 

Sue B. Grafton, and William G. Sewall 

NASA Langley Research Center 

Hampton, Virginia 

ABSTRACT 

Detailed wind tunnel and flight investigations were performed on the FI A-I8 configuration to 

explore the causes of many high-angle-of-attack phenomena and resulting disparities between 

wind tunnel and flight results at these conditions. Obtaining accurate predictions of full-scale flight 

aerodynamics from wind-tunnel tests is important and becomes a challenge at high-angle-of-attack 

conditions where large areas of vortical flow interact. The F/A-18 airplane was one of the first 

high-performance aircraft to have an unrestricted angle-of-attack envelope, and as such the configuration displayed many unanticipated characteristics. 

Results indicate that fixing forebody crossflow transition on models can result in a more 

accurate match of flow fields, and thus a more accurate prediction of aerodynamic characteristics 

of flight at high angles of attack. The wind tunnel results show that small geometry differences, 

specifically nosebooms and aft-end distortion, can have a pronounced effect at high angles of attack and must be modeled in sub-scale tests in order to obtain accurate correlations with flight. 

NOMENCLATURE 

All force, moment, and pressure data have been reduced to coefficient form. All longitudinal 

data are presented in the stability axis system, while all lateral-directional data are presented in the 

body axis system. Moment data are referred to a moment center located at 25% of the mean aerodynamic chord. 

AR aspect ratio 

b span, ft 

b' local span, ft 
CL, max 

c 

FS 

HATP 

HARV 

LEX 

LE 

p 

S 

TE 

y 

a 

2 

l 

drag coefficient, drag/ q ooS 

lift coefficient, lift/q ooS 

maximum lift coefficient 

rolling-moment coefficient, rolling moment/qooSb 

dC 

lateral stability parameter or dihedral effect, d~l per degree 

pitching-moment coefficient, pitching moment/qooSc 

yawing-moment coefficient, yawing moment/qoo Sb 

dC 

directional stability parameter, d~n per degree 

pressure coefficient, ( p - p (0)/ q 00 

side-force coefficient, side force/ q 00 5 

mean aerodynamic chord, ft 

full-scale fuselage station, in. 

High-Angle-of-Attack Technology Program 

High-Alpha Research Vehicle 

leading-edge extension, or wing-body strake 

leading edge 

free-stream Mach number 

surface static pressure, Ib/ft2 

free-stream static pressure, Ib/ft2 

free-stream dynamic pressure, Ib/ft2 

Reynolds number based on mean aerodynamic chord 

Reynolds number based on forebody base diameter 

reference area, ft2 

trailing edge 

local span distance, ft 

angle of attack, deg 

- -- - -------
sideslip angle, deg 

flap deflection angle, deg 

circumferential angle, deg 

INTRODUCTION 

The design of high-performance fighter aircraft has always placed emphasis on maneuverability and agility. To this end the aerodynamic envelope of these aircraft has been expanded to include 

higher usable angles of attack. Significant separated and vortical flows develop over the aircraft at 

high angles of attack and may significantly interact. These types of flow fields result in nonlinear 

aerodynamics and cause significant changes in stability and control characteristics when compared 

with low angles of attack. The changes in stability and control at high angles of attack are difficult 

to predict from ground tests because of the varying sensitivities to test conditions such as Reynolds 

number and Mach number. It has been a goal of the NASA High-Angle-of-Attack Technology 

Program (HA TP) 1 to develop a better understanding of these flow fields and their effect on groundto-flight correlation. 

Obtaining accurate wind-tunnel-to-flight correlation is a significant challenge. High-angleof-attack conditions add a level of complexity to this already difficult task. In the end, a good 

understanding of the behavior of the associated flow fields and their potential interactions is 

necessary to produce useful design methodology and ground-test techniques that can accurately 

predict flight. 

This paper reviews the results of the various wind tunnel and flight investigations of the FI A-18 

configuration leading to a more comprehensive understanding of these disparities and more accurate ground-to-flight correlation at high-angle-of-attack conditions. The results reported herein 

were obtained from various wind tunnel tests of several 6%- and I6%-scale F/A-18 models, and 

flight tests of the F-18 High-Alpha Research Vehicle (HARV) conducted at the Dryden Flight 

Research Center, Edwards, California (DFRC). 

BACKGROUND 

The FI A-18 configuration, due to its interesting high-angle-of-attack characteristics, unrestricted angle-of-attack envelope, and the availability of a flight test aircraft and wind tunnel models, was adopted for much of the work of the HA TP. Since the F/A-18 aircraft was one of the first 

aircraft capable of controlled flight at high angles of attack, it displayed many unusual characteristics that were not well understood during early high-angle-of-attack flight tests. The features of 

the FI A-18 aircraft that made it attractive for high-angle-of-attack research included its contoured 

wing-body leading-edge extensions (LEXs), and its forebody shape, which is approximately elliptical with the major axis in the vertical plane aft of the radome. The design of the FI A-18 aircraft 

3 
used the mixed flow philosophy to obtain good high-angle-of-attack characteristics without significant impact on the conventionallow-angle-of-attack flight regime. At low to moderate angles of 

attack, the relatively low sweep wing works well. But at higher angles of attack, the leading-edge 

extensions (LEXs) generate strong vortices which, in addition to generating lift on the LEXs, 

propagate over the wings, and energize the wing flow fields, thereby extending the usable flight 

envelope beyond maximum lift. This same flow field is also responsible for many of the less 

desirable high-angle-of-attack characteristics of the F/A-18 aircraft. These strong vortices also 

impinge on the vertical tails and under some conditions cause severe buffeting.2- 6 The forebody 

of the FI A-18 aircraft generates a vortex pair at high angles of attack, which interacts with the 

strong LEX vortices, and under some conditions can cause wing rock7,8 and nonlinear effects on 

lateral stability. 

The FI A-I8 development program epitomizes the challenge of ground-to-flight correlation at 

high-angle-of-attack conditions. Early developmental wind-tunnel tests of the FI A-18 aircraft produced inconsistent results. The most notable of these inconsistencies was between the lateral 

stability of large-scale models tested at relatively low Reynolds numbers and small-scale models 

tested at higher Reynolds numbers. The small-scale higher Reynolds number data predicted the 

configuration would remain laterally stable at high angles of attack, while the large-scale lower 

Reynolds number data indicated the configuration would become laterally unstable at angles of 

attack near maximum lift.9 Designers intuitively relied on the higher Reynolds number data as 

being the most representative of flight. However, early flight tests showed the lower Reynolds 

number data better predicted the flight characteristics at high angles of attack. 

Subsequent tests of F/A-18 models and the HARV, as well as other non-F/A-18 tests have 

increased our understanding of high-angle-of-attack flow fields. The flow fields at these high 

angle-of-attack conditions typically consist of separated wakes and potentially interacting vortices. 

The FI A-18 high-angle-of-attack flow field is characterized by interacting forebody and LEX vortices. Since the forebody flow is more sensitive to Reynolds number effects and the LEX flow is 

more sensitive to Mach number effects, it is difficult to predict the high-angle-of-attack flight flow 

field and its interactions from sub-scale wind tunnel tests. The effectiveness of gritting to simulate 

higher Reynolds number boundary-layer cross flow transition at high angles of attack has been 

recently investigated, and has shown promising results. 10, 11 Fixing crossflow transition on the 

FI A-I8 forebody has been shown to reduce differences between wind tunnel and flight aerodynamic results. 12,13 The effects of small geometric differences, such as airdata probes and deformations 

for wind-tunnel supports, can have a pronounced effect at high angles of attack and are responsible 

for some of the differences seen between FI A-18 aircraft and model tests. 

