NASA Technical Memorandum 4774

Effect of Actuated Forebody

Strakes on the Forebody

Aerodynamics of the NASA

F-18 HARV

David F. Fisher, Daniel G. Murri, and Wendy R. Lanser

October 1996

NASA Technical Memorandum 4774

Effect of Actuated Forebody

Strakes on the Forebody

Aerodynamics of the NASA

F-18 HARV

David F. Fisher, Daniel G. Murri,

and Wendy R. Lanser

Dryden Flight Research Center

Edwards, California

National Aeronautics and

Space Administration

Office of Management

Scientific and Technical

Information Program

1996

EFFECT OF ACTUATED FOREBODY STRAKES ON THE FOREBODY

AERODYNAMICS OF THE NASA F-18 HARV

David E Fisher

NASA Dryden Flight Research Center

Edwards, CA

Daniel G. Murri

NASA Langley Research Center

Hampton, VA

Wendy R. Lanser

NASA Ames Research Center

Moffett Field, CA

ABSTRACT

Extensive pressure measurements and off-surface flow visualization were obtained on the forebody and strakes of the NASA F- 18 High Alpha Research Vehicle (HARV) equipped with actuated

forebody strakes. Forebody yawing moments were obtained by integrating the circumferential

pressures on the forebody and strakes. Results show that large yawing moments can be generated

with forebody strakes. At angles of attack greater than 40 °, deflecting one strake at a time resulted

in a forebody yawing moment control reversal for small strake deflection angles. At tx = 40 °

and 50 °, deflecting the strakes differentially about a 20 ° symmetric strake deployment eliminated

the control reversal and produced a near linear variation of forebody yawing moment with differential strake deflection. At ct = 50 ° and for 0 ° and 20 ° symmetric strake deployments, a larger forebody yawing moment was generated by the forward fuselage (between the radome and the apex

of the leading-edge extensions), than on the radome where the actuated forebody strakes were

located. Cutouts on the flight vehicle strakes that were not on the wind tunnel models are believed

to be responsible for deficits in the suction peaks on the flight radome pressure distributions and

differences in the forebody yawing moments.

NOMENCLATURE

ANSER

cn0

Cno, 

Cno, 

actuated nose strakes for enhanced rolling

aircraft yawing moment at 13= 0 °

forebody (ES. = 60 to 190) yawing moment at [3 = 0 ° from integrated pressures

forward fuselage (ES. = 124.5 to 190) yawing moment at 13= 0 ° from integrated

pressures
Cno, rad

Cp

CFD

F.S.

HARV

HATP

M

P

P0

q_,

Re c

y/b

AC n _30Orud

_S

5s ,d

_s, L/R

0

radome (RS. = 60 to 124.5) yawing moment at 13= 0 ° from integrated pressures

pressure coefficient, ( p - PO )/q_

computational fluid dynamics

fuselage station, in.

pressure altitude, ft

High Alpha Research Vehicle

High-Alpha Technology Program

Mach number

local pressure, lb/ft 2

free-stream static pressure, lb/ft 2

free-stream dynamic pressure, lb/ft 2

Reynolds number based on mean aerodynamic chord of 11.525 ft

strake span location divided by the local strake span

aircraft angle of attack, deg, from left wingtip angle-of-attack vane corrected for

upwash and boom bending

aircraft angle of sideslip, deg, average of left- and right-wingtip sideslip vanes

corrected for angle of attack, deg

incremental change in aircraft yawing moment coefficient due to a -30 ° deflection

of both rudders

single strake deflection, deg

differential strake deflection, right strake deflection minus left strake deflection, deg

left and right strake deflection measured from the retracted position, deg

forebody cross-section circumferential angle, deg (0 ° is bottom centerline, positive

is clockwise as seen from a front view, 0 ° to 360 °)

INTRODUCTION

The NASA High-Alpha Technology Program (HATP) was initiated to increase understanding,

improve prediction techniques, provide design guidelines, and investigate new concepts for controis effectors on advanced, highly maneuverable aircraft at high angles of attack. 1 This program

uses the F-18 configuration as a validation and demonstration vehicle. The flight vehicle consists

of a highly modified F-18 referred to as the High Alpha Research Vehicle (HARV). 2 The HATP

consists of wind-tunnel tests of subscale and full-scale models and components, calibration for

computational fluid dynamics codes, piloted simulations, and full-scale flight testing. New ground

2
test,computationalfluid dynamics,andflight testresultshavebeenobtainedandtechniqueshave

beendevelopedasaresultof thisprogram.Theseresultsandtechniqueshavebeensummarizedin

threehighangle-of-attackconferencepublications.3-5

Severalnewconceptsfor control at high anglesof attackhavebeeninvestigatedin thewind

tunnelandwith computationalfluid dynamics.Suchconce]_tsincludeenginethrustvectoring6and

forebodyvortex control using actuatedforebodystrakes/,8andpneumaticblowing.9'10An inflight closed-loopthrust-vectoringsystemhasbeeninstalledonthe HARV, andthe resultshave

beendocumented.11Recently,closed-loopactuatedforebodystrakeshave beeninstalled, and

preliminary resultsarereportedin the presenthigh-angle-of-attackconference.Although not a

partof HATP, an open-looppneumaticblowing flight control systemwas flight-testedon the

X-29A airplaneat highanglesof attack.12'13

This paper examines the local flow aerodynamics of the forebody strakes and their effect on

the forebody aerodynamics of the aircraft. Local forebody pressure distributions and forebody

yawing moment results obtained on the F-18 HARV during the Actuated Nose Strakes for

Enhanced Rolling (ANSER) program are presented. The ANSER acronym refers to "rolling" because the strakes provide the critical yaw control required to coordinate rolling maneuvers about

the velocity vector at high angle-of-attack conditions. 14,15 Forebody yawing moments are presented for maximum single strake deployments for a = 200-65 °. Forebody yawing moments are also

presented for four symmetric strake deployments for a = 30 °, 40 °, 50 °, and 60 °. Pressure distributions are presented at a = 50 ° for two symmetric strake deployments. Comparisons with ground

facility results for a full-scale F-18 model are included.