MODELS, FACILITIES, AND INSTRUMENTATION 

The data presented in this paper were obtained from multiple wind-tunnel entries with several 

F/A-18 6%- and 16%-scale models and flight tests that were conducted with the HARV. The 

ground and flight tests were coordinated as well as possible to obtain meaningful comparisons 

between ground and flight data. This coordination included a common set of measurements, and 

flight conditions (a, p, and Moo) closely matched to the wind tunnel conditions. Also, for model 

4 
data, all leading- and trailing-edge flap positions were made consistent with the flight vehicle 

configuration at that flight condition. The configuration that was tested on the HARV at ex. ~ 26° 

and Moo ~ 0.76 was bLE = 33° and bTE = 0°. 

6% Models 

There were two 6% model configurations that are reported herein. The first is referred to as the 

MDlNavy 6% model (fig. 1). This model is the 6% wind-tunnel model originally used by 

McDonnell Douglas (St. Louis, Missouri) and the U.S. Navy during the design and test of the 

F/A-18 , which was tested in the current F/A-18 configuration. The second 6% model consisted of 

the back end of the MDlNavy 6% model with a more recently constructed and instrumented forebody section (from nose apex back to LEX/wing juncture). This forebody had pressure ports from 

the nose back to the mid LEX region and is referred to as the NASA 6% model (fig. 2). Thus, these 

two models/configurations shared the same wings, empennage, and support system hardware. The 

pressure instrumentation consisted of 440 pressure orifices distributed circumferentially at 5 stations on the nose and spanwise at 3 stations across the LEXs, as shown in figure 3. The pressures 

were measured by internally mounted electronic scanning pressure modules located within the 

forebody. The on-surface flow visualizations were obtained with a mixture of titanium dioxide 

(Ti02), mineral oil, and a small amount of oleic acid added for dispersion.

14 The mixture was 

applied to the model, the tunnel was then brought on condition long enough for the mixture to set, 

and then rapidly brought to a static condition when photographic documentation was taken. The 

off-surface flow visualizations were obtained with water vapor (injected or natural) and a laser 

light sheet in the wind tunnel, this was recorded with video cameras. IS The 6% model wind-tunnel 

Figure 1. MDlNavy 6% F/A-18 model in the NASA Langley 7- x 10-ft High Speed Tunnel. 

5 
tests were conducted in NASA Langley Research Center's 14- x 22-ft Subsonic Tunnel, 7- x lO-ft 

High Speed Tunnel, Low-Turbulence Pressure Tunnel, and in the David Taylor Research Center's 

7- x IO-ft Transonic Wind Tunnel. 

Figure 2. NASA 6% F/A-18 model in the NASA Langley 14- x 22-ft Subsonic Tunnel. 

-1.0 ~ 0 o 

~ 1.0 

ylb' ylb' 

Section A-A Section C-C 

960791 

Figure 3. Pressure measurement locations on F/A-18 forebody and LEX. 

6 
16% Models 

There are two 16% low Reynolds number models reported herein. The first was modified from 

an older preproduction 16%-scale F/A- 18 model configuration and although it was modified to 

more closely represent the FI A-18 configuration as it evolved, still incorporated some differences. 

Figure 4 shows this model and is called the NASA-1 16% F/A-18 model. There are limited flow 

visualization results presented here from this model. The second 16%-scale F/A-18 model represents the current F/A-18C configuration and was of much higher fidelity than the NASA-1 16%-

scale model. This second F/A-18 model is referred to as the NASA-2 16% model and shown in 

figure 5. The NASA-2 16% model was also designed with more instrumentation than the NASA-1 

16% model and all of the 6% models. Both 16% models were constructed of lightweight fiberglass 

since they were also used as free-flight models and therefore were limited to lower dynamic pressure than the 6% models. The NASA-2 16% model also included removable components; some 

that were instrumented to determine the flow field effects of having pressure port roughness 

effects, and others to assess changes in geometry (e.g. the noseboom). The results from the 

NASA-2 16% model will be used for all quantitative data (force, moment, and pressure distributions) since it had a higher degree of fidelity in both geometry and data, and because of its flexibility in replaceable components. The NASA-2 16% model had a more extensive number of surface 

pressure measurements than the instrumented NASA 6% model; however, only those which correspond to the fuselage stations of the NASA 6% model will be reported here (fig. 3). Both of the 

16% wind tunnel models were tested in NASA Langley Research Center's 14- x 22-ft Subsonic 

Tunnel and 30- x 60-ft Tunnel. The on-surface flow visualization was obtained from tests of the 

NASA-l 16% model. The technique was essentially the same as described for the 6% model, with 

the exception of the mixture being thinner to accommodate the lower dynamic pressures. 

Figure 4. NASA-l 16% F/A-18 model in the NASA Langley 14- x 22-ft Subsonic Tunnel. 

7 
Figure 5. NASA-2 16% F/A-18 model in the NASA Langley 30- x 60-ft Tunnel. 

HARV Flight Tests 

Dedicated flight tests were conducted with the F-18 HARV at the NASA Dryden Flight 

Research Center, Edwards, California (fig. 6). Although the vehicle was heavily modified for various phases of the program,16,17 the data shown in this report apply to the basic F/A-18 configuration. Some data were obtained with and some without the LEX fences that are part of the current 

F/A-18 configuration. The F/A-18 HARV has both leading- and trailing-edge flaps that are scheduled with angle of attack and Mach number. At ex 2:: 26° and M ~ 0.76 the leading-edge flaps are 

deflected to their maximum of 33° and the trailing-edge flaps are set to 0°. The HARV was flown 

in the F/A-18 fighter-escort configuration, without stores, and the wingtip missile rails were modified to carry special camera pods and wingtip airdata probes. Also the HARV was tested predominantly without LEX fences installed. The HARV was instrumented with surface pressure port 18 

at the same fuselage stations as shown in figure 3 for the 6% model. On-surface flow visualization 

was obtained in flight with a mixture of an evaporating fluid and dye. 19 The fluid and dye mixture 

was pumped out of the ports, which were later used for pressure measurements. The dye patterns 

set quickly as the fluid evaporated and were preserved until postflight photos could be obtained. 

Off-surface flow visualization was obtained in flight with injected smoke at the apex of the LEX's 

and nose.20,21 

Except where noted, all data presented was obtained from configurations without nosebooms 

or LEX fences. Also, except where noted, all wind tunnel data was obtained from models with only 

the nosering grit pattern used for low-angle-of-attack forebody transition fixing, but not the twin 

strips grit pattern used for high-angle-of-attack forebody crossflow transition. 

8 
EC 88 0095-002 

Figure 6. NASA HARV in flight at the NASA Dryden Flight Research Center, Edwards, California. 

RESULTS AND DISCUSSION 

Basic Aerodynamics 

The basic longitudinal aerodynamic characteristics of the F/A-18 in the high-angle-of-attack 

configurations from wind tunnel tests are shown in figure 7. The longitudinal characteristics 

are similar between the models, with the major differences being small changes in C L, max and 

variations in pitching moment. Maximum lift occurs at approximately a = 40°, with C L max -

1.80. The NASA 6% model has greater pitch stability than the NASA-2 16% model at angles of 

attack from 25° to 40°, and greater nose down pitching moment from a = 10° to a = 45°. Some 

effects of model geometry on this pitching moment difference will be discussed later in this report. 

The lateral-directional characteristics of the NASA 6% and NASA-2 16% models in the highangle-of-attack configuration at a = 40°, are shown in figure 8. An angle of attack of 40° is of significant interest since this is the angle of attack for maximum lift. The NASA 6% model shows 

higher levels of lateral stability than the N ASA-2 16% model at this angle of attack (-4° ::; ~ ::; 4°). 

The effects of Reynolds number, grit, and noseboom on lateral stability will be discussed later in 

the report. 

9 

-----.-

~I 
.3 Model Facility Moo Re c 

0 NASA-216% 30- x 60- ft 0.08 1 x 106 

.2 0 NASA 6% 7- x 10- ft HST 0.30 1.4 x 106 

.1 

Cm 

0 

-.1 

-.2 

2.5 

2.0 

1.5 

CL 1.0 

.5 

0 

-.5 

-10 0 10 20 30 40 50 60 0 .2 .4 .6 .8 1.0 1.2 1.4 1.6 

0:, deg Co 960794 

Figure 7. Longitudinal aerodynamics ofF/A-I8 models. ~ = 0°, C jLE = 33°, C fTE = 0°, CHT = _12°. 

The on-surface flow characteristics of the NASA-1 16% model at a = 36° and MDlNavy 6% 

models at a = 35° (both at Rec - 1 million) are shown in the oil flow visualization of figures 9 and 

10 respectively. While the two models were tested at nearly the same Reynolds number and angle 

of attack the forebody flow field visualizations show some significant differences. The most notable features on the NASA-I 16% model are a void region, seen in black, extending approximately 

from just behind the nose apex aft to the LEX juncture. This feature is most likely a laminar separation bubble, which would indicate that the flow upstream of it is laminar. On the front half of the 

nose the void is terminated by streamlines in an opposing direction. These lines indicate that the 

flow has separated aft of the bubble in this region (most likely laminar separation). On the rear half 

of the nose the void region is terminated by an almond-shaped region where the streamlines are in 

the same general direction as prior to the bubble. This evidence indicates a turbulent reattachment 

in this region. The streamlines in this almond-shaped region show a difference in the texture, 

appearance of higher surface shear can be noticed, a further indication that the flow in this small 

region is turbulent. On the leeward side of the model additional voided regions are noticed which 

appear to be caused by the impingement of the secondary separation vortex. These voided regions 

propagate over the canopy. A similar feature is seen near the leading edge of the LEX. Overall, the 

nature of the flow field of the NASA-l 16% model forebody is transitional, mostly laminar with 

small areas of turbulent flow. 

10 
Model Facility Moo Re c 

.025 

o NASA-2 16% 30-- x 60- ft 0.08 1 x 106 

o NASA6% 7-x10--ftHST 0.30 1.4x106 

.020 .04 

.015 .03 

.010 .02 

.005 .01 

CI 0 Cn 0 

-.005 -.01 

-.010 -.02 

-.015 -.03 

-.020 -.04 

-.025 -.05 L-_-L..._-..l'--_-'-_---'-__ ...I..-_--' 

-12 -8 -4 0 4 8 12 -12 -8 -4 0 4 8 12 

deg deg 960795 

Figure 8. Lateral-directional aerodynamic of F/A-18 models. a = 40°, b jLE = 33°, b fiE = 0°, 

bHT = _12°. 

Figure 9. Forebody oil flow visualization of NASA-2 16% F/A-18 model. a = 36°, Moo = 0.08, 

Rec - 1 x 106. 

11 
Secondary 

separati on 

960798 

Figure 10. Forebody oil flow visualization of NASA 6% F/A-18 model: two views a = 36°, Moo = 

0.22, Rec - 1 X 106. 

The surface flow of the MD/Navy 6% model, shown in figure 10, indicates some distinct differences from that of the NASA-I 16% model, even though they were tested at nearly the same 

Reynolds number and angle of attack. This model does not show any of the laminar separation bubbles that were prominent on the NASA-116% model. The MD/Navy 6% model does show a pooling near the 90° and 270° positions close to 2/3 the length of the nose, which would indicate 

laminar separation. Aft of this the flow appears to remain attached. There are no voided regions on 

the leeward side. There is a demarcation which is most likely caused by secondary separation, but 

is significantly weaker than that seen with the NASA-1 16% model and terminates at the LEX 

apex. Overall, the MD/Navy 6% model also appears to be transitional but with different features 

than that of the NASA-1 16% model. Some of these differences can be attributed to differences in 

12 
dynamic pressure and to the viscosity effects of the oil on changing the local geometry. The 

MDlNavy 6% model was tested at a higher dynamic pressure and thus the oil mixture was somewhat more viscous, so the smaller model would also be more prone to local variations in geometry 

from the oil. The extent of transitional flow on the MDlNavy 6% forebody was more difficult to 

determine than the NASA-1 16% because of effects of the oil on the MDlNavy 6% model and the 

resultant patterns. However, it appears that the flow field present (extent of transitional flow) on 

the forebody of each model was different even though this visualization was obtained at nearly the 

same Reynolds number. 

On-surface flow characteristics of the HARV at a = 34°, and Moo = 0.3 are shown in figure 11. 

The HARV displays a similar surface topology to that of the models with some notable differences. 

The forebody flow of the HARV appears to be fully turbulent aft of the radome. There is a small 

run of laminar flow on the radome, but it has fully transitioned by FS 107.8 Recall that the NASA-l 

16% model showed a significant vortex structure pass over the canopy and the MDlNavy 6%, 

under some conditions, showed a similar although less defined vortex which did not propagate over 

the canopy. Although the HARV had no flow visualization on the canopy the secondary separation 

lines are well defined up to the canopy. 

Figure 11. Forebody flow visualization of NASA HARV a = 34°, Moo = 0.3, Rec - 8.S x 106. 

The off-surface flow visualization of the HARV at a = 39°, and Moo = 0.24 is shown in 

figure 12. This flow visualization was obtained with smoke injection into the feeding sheet of the 

fore body vortex and natural condensation in the LEX vortex. The off-surface visualization 

(fig. 12) shows the vortices propagating over the canopy and interacting with the LEX vortex. It is 

presumed that this type of interaction is also present on the models at some conditions, but it 

appears not to be present at many of the conditions tested. 

13 
Figure 12. Off-surface smoke flow visualization of NASA HARV a = 39°, Moo = 0.24, Rec -

9.7 X 106. 

The forebody surface pressure distributions of the two models and the HARV at a = 40° are 

shown in figure 13. The areas of most interest are the forebody suction peaks (vortex footprint). 

These suction peaks indicate the presence of strong forebody vortices close to the surface, which 

generate high surface velocities and low pressures. These are evident at a = 40° station 142. The 

flight data also shows a faster pressure recovery than the models (75° :::; e :::; 135° and 225° :::; e :::; 

255° at FS 142). The wind-tunnel model data do not show these strong suction peaks (vortex footprint), but otherwise match the HARV data quite well. The character of the flight and wind tunnel 

forebody vortex suction peaks correlates well with what was seen in the flow visualizations; that 

is, the models at these conditions showed less vortical influence near the leeward plane of symmetry when compared with the HARV. 