EXPERIMENT DESCRIPTION

The following description of the vehicle and instrumentation used in this experiment supplied

the basis for data collected.

Vehicle Description

Figure 1 shows the F-18 HARV. This highly modified full-scale development twin-engine,

single-place, fighter/attack (F/A) airplane was originally built for the U.S. Navy by the McDonnell

Douglas Corporation (St. Louis, Missouri) and the Northrop Corporation (Newbury Park,

California). The F-18 HARV is powered by two modified General Electric (Lynn, Massachusetts)

F404-GE-400 afterburning turbofan engines rated at approximately 16,000 lbf static thrust at

sea level.

The F-18 HARV features a midwing configuration with a wing-root leading-edge extension

that extends from the forward portion of the fuselage and blends into the wing. The F- 18 HARV

carries no stores or missiles and is highly instrumented for research purposes. The wingtip launching rails and missiles were replaced with specially designed camera pods and airdata sensors. 16

More specific details of the F- 18 HARV can be found in reference 2.
EC9643540-29

Figure 1.F-18HARV aircraftwithANSERradome,o_-- 30", _is,/dR = 0°/90°"

The HARV was modified by adding externally mounted engine thrust vanes for the deflection

of the exhaust to provide additional pitching and yawing moments. The engines were modified to

accommodate the thrust-vectoring vane installation by removing the divergent flap portion of the

nozzle. The thrust-vectoring capability was provided by controlled deflection of the vanes (three

for each engine), which moved into the engine exhaust plume (ref. 2).

For the ANSER program, a new radome was fabricated at NASA Langley Research Center

(LaRC), Hampton, Virginia, that incorporated hydraulically operated conformal actuated forebody

strakes. Figure 2 shows a schematic of the strakes. These strakes are 48 in. long and are positioned

longitudinally 120 ° up from the bottom of the forebody beginning at 8 in. aft of the forebody apex.

Cove area

Strake retracted _ ._ 90 o

_ 1_ Conformal strake \ / ] deployed

I _ _ _ / 12_o

s I .0om°n°o° /

' L.?"' LJ

Secl|on A-A

960613

Figure 2. Sketch of forebody strakes on the F- 18 HARV.

4
Figure3 showsacloseupphotographof the left strake. When the strakes are closed, there are no

protrusions, and the external mo!d line of the radome conforms to the original radome shape.

Figures 4(a) and 4(b) show the wrplane with the right strake (pilot view) deployed at 30 ° and

90 ° respectively.

Instrumentation

The fuselage forward of the cockpit was extensively instrumented with surface pressure measurements (fig. 5). Five circumferential rings of pressure orifices were installed on the surface of

the radome and forward fuselage, forward of the cockpit canopy at fuselage station (F.S.) 70,

F.S. 85, F.S. 107, F.S. 142, and F.S. 184. On each forebody strake, three rows of orifices were installed on the outboard surface at F.S. 70, F.S. 85, and F.S. 107, and two on the inboard surface at

F.S. 85 and F.S. 107. In addition, three rows of orifices (F.S. 70, F.S. 85, and F.S. 107) were installed on the cove section of the forebody. This region of the forebody is exposed when the strake

is actuated. The forebody pressures were measured with a +216 lbf/ft 2 differential pressure transducer, and the strakes used +720 lbf/ft 2 transducers. The accuracy was estimated to be +1 lbf/ft 2

for the forebody pressures and +3 lbf/ft 2 for the strake pressures. More detailed information of the

instrumentation has been given previously. 17

Airspeed, altitude, angle of attack (o0 and angle of sideslip (13) were measured using airspeed

booms mounted on specially designed wingtip photograph pods. A swiveling probe that was designed with four vanes to align the pitot-static probe head with the local airstream was on the

left wingtip. 16 Aircraft angle of attack was measured by using a vane on the right wingboom and

Figure 3. Close-up photo of left forebody strake.
EC9643598-9

(a)8s,UR = 0°/30°'

Figure 4. F-18 HARV with ANSER radome and right forebody strake deployed.

6
(b) 8s,/JR= 0°/90°"

Figure4. Concluded.

7
Nose ring

pressure

F.S. 70

F.S. 85

F.S. 107

F.S. 142

F.S. 184

Figure 5. Location of pressure orifices on F-18 HARV forebody.

960617

corrected for upwash and boom bending. Angle of sideslip was determined as the average of the

left- and right-wingboom sideslip vane measurement corrected for angle of attack.

Data from these measurements, as well as from the standard aircraft control positions, inertial

systems, and accelerometer parameters, were transmitted to a ground station. These critical

parameters were monitored by engineers and technicians in real time on strip charts and displays.

These data were monitored in near real time at NASA Dryden Flight Research Center, Edwards,

California, and Langley Research Center, Hampton, Virginia.

On selected flights, smoke flow visualization was used to mark and identify the off-surface

forebody/strake vortical flows. For these flights, the nose cap on the ANSER radome was replaced

with one that had two 1.0 in. diameter smoke ports, placed symmetrically 60 ° up from the bottom

centerline. Figure 6 shows a closeup of the right smoke port and radome after a flight. White smoke

from a smoke generator system 18 used previously 19 on the F-18 HARV was fed from the smoke

generator system through a single 1.5-in. diameter tube to the two 1.O-in. diameter symmetrically

located ports on the nose cap. Also shown in figure 6 are six cutouts on the lower portion of the

right strake that provided clearance for the radome bulkheads. These cutouts will be discussed later

in Results and Discussion.

DATA REDUCTION TECHNIQUES

For the forebody and strake pressures, in-flight zero differential pressures were taken before

each test point and were used in postflight data reduction to correct the data for calibration offsets.

8
Figure6. Close-upof right smokeport andstrakecutoutsonANSERradome.

Time segmentsof 1.0 secdurationwere usedfor dataanalysispurposeswith approximately

25timepointsaveraged.

Forebodyyawingmomentcoefficientsweredeterminedby integratingthefive rings of pressureson theforebody,thepressuresonthecoves,andthepressuresonthestrakesovertheirrespectiveprojectedforebodysideareas.Fortheforebody,thisareaextendedfrom theforebodyapex

to theapexof thewing leading-edgeextensions.Pressureswereintegratedfor flight conditionsin

whichtheangleof sideslipwaslessthan-I-0.5 ° to determine the yawing moment at 0 ° sideslip.

During data analysis, pressure coefficients were noted on the lower fuselage centerline that

were greater than 1.0 during extremely high-angle-of-attack flight, (x > 60 °. This anomaly had not

been noted in earlier flight tests. 17,2° At these angles of attack, the aircraft is no longer able to

maintain constant altitude. In fact, the aircraft descended at a rate of 200-300 ft/sec. During

subsequent hangar tests of the swiveling probe pitot- and static-pressure orifices, an approximate

0.3-sec lag was noted. Possibly, the flexible pressure tubing for the swiveling probe pitot and static

pressures had become restricted as compared to previous flight data. For the flight conditions of

the data within this paper, (M < 0.4 and altitudes near 25,000 ft), this lag was estimated to be approximately 0.6 sec after adjusting for the change in density and viscosity at altitude. 21 Therefore,

for these results, the pitot static data were time skewed by 0.6 sec.

FLIGHT TEST CONDITIONS

Data were obtained in quasi-stabilized 1-g flight maneuvers at a nominal altitude of 25,000 ft

and M < 0.4 for o_ = 200-65 °. At angles of attack greater than approximately 30 °, constant altitude

9
could not be maintained, and these data were obtained in a descent. The pressure distribution data

were obtained during open-loop flight maneuvers, and the flow visualization data were obtained

during open- and closed-loop maneuvers. For the data presented, with few exceptions, 1131< 0.5 °.

Table 1 lists the test points and flight conditions.

Table 1. Flight conditions for HARV ANSER pressure distribution test points.

a, 9, Rec 8s' 8s'