While the wind tunnel model forebody pressure data matched flight well everywhere except 

for the forebody vortex suction peaks, the models did not correctly predict the full-scale flight 

lateral-directional behavior at a = 40° (maximum lift). This strongly implies that the forebody flow 

field contributes to this disparity in lateral stability. This forebody flow characteristic is consistent 

with the behavior seen on ogive cylinders at high angles of attack, which is strong vortex development at laminar and fully turbulent conditions and significantly less at transitional conditions.22- 24 

This implies that correct boundary-layer simulation on wind tunnel models is important to accurate 

prediction of airplane lateral-directional stability characteristics. 

14 
Cp 

Cp 

Cp 

-1.5 

-1.0 

-.5 

0 

.5 

1.0 

-1.5 

-1.0 

-.5 

0 

.5 

1.0 

-1.5 

-1.0 

-.5 

0 

.5 

Model Facility Moo 

o NASA 6% 7- x 10-- ft HST 0.30 

o NASA-2 1S% 30-- x SO-- ft 0.08 

£', HARV Flight 0.27 

f;{ 

@~ Station107 

2 

ti 

Station 142 

Station 184 

Rei; 

1.4 x 106 

1.0 x 106 

8.7 x 10S 

b. 

o 

o 

o 

1.0 ~ __ ~ ____ ~ __ ~ ____ ~ __ ~ __ ~ ____ ~ __ ~ ____ ~ __ ~ ____ ~ __ ~ 

o 30 60 90 120 150 180 210 240 270 300 

e, deg 

330 3S0 

960801 

Figure 13. Forebody surface pressure distribution ofF/A-I8 models and HARV a = 40°. 

15 
Mach Effects 

Effects of compressibility become apparent at transonic conditions when local shock waves develop. These shock waves cause premature boundary-layer separation and vortex bursting,25 and 

change the nature of the highly interactive separated flow field. The result is seen in nonlinear aerodynamics at high angles of attack. However, where significant local flow acceleration occurs 

effects can be seen at significantly lower free-stream Mach numbers. Highly cambered or swept 

surfaces with sharp leading edges (e.g. strakes and LEXs) may show some compressibility effects 

even below Moo = 0.3. 

The effects of Mach number on lateral-directional characteristics for the MD/Navy 6% model 

is shown in figure 14. These effects are only shown for the MD/Navy 6% model, since the 16% 

model was limited to low dynamic pressures. These tests were conducted at the David Taylor 

Research Center 7- x 10-ft Transonic Tunnel, which runs at reduced total pressure at higher Mach 

numbers. This reduced, but did not eliminate, variation in Reynolds number with Mach number. 

There are significant changes in the lateral-directional stability for increasing Mach numbers. Figure 14 shows the basic trend, which is an increase in lateral stability and a decrease in directional 

stability with increasing Mach number and increasing angle of attack at angles of attack a > 30° 

.002 

0 

I~ - .002 

- .004 

- .006 

.004 

.002 

cn ~ 0 

- .002 

-.004 

0 5 10 15 20 25 30 

Ct, deg 

35 40 45 50 

Moo 

o 0.40 

o 0.60 

o 0.70 

{:; 0.80 

~ 0.90 

55 60 

960802 

Figure 14. Effects of Mach number on lateral-directional stability of NASA 6% F/A-18 models. 

bjLE = 33°, bjTE = 0°, 0° ::; ::; 4°. 

16 
and Moo < 0.9. At Moo = 0.9 there is no change in lateral stability from Moo = 0.8 and directional 

stability increases slightly for increasing angles of attack. At high angles of attack Cl is a stronger influence on departure resistance than is Cn ' and it is typical for fighter-type coAfigurations 

to have directional instability at high angles of ~ttack. Gradual changes at lower Mach numbers 

are likely caused by locally transonic effects in the vortices. The large changes in C, between 

Moo = 0.6 and Moo = 0.7 are most likely a result of the formation of local shocks on thl wing and 

LEXs. In general the MD/Navy 6% model displays laterally stable behavior above a = 10° and 

through Moo = 0.9. This trend had typically been observed throughout the early development of 

the FI A-18 aircraft. 

The forebody and LEX pressure distributions at various Mach numbers at a = 40° are shown 

in figures 15(a) and 15(b) respectively. The LEX pressure distributions were obtained in the David 

Taylor Research Center 7- x lO-ft Transonic Tunnel. There is a significant effect of Mach number 

on the LEX surface pressure distributions (fig. IS(b)). The LEX pressure distributions decrease and 

flatten out with increasing Mach number. The effects of Mach number on the LEX pressures are 

seen to persist below Moo = 0.3, typically the lower limit of significant compressibility effects. This 

is most likely caused by the high accelerations around the leading edge of the LEX resulting in 

local transonic flow. The effects on the forebody are less pronounced than the LEX. The forebody 

surface pressures are relatively insensitive to Mach number up to Moo = 0.6. The Moo = 0.8 data 

shows the suction peaks associated with the development of strong forebody vortices. This change 

at Moo = 0.8 is likely a result of shock induced separation on the forebody. 

Effect of Reynolds Number and Forced Transition 

Reynolds number effects on lifting surfaces have long been known to have a significant impact 

on model-to-flight correlations. As previously shown boundary-layer transition and therefore Reynolds number effects on forebody surfaces also have a significant effect on ground-to-flight correlations at high angles of attack. Forced transition techniques analogous to those developed for 

attached flows on wings and other surfaces26 have been developed for forebodies at high angles of 

attack in conventional wind tunnels. lO By properly transitioning the forebody boundary layer, 

simulating higher Reynolds numbers, and testing at the correct Mach number it is possible to more 

accurately predict flight results. The grit pattern that was used in the reported FI A-18 tests included 

a standard nosering, which has been used for conventionallow/moderate angle-of-attack testing, 

and a so-called twin strip pattern shown in figure 16 in which two strips are placed longitudinally 

on both sides of the forebody approximately 54° from the windward attachment line. 

The effects of a fully turbulent boundary layer on the surface flow characteristics of the 

NASA 6% model in the NASA Langley 7- x lO-ft High Speed Tunnel were assessed by adding 

additional Ti02 particulates (to increase roughness) and testing at a higher dynamic pressure. The 

result is shown in figure 17 at a = 35° and Moo = 0.5, and many differences can be seen from that 

seen in figure 10. Recall figure 10 (Ree = 1 x 106, Moo = 0.22) which showed very weak structure 

on the lee side of the forebody, the structures terminated at the LEX apex and did not pass over 

the canopy. Figure 17 now shows the strong vortex features on the forebody propagating over the 

canopy at a = 35° and Moo = 0.5 (Red = 600,000). Also the previously seen pooling and wedges 

17 

-----------
------ ~-----------. ----- ----- ~------

-1.5 

-1 .0 

-.5 

Cp 

0 

.5 

1.0 

-1.5 

-1.0 

-.5 

Cp 

0 

.5 

1.0 

-1.5 

-1 .0 

- .5 

Cp 

0 

.5 

1.0 

0 30 60 90 120 150 180 210 240 270 300 

e, deg 

(a) Forebody. 

Moo 

o 0.20 

o 0.40 

o 0.60 

£:, 0.80 

330 360 

960803 

Figure 15. Effect of Mach number on surface pressure distributions of NASA 6% F/A-18 

model. a = 40°. 