Test point deg deg M E-6 qo* hp right left 5s,d

flt0336.nz.02 20 0 0.328 12.2 59.4 24886 --0.3 89.6 -89.9

flt0336.nz.04 20 -0.1 0.35 13.07 67.7 24923 89.6 --0.3 89.9

flt0333.nz.02 24.6 0.3 0.303 12.73 58.6 21545 -0.2 89.7 -89.9

flt0343.nz.08 25.1 0 0.335 12.42 61.4 25070 -0.5 -0.5 -0.1

flt0333.nz.06 25 0.8 0.27 11.68 47.9 20073 89.6 -0.2 89.8

flt0333.nz.01 29.6 0 0.288 11.79 51.1 22302 -0.2 89.7 -89.8

flt0343.nz.07 30 0.3 0.273 9.52 37.3 27162 -0.5 59.7 -60.2

flt0333.nz.05 29.7 0.5 0.288 12.07 52.7 21634 -0.3 29.7 -29.9

flt0336.nz.ll 30.2 0.2 0.259 10.46 41 22598 -0.3 19.7 -20

flt0352.nz.03 32.6 0.2 0.27 10.24 40.1 24915 4.6 14.7 -10.1

flt0336.nz.01 30.1 0.1 0.294 10.8 46.7 25345 -0.2 9.7 -9.9

flt0335.nz.12 29.8 0.1 0.294 11.23 47 25306 --0.6 -0.6 0

flt0352.nz.02 30.2 0 0.309 12 54 24402 9.7 9.7 0

flt0336.nz.03 30.2 0 0.323 11.42 53.4 26631 9.7 -0.2 10

flt0352.nz.04 32.3 0.1 0.252 9.87 36.7 23926 14.7 4.6 10.1

flt0336.nz.13 29.9 0.1 0.273 10.84 44.5 23027 19.7 -0.4 20.1

flt0336.nz.06 29.8 -0.2 0.304 10.95 48.5 26077 29.7 -0.3 30

flt0336.nz.07 30.3 0.1 0.328 11.56 54.7 26801 59.7 -0.3 60

flt0336.nz.08 29.8 -0.2 0.297 11.24 49.3 24625 89.6 -0.3 90

flt0335.nz.02 35.1 -0.2 0.307 10.4 42.6 29380 -0.2 89.6 -89.9

flt0335.nz.03 35.1 -0.1 0.301 10.65 43.8 27904 -0.2 -0.2 0

flt0336.nz.09 35.2 --0.2 0.259 10.42 40.6 22809 89.6 -0.3 90

flt0335.nz.21 39.9 0 0.27 10.18 38.4 25903 -0.6 89.6 -90.2

flt0334.nz.02 40 0.1 0.279 10.42 41.3 25779 -0.2 59.7 -59.9

flt0337.nz.09 40.6 0.3 0.24 9.72 34.9 22739 4.6 34.7 -30

flt0334.nz.06 39.3 -0.4 0.284 10.03 39.6 27582 -0.3 29.6 -29.9

fit0337.nz.03 39.9 0 0.278 9.92 39.5 26615 9.7 29.7 -19.9

flt0351.nz.01 40 0.3 0.29 10.18 39.8 28398 -0.2 19.7 -19.9

flt0352.nz.05 39.2 -0.3 0.268 10.05 38.7 25372 4.6 14.7 -10.1

flt0336.nz.10 40.6 -0.1 0.258 9.76 37.3 24617 -0.3 9.7 -9.9

flt0338.nz.02 39.7 -0.1 0.29 10.5 43.5 26332 14.7 24.7 -9.9

flt0338.nz.03 39.8 -0.1 0.271 10.21 40.1 25133 -0.2 -0.2 0

flt0352.nz.01 40 -0.1 0.272 9.64 36.9 27192 9.7 9.7 0

flt0337.nz.13 40 0.2 0.241 9.8 35.5 22531 19.7 19.7 0.1

flt0338.nz.05 39.6 -0.1 0.3 10.44 43.9 27746 24.7 14.7 10

flt0336.nz.12 39.9 -0.1 0.273 10.21 41.3 24907 9.7 -0.4 10.1

flt0352.nz.06 39.6 -0.1 0.263 10.02 38.1 25038 14.7 4.5 10.2

flt0333.nz.10 40 0.1 0.201 8.87 27.4 20163 19.7 -0.3 20

Cn0, fb

Cn0,

rad Cn0, ff

0.0033 0.0020 0.0013

-0.0032 -0.0018 -0.0014

0.0059 0.0032 0.0027

0.0001 0.0002 -0.0001

-0.0033 -0.0034

0.0049 0.0058

-0.0068

0.0108

0.0084 0.0037 0.0048

0.0033 0.0012 0.0021

0.0015 0.0006 0.0009

0.0007 0.0001 0.0005

0.0001 0.0003 -0.0002

0.0003 0.0003 0.0000

0.0000 0.0001 0.0000

-0.0005 0.0000 -0.0004

-0.0009 -0.0001 -0.0008

-0.0019 -0.0004 -0.0016

-0.0034 -0.0009 --0.0026

-0.0083 -0.0032 -0.0051

-0.0103 -0.0044 -0.0059

0.0162 0.0070 0.0092

0.0003 0.0003 0.0000

--0.0156 -0.0063 -0.0094

0.0218 0.0093 0.0125

0.0173 0.0062 0.0111

0.0090 0.0022 0.0068

0.0056 0.0016 0.0039

0.0070 0.0018 0.0052

0.0000 0.0002 -0.0002

0.0001 0.0001 0.0(0

-0.0023 0.0000 -0.0022

0.0035 0.0007 0.0028

0.0005 0.0005 0.0000

0.0010 0.0004 0.0006

0.0000 0.0002 -0.0001

-0.0032 -0.0003 -0.0028

0.0019 0.0007 0.0012

0.0011 0.0006 0.0006

-0.0009 0.0004 -0.0013

10
Table 1. Continued.

o_, fS, Rec ks' ks'

Test point deg deg M E-6 q,,_ hp right left _Ss,d Cn0, fb

flt0334.nz.07 40.4 0 0.255 9.73 35.7 25100 29.7 -0.2 29.9 -0.0059

fitO337.nz.06 40.3 0.1 0.256 10.09 38.4 23469 34.7 4.5 30.2 -0.009

fltO334.nz.04 40 -0.1 0.285 9.9 38.9 28168 59.7 -0.2 59.9 --0.0171

flt0338.nz.06 39.8 0.1 0.277 9.93 39.6 26449 89.6 -0.3 89.9 -0.0197

flt0335.nz.25 44.8 0.1 0.301 10.12 40.3 29833 -0.6 89.7 -90.2 0.0297

flt0337.nz.05 44.9 0 0.265 9.86 37.9 25413 -0.1 -0.1 0 0.0000

flt0335.nz.27 44.5 0.3 0.313 11.2 47.4 27928 89.6 -0.5 90.2 -0.0286

flt0336.nz.14 50.2 -0.2 0.262 9.32 35.5 26365 -0.3 89.6 -90 0.0337

flt0335.nz.19 50 0.2 0.243 10.12 36.6 22142 -0.6 59.7 -60.3 0.0241

fltO337.nz.08 49.5 0.2 0.253 9.66 35.5 24830 4.6 34.7 -30 0.0127

flt0337.nz.07 49.8 -0.1 0.315 9.35 38.9 32539 -0.2 29.7 -29.8 0.0017

flt0343.nz.04 49.7 -0.1 0.276 9.97 39.7 26274 9.7 29.7 -20 0.0076

flt0352.nz.09 48 0.2 0.261 9.84 36.6 25529 19.7 39.6 -19.9 0.0152

flt0351.nz.02 49.8 0.1 0.257 10.56 38.5 23680 -0.1 19.7 -19.9 -0.0091

flt0352.nz.07 49.3 -0.3 0.263 9.39 34.6 27206 24.7 34.7 -10 0.0111

fit0335.nz.16 49.7 0.1 0.296 10.38 41.9 28185 14.7 24.7 -10 0.0035

flt0338.nz.01 50.1 0 0.292 9.07 35.3 31342 14.7 24.7 -9.9 0.0028

flt0333.nx.01 49.9 -0.1 0.248 10.04 37 22878 -0.2 9.7 -9.9 -0.0081

fit0337.nz.ll 49.8 0.2 0.271 9.24 34.9 28272 -0.3 -0.2 -0.1 0.0001

flt0357.nz.03 50.9 0.2 0.241 9.29 31.3 25523 29.7 29.6 0 -0.0010

flt0335.nz.17 50 0.3 0.256 10.18 37.3 24288 19.7 19.7 0 0.0017

flt0333.nx.02 49.7 -0.2 0.241 9.94 36 22205 9.7 -0.2 9.9 0.0084

fltO352.nz.08 48.7 -0.7 0.24 9.56 32.9 23253 34.7 24.7 10 -0.0095

fltO335.nz.18 49.8 -0.2 0.237 10.11 35.9 21429 24.7 14.7 10 -0.0035

fltO352.nz.10 50.2 -0.4 0.227 9.45 31.6 22374 39.7 19.7 20 -0.0162

fltO335.nz.23 50 0 0.279 10.19 38.7 27274 29.7 9.7 20 -0.0074

flt0339.nz.06 49.7 0.2 0.257 9.74 36.5 24929 19.7 -0.3 20.1 0.0082

fltO357.nz.02 47.2 -0.4 0.235 10.81 38.2 19568 44.7 14.7 30 -0.0184

flt0335.nz.29 50.1 0 0.248 10.19 36.8 23057 34.7 4.6 30.1 -0.0127

fit0335.nz.15 49.3 0 0.265 10.88 42.4 22879 29.7 -0.6 30.2 -0.0040

flt0337.nz.12 50.1 0.1 0.257 9.32 34.2 26369 59.7 -0.2 60 -0.0264

flt0336.nz.15 50.2 0.1 0.242 9.4 33.9 23909 89.7 -0.4 90 -0.0345

flt0341.nz.09 55.1 -0.4 0.257 9.67 36 25255 -0.2 89.7 -89.9 0.0368

flt0341.nz.07 55.2 0 0.25 8.87 31.8 26772 -0.2 -0.2 0 -0.0002

flt0341.nz.10 55.1 0.4 0.233 9.59 33.3 22367 89.6 -0.2 89.8 -0.0379

flt0342.nz.06 59.8 -0.1 0.276 11.07 45.8 22850 -0.5 89.7 -90.2 0.0356

flt0342.nz.04 59.7 0.4 0.287 11.22 47.5 23839 -0.5 59.7 -60.2 0.0203

fit0375.nz.07 58.2 -0.2 0.267 10.22 40.8 24087 14.7 44.7 -29.9 0.0247

flt0342.nz.01 60.8 0.1 0.271 11.05 45.3 22288 -0.3 19.7 -20 -0.0112

flt0375.nz.14 60.4 0.2 0.282 9.37 37.7 28348 19.7 39.6 -19.9 0.0219

flt0343.nz.05 59.9 0.4 0.298 10.35 43.9 27434 -0.5 9.7 -10.1 -0.0084

flt0358.nz.01 60.2 0.2 0.284 11.4 45.1 24594 24.7 34.7 -10 0.0210

flt0357.nz.04 59.1 0.1 0.23 9.9 33 21994 29.7 29.6 0 0.0087

flt0358.nz.02 59.8 0.3 0.264 10.76 40.4 23821 34.7 24.7 10 -0.0120

Cn0,

rad Cno, ff

-0.0007 -0.0051

-0.0021 -0.0069

-0.0053 -0.0118

-0.0068 -0.0129

0.0122 0.0175

0.0003 -0.0002

-0.0106 -0.0180

0.0149 0.0187

0.0093 0.0148

0.0024 0.0103

0.0006 0.0011

0.0015 0.0061

0.0043 0.0110

-0.0012 -0.0079

0.0025 0.0086

0.0007 0.0028

0.0007 0.0022

-0.0011 -0.0070

0.0002 -0.0001

-0.0001 --0.0009

0.0002 0.0015

0.0019 0.0065

-0.0018 -0.0077

-0.0006 -0.0029

-0.0043 -0.0119

-0.0014 -0.0060

0.0019 0.0063

-0.0051 -0.0133

-0.0023 -0.0104

-0.0005 -0.0035

-0.0092 -0.0172

-0.0138 -0.0207

0.0180 0.0189

0.0002 -0.0004

-0.0170 -0.0209

0.0210 0.0146

0.0121 0.0082

0.0087 0.0160

-0.0068 -0.0044

0.0080 0.0139

-0.0014 -0.0070

0.0058 0.0153

0.0015 0.0072

-0.0025 -0.0095

11
Table 1. Concluded.

(x, 9, Rec ks' 6s'

Test point deg deg M E-6 q*_ hp right left fis,d

flt0343.nz.06 59.9 -0.3 0.284 11.04 46.2 24118 9.7 -0.5 10.2

flt0342.nz.02 59.3 0.5 0.248 8.76 31.6 26531 19.2 -0.3 19.5

flt0375.nz.05 59.9 -0.2 0.284 10.47 43.7 25235 39.7 19.7 20

flt0342.nz.03 59.7 0.2 0.266 9.35 36.2 26704 29.7 -0.3 30

fit0342.nz.05 59.6 0.4 0.288 11.11 47.1 24213 59.7 -0.5 60.2

flt0342.nz.07 60.1 0.5 0.28 10.29 42.6 25871 89.6 -0.4 90

flt0342.nz.09 64.6 0.7 0.3 11.1 48 24933 -0.5 89.7 -90.1

fh0343.nz.02 64.5 1.7 0.28 10.5 41.5 26253 -0.3 -0.3 0

flt0342.nz.10 64.9 1.1 0.27 11.22 45.6 22214 89.6 -0.4 90

Cn0,

CnO, tb rad CnO, ff

0.0173 0.0061 0.0112

0.0191 0.0080 O.Ol l 1

-0.0154 -0.0061 -0.0093

0.0132 0.0053 0.0079

-0.0182 -0.0105 -0.0076

-0.0309 -0.0192 -0.0117

0.0291 0.0222 0.0069

-0.0003 0.0010 -0.0013

-0.0260 -0.0206 -0.0053

RESULTS AND DISCUSSION

This section first discusses the forebody yawing moments to show the overall effect of the forebody strakes. The forebody and strake pressure distributions are then discussed to explain causes

for some forebody yawing moment results. Flow visualization is used to support the interpretation

of the pressure distributions. Finally, pressure distributions from a full-scale wind tunnel model are

presented for comparison.

Forebody Yawing Moments

Figure 7 shows the yawing moment at 0 ° sideslip from the 30- by 60-Ft 22 and 80- by 120-Ft 23

Wind Tunnels for the complete airplane are presented as a function of angle of attack for left and

.o6 ..... _lt .... '_..... ; .....

! , E L ----i:---::, ;;--

°no'. hiiill

Cno, 0 .... _ .......

- .02

-._ ..... !..... i..... !..... i- -_..... i.....

i ', ', , , ,8s, L/R" 0°/90°

-- .08

0 10 20 30 40 50 60 70 80 90

Ct,deg

960619

---o-- Cno, 30 x 60, ref. 22

Cn0, 80 x 120, re/. 23

Cn0, fb, flight

---O-- Cn0 ' fb, 80 x 120

Figure 7. Comparison of forebody yawing moments from flight and wind tunnel and total aircraft

yawing moments from wind tunnel.

12
right strake deflections, 5s, L/R = 90°/0°' 0°/0% and 0°/90 °. Figure 7 also shows yawing moments

for just the forebody alone (F.S. 60 to F.S. 190) from flight a:_ the 80- by 120-Ft Wind Tunnel.