18 
-3.0 

-2.5 

-2.0 

Cp -1.5 

Moo 

o 0.20 

o 0.30 

o 0.40 

[::, 0.60 

b. 0.70 

~ 

D 0.80 

0 0.90 

-1.0 

-.5 

o ~~--~--~~--~--~~--~--~~ 

-3.0 

-2.5 

-2.0 

-1.0 

-.5 

-3.0 

-2.5 

-2.0 

-1.5 

Cp 

-1.0 

-.5 

o ~~--~--~--~--~--~~--~--~--~ 

-1.0 -.8 -.6 - .4 -.2 0 .2 .4 .6 .8 1.0 

ylb' 

(b) LEX, upper surface. 

Figure 15. Concluded. 

960804 

19 
ASA -:LARC 

7 B 10H'ST 

Figure 16. Photograph of twin strip forebody transition grit on NASA 6% F/A- 18 model. 

Figure 17. Forebody flow visualization of NASA 6% FI A-18 model with forced transition. a = 35°, 

Moo = 0.5, Rec - 2.1 x 106. 

are no longer present and the separation lines are much more defined, an indication that they were 

likely generated by much higher surface shear and stronger vortical flow. This difference in flow 

visualization shows the local flow differences with a change in free-stream condition and boundary layer state. 

20 
At high angles of attack the local surface flow has a significant component in the cross-flow 

direction. Therefore, the twin strips are placed so that they transition each of the streamlines as they 

move around the body. Figure 18 shows the surface pressure distribution at fuselage stations 107, 

142, and 184 with and without this fixed cross-flow transition at a = 36.4° and Moo = 0.30. The 

comparison shows that the major effect is the much more evident vortex footprints (suction peaks 

-1.5 

- 1.0 

-.5 

Cp 

0 

.5 

1.0 

- 1.5 

-1.0 

- .5 

Cp 

0 

.5 <> 

1.0 

- 1.5 

-1.0 

- .5 

Cp 

0 

<> .5 

1.0 

0 

0 

<> 

<> 

30 60 

Station1 07 

Station 142 

Station 184 

Grit patte rn 

o Nosering only 

o Twin stri ps, nO.180 

~~<> O o 

<> 

<> 

<> 

<> 

90 120 150 180 210 240 270 300 330 360 

e, deg 960807 

Figure 18. Effect of fixed transition on forebody pressure distribution of NASA 6% F/A-18 model. 

a = 36.4°, Moo = 0.3, Re?; - 1.4 x 106. 

21 
at FS 142) on the leeward side when the crossflow is tripped. This, again, is evidence of the strength 

of the forebody vortices, which are attenuated at transitional conditions.22,24 A comparison of the 

6% model test at a = 40° and Moo = 0.30 with fixed cross-flow transition to the HARV is shown 

in figure 19. This shows a well-matched surface pressure distribution, with the forebody suction 

peaks just slightly lower for the NASA 6% model than for the HARV. 

-1.5 

-1.0 

-.5 

Cp 0 

0 0 

00 

.5 

1.0 

-1.5 

-1.0 

-.5 

Cp 

0 

<> 

.5 

1.0 

-1.5 

-1.0 

-.5 

Cp 

0 

0 

.5 

1.0 

0 30 60 90 

Model Facility 

o NASA 6% 7- x 10- ft HST 

o HARV Flight 

Station107 

Station 142 

Station 184 

120 150 180 210 240 270 300 

8,deg 

Rei; 

1.4 x 106 

13.5 x 106 

o 

0<0 

o 

o 

~O 

o 

330 360 

960808 

Figure 19. Comparison of forebody pressure distribution of NASA 6% F/A-18 model with fixed 

transition (twin strips no. 180 grit) to HARY. a = 40°, Moo = 0.3. 

22 

.-- - --. 
Figure 20 shows the effect of fixed forebody cross-flow transition on the lateral-directional 

stability of the NASA 6% model at a = 36.4° and Moo = 0.30, with both 33° (fig. 20(a)) and 25° 

(fig. 20(b)) leading-edge flaps. Figure 21 shows the effects of fixed transition on the NASA-2 

16% model at a = 40° and Moo = 0.08, with both 33° (fig. 21(a)) and 25° (fig. 21(b)) leading-edge 

flaps. At these conditions the predominant effect of the fixed forebody transition is to increase lateral stability (_40 < ~ < 4°). The direct effect of fixing crossflow transition is to increase the 

strength of the forebody vortices at high angles of attack. This indicates that, at least for the 

F/A-18 , fixing transition in sub-scale tests ha a favorable effect on the lateral stability when there 

is an interaction of the forebody vortices and the LEX flow field. This effect is more pronounced 

with the 25° leading-edge flap deflection than with the 34° deflection, probably because of the 

more separated and presumably weaker wing flow field with the 25° leading-edge flap deflection. 

In other words, with the 34° leading-edge flaps deflection the wing flow is much more attached 

and stronger and therefore more likely to resist the influence from the LEX vortices. The lateral 

stability changes with leading-edge flap deflection were much more pronounced with the NASA 

Grit pattern 

o Nosering only 

.025 o Twin strips, no. 180 

.020 .04 

.015 .03 

.010 .02 

.005 .01 

CI 0 Cn 0 

-.005 - .01 

- .010 - .02 

- .015 - .03 

- .020 -.04 

- .025 - .05 

-12 -8 - 4 0 4 8 12 -12 -8 -4 0 4 8 12 

~, deg ~, deg 960809 

(a) 0jLE = 33°. 

Figure 20. Effect of transition on lateral-directional stability of NASA 6% FI A-18 model. a = 

36.4°, Moo = 0.3, Rec - 1.4 x 106. 

23 
Grit pattern 

o Nosering only 

.025 o Twin strips, no. 180 

.020 .04 

.015 .03 

.010 .02 

.005 .01 

CI 0 Cn 0 

-.005 -.01 

-.010 -.02 

-.015 -.03 

-.020 -.04 

-.025 -.05 

-12 -8 -4 0 4 8 12 -12 -8 -4 0 4 8 12 

~,deg deg 960810 

(b) 8 jLE = 25°. 

Figure 20. Concluded. 

6% model than the NASA-2 16% model. This effect from the forebody flow field is dependent on 

several conditions and changes in any of these conditions can significantly effect the interaction. 

The effect of pressure ports on a model can be similar to grit, that is they can cause the flow to 

transition. Whether the ports significantly disturb the flow is dependent on the ize and density of 

the ports as well as the location. The NASA-2 16% model was fabricated with removable forebody 

sections both with and without pressure ports which made it possible to ascertain the effect of the 

pressure ports on the forebody . The forebody section was of the most interest, since variations in 

transition on this section has been seen as a significant factor on the high-angle-of-attack characteristics of the FI A-18. The effect of these pressure ports on the lateral-directional behavior is 

shown in figure 22. The effect of the pressure ports was very similar to that of gritting the forebody, 

such as shown in figure 21. The ports acted to transition the flow, similar to gritting, which 

strengthens the forebody vortices and subsequently increases lateral stability. 

24 
Grit pattern 

o Nosering only 

.025 o Twin strips, + nosering 

.020 .04 

.015 .03 

.010 .02 

.005 .01 

CI 0 Cn 0 

-.005 -.01 

-.010 -.02 

-.015 -.03 

-.020 -.04 

-.025 -.05 

-12 -8 -4 0 4 8 12 -12 -8 -4 0 4 8 DC 

p,deg p,deg 960811 

(a) 0jLE =33°. 