Note that when the left strake is deflected, 90°/0 °, the yawi_ aoment is positive or to the right

and the converse is true when the right strake is deflected, ,90 °. This same figure shows the

strong effectiveness of the actuated forebody strakes at high angle of attack, especially when compared to the rudder. The maximum forebody yawing moment from flight is approximately one-half

of the total maximum aircraft yawing moment from the wind tunnels. At ct = 50 ° the flight forebody yawing moment is approximately 80 percent of the wind-tunnel value. The peak forebody

yawing moment from flight was at tx = 55 ° and at ct = 59 ° (maximum test o0 for the wind tunnel.

For the whole airplane, the maximum yawing rate was at ct = 50 °. At _ = 59 ° for the wind tunnel

and t_ = 65 ° for flight, the forebody and total aircraft yawing moments were approximately equivalent. The differences between the flight and wind-tunnel forebody yawing moments will be

discussed further in the Pressure Distributions subsection.

During development in wind-tunnel tests, 15 deflecting one strake at a time at high angles of

attack could result in a small but undesirable control reversal at small strake deflections. To

overcome this undesirable characteristic for closed-loop control, a solution was developed that deploys the strakes symmetrically as angle of attack increases to ct = 30 ° and greater. When a yawing

moment is desired under these conditions, the strakes are deflected differentially about a symmetric strake deployment (fig. 8). For example, using the 20 ° symmetric strake schedule, (fig. 8(c)),

(_sd = 0° would correspond to 5 s L/R - 20°/20°, _)s,d = 10° would correspond to ks, L/R = 15°/25°,

_)s',d 20 ° would correspond to _s,/JR = 10°/30° and so on. For all angles of attack, however, the

maximum yaw control deflection would always consist of one strake fully deployed (90 °) and the

other strake fully retracted (0°).

90

60

30

0

- 30

___i ('e) 0 ° ,y'mmetric i / q

slreke depl°ymenl /_(/

Left etrake

- - - Right etrake

i (1=)10 ° symmetric i / "1

$treke deployment , / |

..... ..... ...... i

! ! i .... -J

i _ i i t

I i i i , ,

__ (c) 20 ° symmetric i //]

..... i ..... st,rake ..... deployment ..... ......... _/ J

/

, _ , .

i i i i i

i i i , i

i i i i i

_i (tl) 30 ° eymmetri'c i / /

etrake deployment ; /

..... ....i.... .....

_ i J J i i

, / i i i t

i .*" i * i b

i i t i i

- 90 - 60 - 30 0 30 60 90

_s, d, deg

960620

Figure 8. Symmetric strake deployment schedules.

13
Figure9 showstheforebodyyawingmomentsfrom flight asa functionof differential strake

deflection,8s,d, (right strake deflection minus the left strake deflection). At ot = 30 °, these data

show the variation of forebody yawing moment with differential strake deflection for the 0 ° and

10 ° symmetric strake deployments (fig. 9(a)). For the 0 ° symmetric strake deployment, the left and

right strakes are at 0 ° (closed) when no yawing moment is desired. For the 10 ° symmetric strake

deployment, the left and right strakes are deployed symmetrically to 10° when no yawing moment

is desired. Very little difference is noted between the 0 ° and 10 ° symmetric strake deployments.

At ot = 40 °, data for the 0 ° symmetric strake deployment show a control reversal at 8s, d = +10 °

(fig. 9(b)). In this case as the strakes begin to open, the yawing moment is opposite to the desired

direction. Data for the 10 ° symmetric strake deployment are better but still show a slight reversal.

In this case at o_ = 40 °, the 20 ° symmetric strake deployment eliminates the control reversal and

results in a nearly linear variation of forebody yawing moment.

At o_= 50 °, a similar trend is seen (fig. 9(c)). The 0 ° symmetric strake deployment results in a

large control reversal. As at ot = 40 °, the 20 ° symmetric deployment eliminates the control reversal

and results in a nearly linear variation of forebody yawing moment for differential strake deflections. The 30 ° symmetric strake deployment also eliminates the control reversal, but the curve does

not have the desired linear slope.

At ot = 60 °, the control reversal in yawing moment for the 0 ° symmetric strake deployment becomes larger and asymmetric (fig. 9(d)). The forebody yawing moment at 8s, d = 20 ° (0o/20 °) is

approximately double that at 8s, d = -20 ° (20°/0°). The 30 ° symmetric strake deployment is an

improvement but is not as linear as desired.

For comparison, figure 10 shows the forebody yawing moments at ct = 50 ° for the 0 ° and 20 °

symmetric strake deployments from the 80- by 120-Ft Wind Tunnel. As compared with the

flight forebody yawing moments, the yawing moment reversal for the 0 ° symmetric strake deployment from the wind tunnel is only about one-half the flight value (fig. 9(c)). The 20 ° symmetric

strake deployment is much improved but not quite as linear as the flight case. The maximum

forebody yawing rates from the wind tunnel at 8s, d = _+90° is approximately 25 percent higher than

the flight values. The explanation for these differences will be given later under the Pressure

Distributions subsection.

In figure 11, the flight forebody yawing moments at tx = 50 ° are broken down into two parts

for the 0 ° and 20 ° symmetric strake deployments: the ANSER radome including strakes (F.S. = 60

to 124.5) and the forward fuselage between the radome and the apex of the wing leading-edge

extension (F.S. = 124.5 to 190). For the 0 ° symmetric strake deployment, the majority of the yawing moment results from the side force on the forward fuselage, including the region of yawing

moment reversal, 18s,d I < 30 ° (fig. 1 l(a)). Using the 20 ° symmetric strake deployment, the yawing

moment reversal is no longer present for the radome and the forward fuselage (fig. 1 l(b)). Again,

as for the 0 ° symmetric strake schedule, the strakes generate little yawing moment at 18s,dl < 30 ° as

shown by the radome data. However, the vortices the strakes generate create a significant yawing

moment further aft on the forward fuselage. Approximately 80 percent of the forebody yawing moment for 18s,dI < 30 ° at Ot = 50 ° is caused by the side forces on the forward fuselage. At _s,d = +90°

the forward fuselage accounts for 55-60 percent of the forebody yawing moment.

14
Cno, tb

Strake deployment

0 ° Symmetric

-- -D-- - 10 ° Symmetric .04 .....