Figure 21. Effect of transition on lateral-directional stability of NASA-2 16% F/A-18 model. ex. = 

40°, Moo = 0.08, Rec - 1 x 106. 

Geometry Effects 

Geometry effects are an obvious area of concern for correlation between models and model to 

flight. Small geometry changes that have minimal impact at low angles of attack can have large 

and pronounced effects at high angles of attack. For sharp and high fineness ratio forebodies at high 

angles of attack, even small variations at the nose can change the vortex shedding and the subsequent flow field.27 The ramifications for high-angle-of-attack testing is that geometric fidelity in 

the forebody region warrants serious attention. The conventional sensitivities for low/moderate 

angle-of-attack testing must still be taken into consideration. 

25 
Grit pattern 

o Nosering only 

.025 o Twin strips, + nosering 

.020 .04 

.015 .03 

.010 .02 

.005 .01 

C1 0 Cn 0 

-.005 -.01 

-.010 -.02 

-.015 -.03 

-.020 -.04 

-.025 -.05 

-12 -8 -4 0 4 8 12 -12 -8 -4 0 4 8 12 

~,deg ~,deg 960812 

(b) 0 jLE = 25°. 

Figure 21. Concluded. 

Aft End Geometry Effects 

Subtle geometry differences other than those at the forebody can also have a significant impact 

on high-angle-of-attack characteristics. The MDfNavy 6% scale model was designed to be tested 

at high dynamic pressures, and as such required a support sting capable of supporting the large 

loads, especially normal force and pitching moment. The support sting was large and elongated in 

the vertical axis so the model aft end had to be flared (distorted) to allow sufficient clearance from 

the sting (fig. 23, both the MDfNavy and NASA 6% models had a common aft end). Although this 

slight flaring of the aft end would appear to have minimal consequences it was suspected for being 

partly responsible for the significant pitching moment differences between the 6% and 16% models 

(fig.7). At the higher angles of attack the NASA 6% model shows higher levels of nose down pitching moment than the NASA-2 16% model at a > 10°, even when the NASA 6% model was tested 

26 
Forebody 

o Pressure instrumented 

.025 o Non-pressure instrumented 

.020 .04 

.015 .03 

.010 .02 

.005 .01 

C1 0 Cn 0 

- .005 -.01 

-.010 -.02 

-.015 - .03 

-.020 -.04 

-.025 - .05 

-12 -8 -4 0 4 8 12 -12 -8 -4 0 4 8 12 

p,deg p,deg 960813 

Figure 22. Effect of pressure ports on lateral-directional aerodynamics of NASA-2 16% F/A-18 

model in the NASA Langley 30- x 60-ft Tunnel. a = 40°, Moo = 0.08, 0jLE = 33°, Rec - 1 X 106. 

Figure 23. Photograph of NASA 6% FI A-18 model showing aft end flare . 

27 
at similar Reynolds or Mach numbers. A test was conducted of the NASA-2 16% model with pieces added to roughly represent the flared aft end of the MDlNavy and NASA 6% models (fig. 23). 

The results of the NASA-2 16% model with the flare simulated (fig. 24) made up approximately 

half the difference in pitching moment at high angles of attack (25° < a < 45°) compared to the 

NASA 6% model. This indicates that the change in the back end geometry was likely responsible, 

at least in part, for this change in the high-angle-of-attack pitch behavior. 

Model Facility M Aft end Grit Rei: 

0 NASA-216% 30- x 60- ft 0.08 Not flared No. 36 1.0 x 106 

0 NASA 6% 7- x 10- ft HST 0.30 Flared No. 180 1.4 x 106 

<> NASA-216% 30- x 60- ft 0.08 Flared No. 36 1.0x106 

.3 

.2 

.1 

o 

-.1 

-.2 1--__ ..1--__ -'--__ -'-__ --'-__ ---' ___ "--_-----' 

-10 o 10 20 30 40 50 60 

ex, deg 960615 

Figure 24. Effect of aft end flare on pitching moment ofF/A-18 models. ~ = 0°, 8 jLE = 33°, 8 jTE = 

0°, 8HT = _12°, twin strips. 

Symmetric Forebody Strakes 

Controlling separation on the forebody has been investigated for use as a yaw control device at 

high angles of attack.28- 31 During one such investigation of the effectiveness of deployable 

forebody strakes on the forebody of the NASA-2 16% model, symmetric deployments of the forebody strakes were tested. Figure 25 shows the results of symmetric strake deflections up to 30°. 

Increasing symmetric deflections of the forebody strakes at the deflection angles shown, increases 

lateral stability at angles of attack from approximately 30° to 50°. This is most likely caused by the 

increasing strength of the forebody vortices and favorable interaction with the LEX vortices. The 

effect is most pronounced near a = 40°, the angle of attack of maximum lift, which is where the 

disparity in lateral stability between the models was greatest. 

28 
.001 Deflection, 

deg. 

0 0° 

0 0 10° 

0 20° 

f::::. 30° 

-.001 

CI 

~ -.002 

-.003 

-.004 

-.005 

0 10 20 30 40 50 60 70 80 90 

a, deg 960816 

Figure 25. Effect of symmetric forebody strake deflections on lateral stability of the NASA-2 16% 

F/A-18 model. Moo =0.08, bjLE =33°, bJTE =0°, ~~=±4°, Rec -1 x 106. 

Noseboom Effects 

The effect of adding a noseboom on a configuration can be extremely significant. Subtle changes in the geometry near the tip of a sharp or high fineness ratio forebody, at high angles of attack, 

can have a dramatic effect on the separation about that forebody, and subsequently the forebody 

flow (vortices) and downstream interactions. The noseboom has the combined effect of changing 

the tip geometry, increasing the effective fineness ratio, and propagating its own wake. This effect 

can be seen in the laser vapor screen off-surface flow visualization shown in figure 26. This was 

Figure 26. Flow visualization of NASA 6% F/A-18 forebody with and without noseboom. ex = 

50°, ~ = 0°, Moo = 0.6, FS = 184. Rec - 1.3 x 106. 

29 
obtained with the NASA 6% model at Moo = 0.6 and a = 50°, at the David Taylor Research Center 

7- x 10-ft Transonic Tunne1.25 Although this was obtained at an angle of attack and Mach number 

greater than that discussed previously, it clearly shows the effect of the noseboom on the forebody 

flow field. Without the boom a fairly symmetric pair of well developed vortices can be observed 

at a station just forward of the canopy junction. With the noseboom on, the forebody vortices are 

significantly diminished in size, no longer appear as two separate and distinct vortices, and a significant wake can clearly be seen, presumably generated by the boom itself. 

The effects of the noseboom on the forebody surface streamline of the NASA-1 16% model 

at a = 40° are shown in figure 27. Figure 27(a) shows the forebody surface flow without the noseboom and figure 27(b) shows the forebody surface flow at the same conditions with the noseboom. 

As would be expected with forebody vortices being attenuated, the surface streamlines are significantly softened when compared to those of the forebody without the boom. There are also weaker 

secondary separation lines that do not per ist over the canopy. Figure 28 shows the forebody surface pressure distribution of the NASA 6% model with and without the no eboom at a = 37° and 

Moo = 0.3. The major effect of the noseboom is to reduce the suction pressure peaks at FS 142, and 

show an indication of reduced forebody vortex strength. The lateral-directional characteristics of 

the NASA-2 16% and NASA 6% model with and without the noseboom, at a = 40°, are shown 

in figures 29 and 30, respectively. However, even though the effect of the boom on the forebody 

flow field is similar (reduced forebody vortex strength) between the two scale models, there is less 

of a sensitivity from the boom on the lateral-directional behavior of the NASA-2 16% than the 

NASA 6% with bjLE = 33°. 

Leading-Edge Flap Effect 

Leading-edge flaps can have a significant influence in this forebody/LEX flow interaction and 

therefore high-angle-of-attack lateral-directional stability. By reducing wing separation at a given 

angle of attack, leading-edge flaps can reduce the effects of the upstream flow field (i.e. the forebody and LEX flow fields). In the case of F/A-18 development, the lateral instability that was 

observed in the C L max region was alleviated by increasing the leading-edge flap deflection from 

25° to 33°. The effect of the leading-edge flap deflection and noseboom on lateral-directional stability of the two models is shown in figure 29 for the NASA-2 16% (a = 40°) model and figure 30 

for the NASA 6% model (a = 37°). In both cases the destabilizing effect of the boom was reduced 

by increasing the leading-edge flap deflection. As seen previously the effect was more pronounced 

on the NASA 6% model than on the NASA-2 16%. This delayed wing flow separation to a higher 

angle of attack, and lessened the unstable effects of the LEX vortices acting on the mostly separated flow over the wing surfaces. Since the forebody vortices interact favorably with the LEX and 

the noseboom disrupts the forebody flow field and attenuates the forebody vortices, the noseboom 

is a destabilizing influence. However, with the leading-edge flap deflection increased to 33° this 

destabilizing influence is mitigated as shown in figure 30(b). The NASA-2 16% model showed 

much less sensitivity in lateral aerodynamics from this change in flap deflection. 

30 
(a) Noseboom off. 

(b) Noseboom on. 

Figure 27. Forebody flow visualization of NASA-2 16% F/A-18 model with and without noseboom. ex = 40°, ~ = 0°, Moo = 0.08, Rec - 1 X 106. 

31 
-1.5 

-1.0 

-.5 

o 

.5 

<> 

<> 

<> Station107 

Noseboom 

o Off 

<> On 

1.0 '--_-'-_----' __ ...1....-_--'-_---:-'--_....L...._----'-__ "---_....L...._---' __ ....L...._~ 

-1.5 

-1.0 

-.5 

o 

.5 <> 

Station 142 

1.0 '--_--'-_----' __ ....1...-_--'-__ '--_-'-_--'-__ "'--_-'-_---' __ --'-_---' 

-1.5 

-1.0 

-.5 

o 

.5 

Station 184 

1.0 '--_-'-_----' __ ...1....-_--'-__ '--_-'-_--'-__ "'--_-'-_----' __ --'-_---' 

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

8, deg 960820 

Figure 28. Effect of nose boom on forebody surface pressure distribution of NASA 6% F/A-18 

model. a = 37°, ~ = 0°, Moo = 0.3, Rec - 1.4 x 106, twin strips no. 180 grit. 

32 
Noseboom 

0 Off 

.025 0 On 

.020 .04 

.015 .03 

.010 .02 

.005 .01 

C1 0 Cn 0 

-.005 -.01 

- .010 - .02 

-.015 - .03 

-.020 -.04 

-.025 -.05 

-12 -8 -4 0 4 8 12 -12 -8 -4 0 4 8 12 

~, deg ~, deg 960821 

(a) ()jLE = 33°. 

Figure 29. Effect of noseboom on lateral-directional aerodynamics of NASA -2 16% FI A -18 model. 

ex = 40°, P = 0°, Moo = 0.08, Rec - 1 x 106, twin strips no. 36 grit. 

33 

------
--- 

Noseboom 

o Off 

.025 o On 

.020 .04 

.015 .03 

.010 .02 

.005 .01 

C1 0 Cn 0 

-.005 -.01 

-.010 -.02 

-.015 -.03 

-.020 -.04 

-.025 -.05 

-12 -8 -4 0 4 8 12 -12 -8 -4 0 4 8 12 

~,deg ~,deg 960822 

(b) 0jLE = 25°. 

Figure 29. Concluded. 

34 

------------
Noseboom 

0 Off 

.025 0 On 

.020 .04 

.015 .03 

.010 .02 

.005 .01 

C1 0 Cn 0 

-.005 -.01 

-.010 -.02 

-.015 -.03 

-.020 - .04 

-.025 -.05 

-12 -8 -4 0 4 8 12 -12 -8 -4 0 4 8 12 

p,deg p,deg 960823 

(a) 0 jLE = 33°. 

Figure 30. Effect of nose boom on lateral-directional aerodynamics of NASA 6% F/A-18 model. 

a = 37°, ~ = 0°, Moo = 0.3, Rec - 1.4 X 106, twin strips no. 180 grit. 

35 
Noseboom 

0 Off 

.025 0 On 

.020 .04 

.015 .03 

.010 .02 

.005 .01 

C1 0 Cn 0 

-.005 -.01 

-.010 -.02 

-.015 -.03 

-.020 - .04 

-.025 -.05 

-12 -8 -4 0 4 8 12 -12 -8 -4 0 4 8 12 

~.deg ~.deg 960824 

(b) 8 jLE = 25°. 

Figure 30. Concluded. 

CONCLUDING REMARKS 

A series of ground based and flight investigations have been conducted on the FI A-18 configuration as part of the NASA High-Angle-of-Attack Technology Program. One of the objectives of 

these investigations was to determine the causes of disparities between ground and flight results. 

The re ults indicate that cros flow on the forebody, leading-edge extension (LEX) flow fields, 

nosebooms, and aft distortion can have a paramount effect on ground-to-flight correlation at high 

angles of attack. 

The differences seen in various sub-scale model tests at high angles of attack are due in part to 

the differences in forebody flow field and its subsequent interactions with the LEX and downstream flow fields. The forebody flow field is most sensitive to Reynolds number which drives its 

boundary layer state, and becomes increasingly important at high angles of attack. However, the 

LEX flow field is most sensitive to Mach number. It is difficult to match both conditions simultaneously during ground tests. Therefore, significant interactions at full-scale flight conditions may 

not be accurately predicted. 

36 
Fixing transltlon can simulate higher forebody Reynolds numbers, and therefore provide 

better predictions of full-scale flight results from sub-scale wind tunnel tests. At high angles of 

attack it is necessary to fix crossflow transition on the forebody in addition to conventional transition fixing for lower angles of attack. The twin strip technique has been shown to be effective on 

the F/A-18 configuration. 

Pressure ports and other subtle configuration differences can have a significant effect on the 

forebody flow field, and on the overall configuration flow field. The effect of pressure ports is to 

add roughness similar to grit. These are more apparent at high angles of attack where the forebody 

flow field becomes more significant. 

Geometric differences can often have pronounced effects at high angles of attack. Aft end distortions, due to wind tunnel support, on some F/A-18 models resulted in increased nose down 

pitching moments compared to the undistorted geometry. Symmetric forebody strakes can have a 

significant effect on the forebody vortices and increase lateral stability of the F/A-18 at angles of 

attack near that of maximum lift. Nose probes or other protuberances on the forebody, especially 

near the tip, have a significant influence on the forebody flow field. These influences can reduce 

forebody vortex strength at high angles of attack and cause a reduction in lateral stability on the 

F/A-18. Increasing leading-edge flap deflection at high angles of attack decreases wing flow separation and reduces the effects of upstream influences; that is the LEX and forebody vortex flows 

in the case of the F/A-I8. The combined effect of the nose probe and reduced leading-edge flap 

deflection can result in lateral instability at angles of attack near maximum lift for the F/A-18. 