.o31.......i:.......i:.......i:......i:......._:......t

•02 ........ :....... _....... ',....... J,....... _...... "1

• ...... i....... i... t

o........ _....... :- -' ....... ! ......

-.o_......_:...... "....... :_......_......'

-°_l......_-_ ...... i......1

_ 04 _ _ I ', I i I

• - 90 - 60 - 30 0 30 50 90

_s, d' deg

960621

Cno, fb

.04

.03

.02

.01

0

- .01

- .02

- .03

- .04 90

Strake deployment

----o-- 0° Symmetric

- - o--- 10 ° Symmetric

-- _ - 20 ° Symmetric

....... _....... ;....... '_......i.......i.......

i l

i p i i ,

_-, .............. :....... _... _......

p i

t i

.............._.......,...............,L......

i

, ; , ', I

- 60 - 30 0 30 60 90

5s, d, deg

960622

(a) cz = 30 °. (b) c_ = 40 °.

Cno, fb

.04

.03J

.02

.01

0

- .01

--,0_11

- .03

- .04 90

Strake deployment

----o-- 0 ° Symmetric

--0-- 20 ° Symmetric

--z_-- 30 ° Symmetric

i i i ! i

_,.__.__.......:......._.......,L....... _i_-_.+-s,--_-_ l %'e.. , i i

....... i.......

: ' re.' ,, L

,, ,, ,, _.., .. ',

........ ,,....... _....... ,,....... _......... 2

........_---I ......_"_

I I I I I

- 60 - 30 0 30 60 9O

_e, d, deg

960623

Cno, fb

.04

.03

.02

.01

0

- .01

- .02

- .03

Strake deployment

0 ° Symmetric

-- .-D-- - 30 ° Symmetric

_-i..... .-.-_ ....... i...... _....... !.......

....::_: -_:":_ir_.......;......

....... i ...... _,...... • ...... ,_....... !!!!ili ii: 

90 - 60 - 30 30 60 90

_s, d' deg

960624

(c) (x = 50 °. (d) (x = 60 °.

Figure 9. Forebody yawing moments as a function of differential strake position, 13= 0°.

Strake deployment

0° Symmetric

---o-- 20 ° Symmetric .05 .....

___ ', ,, ,, :

.04 .... _....... _...... i....... !.......

.o_t'-..... ?_---f ....... ?...... !....... ,_.......

.o,_......_,,_-:_--_ oL!' I' '

.o_ ......_...... _--._--,......,....... ,.....

', : _.' ,, :

Cno, fb 0 ......-_.......T...... _.......,_.......

- .01 ' ' ' ----! .......

-.°3.......__--_-- ....

- .04 ......-_......._,_ ......;...... ,_-_._

-.0 , I _ '_

_-_o -60 -30 0 30' 60' 90

_s, deg

_062_

Figure 10. 80- by 120-Ft Wind Tunnel forebody yawing moments as a function of differential

strake position, [3 = 0 °, 0 ° and 20 ° symmetric strake deployments•

15
Cno

Forebody

-- -O-- - Radome

- - -zl- - Forward fuselage

°4_,._ T ! ! ! ! /

•°3_-_t......_......._......_.......i......1

•o,_:-L._--i.......i.....i......_......1 .oi -_....... _- -_....... _......

_.o,I_.....________:.:%.__.____, --o't...... .......

::.r-o....-!:.... ;:-....-!-.... -j:.....

_s, d' deg

960626

(a) 0 ° symmetric strake deployment.

Cno

Forebody

-- -O-- - Radome

- - -_1- - Forward fuselage

.04 : _ : :

......i.......................

.02, ,:::_- - - -,'-:'-._,: - _....... ',....... J....... F.......

I -. _-:.'"-Z_ "'"-Z_ _'_ :' ', '. ',

.01 :--- --" D-:.: __ ...... I ....... ,".......

o.......: i-----_-o-- "_-_ - ._,_+-..._ -._, _ - -_.......

-.of ...... _....... !...... :------'_---:----o-.-....

-.o3....... _ .... !...... !:',-._

- 04 " ; _ " ;

• - 90 - 60 - 30 0 30 60 90

(Ss, d' deg

960627

(b) 20 ° symmetric strake deployment.

Figure 11. Forebody, radome, and forward fuselage yawing moments as a function of differential

strake position, o_= 50 ° and 13= 0 °.

Pressure Distributions

Figure 12 shows the forebody and deployed strake pressure distributions for o_= 50 ° for the 0 °

and 20 ° symmetric strake deployments. This figure shows the aerodynamic mechanisms responsible for the yawing moments generated by the strakes. For the 0 ° symmetric strake deployment,

large yawing moment reversals were present. For the 20 ° symmetric strake deployment, the yawing

moments were linear. Separate plots are shown for the forebody and the right and left strakes. Note

that the pressure distributions are offset by Cp = 1.0 for each consecutive fuselage station for clarity.

0 ° Symmetric Strake Deployment

Figure 12(a) shows the forebody pressure distributions as a function of the forebody crosssection circumferential angle, 0. The scale for the circumferential angle is reversed so that the pressure distributions can be viewed in the pilot's perspective. The 0 = 0 ° and 360 ° are on the lower

centerline; 0 = 180 ° is on the top centerline; 0 = 90 ° is on the right side of the fuselage; and 0 =

270 ° is on the left. This convention was established in previous papers. 17,20 The symbols and solid

curve indicate the pressure distribution from the HARV ANSER radome. The footprints of the forebody vortices can be seen by the suction peaks in the pressure distributions at 0 -- 160 ° and 200 °

for F.S. 85, F.S. 107, F.S. 142, and F.S. 184. The large suction peaks at 0 -- 90 ° and 270 ° for F.S. 70,

F.S. 85, and F.S. 107 are caused by the acceleration of the flow around the fuselage that is nearly

circular in cross-section. The small suction peak for F.S. 142 at 0 -- 110 ° and 250 ° is caused by a

small antenna cover just forward of F.S. 142. The dashed curve indicates pressure distribution data

obtained from the original HARV radome. 17 The most obvious difference between ANSER radome results and the original HARV radome results is the deficit in the suction peak at 0 -- 80 °

to 110 ° and 250 ° to 280 °, starting at F.S. 70 and progressing aft to F.S. 107. The original HARV

radome suction peak is much smoother and fuller. This deficit region is below the strake. The joint

at 0 = 280 ° (where the white and black paint meet below the strake) (fig. 3) could cause boundarylayer transition. Previously, however, boundary-layer trips very near this location on the original

16
Left Right

-a !