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Nov. 1, 1992, pp. 23-59. 

2Lee, B.H.K, Brown, D., Zgela, M., and Poirel, D., "Wind Tunnel Investigation and Flight 

Tests of Tail Buffet on the CF-18 Aircraft," AGARD Specialists' Meeting on Aircraft Loads due to 

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3Meyn, L.A. and James, KD., "Full-Scale Wind Tunnel Studies ofF/A-I8 Tail Buffet," AIAA93-3519, Monterey, CA, 1993. 

4Meyn, L.A., James, KD., and Geenen, R.J., Correlation of FIA-J8 Tail Buffet Results, NASA 

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37 
l 

6Shah, Gautam H. , "Wind Tunnel Investigation of Aerodynamic and Tail Buffet Characteristics 

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8Fisher, David F, DelFrate, John H. , and Richwine, David M. , In-Flight Flow Visualization 

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9Erickson, Gary E. , Water Tunnel Flow Visualization and Wind Tunnel Data Analysis of the 

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loHall, RM. and Banks, D.W., Progress in Developing Gritting Techniques for High Angle of 

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llHall, RM., Erickson, G.E., Banks, D.W., and Fisher, D.F, "Advances in High-Alpha 

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"Comparison of Wind Tunnel and Flight Forebody Flow Fields on the NASA F-18 High-Alpha 

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18Fisher, David F, Banks, Daniel W. , and Richwine, David M. , F-18 High Alpha Research 

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38 
,---

19Fisher, David E, Richwine, David M., and Banks, Daniel W., Surface Flow Visualization of 

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22Lamont, P.J., The Effect of Reynolds Number on Normal and Side Forces on Ogive-Cylinders 

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23 Keener, Earl R , Flow-Separation Patterns on Symmetric Forebodies, NASA TM-86016, 

Jan. 1986. 

24Hunt, B.L., Asymmetric Vortex Forces and Wakes on Slender Bodies (Invited paper), 

AIAA-82-1336, Aug. 1982. 

25Erickson, Gary E., Wind Tunnel Investigation of Vortex Flows on FIA-18 Configuration at 

Subsonic Through Transonic Speeds, NASA TP-3111, Dec. 1991. 

26Braslow, A.L., Harris, RV. , Jr., and Hicks, RM., Use of Grit-Type Boundary-Layer Transition Trips on Wind-Tunnel Models, NASA TN-D-3579, Sept. 1966. 

27Moskovitz, Cary A., Dejarnette, ER, and Hall Robert M., Effects of Nose Bluntness, Roughness, and Surface Perturbations on the Asymmetric Flow Past Slender Bodies at Large Angles of 

Attack, AIAA-89-2236, Jan. 1989. 

28Murri, Daniel G., Biedron, Robert T., Erickson, Gary E., Jordan, Frank L., Jr. , and Hoffler, 

Keith D., "Development of Actuated Forebody Strake Controls for the F-18 High Alpha Research 

Vehicle," NASA CP-3149, pp. 335-380. Presented at the NASA High-Angle-of-Attack Technology 

Conference, Hampton, Virginia, Oc.t 3D-Nov. 1, 1990. 

29Erickson, Gary E., and Murri, Daniel G., "Forebody Strakes for High-Angle-of-Attack Vortex 

Flow Control - Mach Number and Strake Planform Effects," NASA CP-3149, pp. 381-480. 

Presented at the NASA High-Angle-of-Attack Technology Conference, Hampton, Virginia, 

Oct. 3D-Nov. 1, 1990. 

30Murri, Daniel G. , Shah, Gautam H., DiCarlo, Daniel J., and Trilling, Todd W., "Actuated 

Forebody Strake Controls for the F-l8 High-Alpha Research Vehicle," Journal of Aircraft, vol. 32, 

no. 3, May-June 1995, pp. 555-562. 

31Murri, Daniel G., Fisher, David E, and Lanser, Wendy R, "Effect of Actuated Forebody 

Strakes on the Forebody Aerodynamic of the NASA HARV," Presented at the NASA High-Angleof-Attack Technology Conference , Hampton, Virginia, Sept. 17-19, 1996. 

39 

J 
REPORT DOCUMENTATION PAGE Form Approved 

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1. AGENCY USE ONLY (Leave blank) 1 2. REPORT DATE 1 3, REPORT TYPE AND DATES COVERED 

January 1997 Technical Memorandum 

4, TITLE AND SUBTITLE 5. FUNDING NUMBERS 

The FI A-18 High-Angle-of-Attack Ground-to-Flight Correlation: Lessons 

Learned 

6. AUTHOR(S) WU 505-68-30 

Daniel W. Banks, David F. Fisher, Robert M. Hall, Gary E. Erickson, 

Daniel G. Murri, Sue B. Grafton, and William G. Sewall 

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 8. PERFORMING ORGANIZATION 

REPORT NUMBER 

NASA Dryden Flight Research Center 

P.O. Box 273 H-2149 

Edwards, California 93523-0273 

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORINGIMONITORING 

AGENCY REPORT NUMBER 

National Aeronautics and Space Administration 

Washington, DC 20546-0001 NASA TM-4783 

11 . SUPPLEMENTARY NOTES 

Presented at NASA Langley High-Angle-of-Attack Technology Conference, Langley Research Center, Hampton, Virginia, 

September 17-19, 1996, Daniel Banks and David Fisher, Dryden Flight Research Center, Edwards, CA; Robert Hall, Gary 

Erickson, Daniel Murri, Sue Grafton, and William Sewall, NASA Langley Research Center, Hampton, VA. 

12a. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE 

Unclassified- Unlimited 

Subject Category 02 

13. ABSTRACT (Maximum 200 words) 

Detailed wind tunnel and flight investigations were performed on the F/A-18 configuration to 

explore the causes of many high-angle-of-attack phenomena and resulting disparities between wind 

tunnel and flight results at these conditions. Obtaining accurate predictions of full-scale flight aerodynamics from wind-tunnel tests is important and becomes a challenge at high-angle-of-attack conditions 

where large areas of vortical flow interact. The FI A-18 airplane was one of the first high-performance 

aircraft to have an unrestricted angle-of-attack envelope, and as such the configuration displayed many 

unanticipated characteristics. 

Results indicate that fixing forebody crossflow tran ition on models can result in a more accurate 

match of flow fields, and thus a more accurate prediction of aerodynamic characteristics of flight at 

high angles of attack. The wind tunnel results show that small geometry differences, specifically 

nosebooms and aft-end distortion, can have a pronounced effect at high angles of attack and must be 

modeled in sub-scale tests in order to obtain accurate correlations with flight. 

14. SUBJECTTERMS 

Fighter aircraft; Forebody flows; High angle of attack; Vortex flows; Vortex 

interactions 

17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 

OF REPORT 

Unclassified 

NSN 7540-01-280-5500 

OFTHIS PAGE OF ABSTRACT 

Unclassified Unclassified 

Available from the NASA Center for AeroSpace InformatIOn, 800 Elkndge Landmg Road, 

Linthicum Heights, MD 2 1090; (301)621-0390 

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43 

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A03 

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Unlimited 

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298·' 02