%1 k

,, i

360 300 240 180 120 60 0

e, deg

960628

- - - F.S. 184

F.S. 142

F.S. 107

F.S. 85

F.S. 70

Radome

HARV ANSER

------ HARV original, ref. 17

Note: Cp offset

(a) 6s, L/R = 0°/0° with comparison to previous flight results.

Figure 12. Forebody and strake pressure distributions with the ANSER radome at o_= 50 ° for the

0 ° symmetric strake deployment, pilot's view.

HARV radome 24 did not show a deficit in this region but did show the enhanced vortex footprints

at F.S. 85. The radome was checked for roundness in this region at F.S. 70 and F.S. 85, and the

local curvature appeared to be satisfactory. The strakes were sealed from the interior of the radome

and from top to bottom. The gap between the retracted strake and fuselage was on the order of 0.1

to 0.2 in. Large cutouts existed as shown previously in figure 6, some 1- by 1-in., and some 1.5- by

2-in. This deficit in the suction peak will be discussed further when the ANSER radome pressure

distributions are correlated with the 80- by 120-Ft Wind-Tunnel model pressure distributions.

Figure 12(b) shows the pressure distributions at ct = 50 ° for the right strake extended to 10 °. At

this strake deflection (fig. 9(c)), a strong yawing moment reversal with the yawing moment positive or to the right occurred. At this angle of attack, the footprints from forebody vortices that can

be seen at 0 -- 160 ° and 200 ° become asymmetric at F.S. 142 and F.S. 184. The pressure distribution on the right side, especially between 0 = 60 ° to 120 °, tend to have lower pressures than on the

left side at the same corresponding angles. This circumferential range of angles, 0 -- 60 ° to 120 ° on

the right side and the corresponding angles (240 ° to 300 °) on the left side are important because

they represent approximately 87 percent (sin 60 °) of the forebody side area. At this strake deflection angle, the pressures on the outboard and inboard strake surfaces are below ambient (Cp < O)

with the outboard surface pressures less than the inboard surface pressures. Pressures on the forebody and the strake result in a yawing moment to the right, which was not the desired direction

(figs. 9(c) and 11 (a)).

With the right strake deflected 20 °, the pressure distributions become more asymmetric with

large suction peaks at 0 = 160 ° for F.S. 85 and F.S. 107 caused by the right forebody/strake vortex

(fig. 12(c)). The right vortex footprints at F.S. 142 and F.S. 184 are diminished because the right

17
Left Right

-2

-1

0

0

Cp

0

O'

0

1

360

I I 1 I I I

i

300 240 180 120 60 0

e, dag

F.S. 184

F.S. 142

>F.S. 107

F.S. 85

F.S. 70

(b) _)s,LIR = 0°/10°"

-2

-1

0

Note: Cp offset

Open symbolsoutboard surface

Closed symbols -

Inboard surface

T /

0 .5 1.0

y/b, right strske

96O629

Cp

-2

-1

0 d

O I

Left Right

F.S. 184

1 F I I I

i , i i

300 240 180 120 60 0

e, deg

F.S. 142

>F.S. 107

F.S. 85

F.S. 70

-2

I 0

360 -.5

(c) ks, L/R = 0°/20°"

Figure 12. Continued.

Note: Cp offset

Open symbols -

outboard surface

Closed symbols -

Inboard surface

! I

F.S. 107 ',

- 1 ...... _--

F.S. 85

i i ,

0 .5 1.o

y/b, right strake

96O63O

vortex is beginning to lift from the forebody surface. Again, the pressures on the right forebody

tend to be less than those on the left forebody, particularly at F.S. 142 and F.S. 184. Pressures on

the strake are less than ambient, and the outboard surface pressures are generally less than the inboard surface pressures. The suction peak on the inboard surface at y/b = 0.85 at F.S. 85 is approximately the same magnitude (Cp = -1.2) as the one on the fight forebody and is caused by the

18
proximity to theright forebody/strakevortex.Again,thecombinationof theforebodyandstrake

pressuresresultsin a yawingmomentto theright,notthedesireddirection(figs. 9(c) and 1l(a)).

In figure 12(d),theright strakeis deflected30°. This deflectionresultsin a strakethat is in a

nearlyverticalorientationwith thefuselage.A largesuctionpeakoccursat0 = 160° at F.S.85and

iscausedby therightforebody/strakevortex.Thissuctionpeakdiminisheswhenmovingaft onthe

fuselageandis causedbythevortexlifting fromthesurface.Theforebodyvortexontheleft,noted

by the suctionpeaksat 0 = 200°, developsand lifts from the surfacelaterthanthe right forebody/strakevortex.At F.S. 107,F.S. 142,andF.S. 184,thepressuresontheleft sideat 0 ---240°

to 300° aregenerallylessthanthecorrespondingpressuresontheright side.Onthestrake,thepressuresareall still lessthanambient. Thesuctionpeakfrom thestrakevortexontheinboardsurface

aty/b = 0.85 has become greater in magnitude (Cp = -1.95) than the right forebody suction peak

(Cp = -1.5). At F.S. 107, the inboard and outboardstrake surface pressures are approximately the

same. The net result of the forebody and strake pressure integration is a small yawing moment to

the left, the desired direction (figs. 9(c) and 11 (a)).

When the right strake extends to 60 ° (fig. 12(e)), the forebody pressures on the left side at 0 =

240 ° to 300 ° are noticeably less than the corresponding pressures on the right side for all pressure

orifice stations resulting in a yawing moment to the left. At this strake deflection, the strake is approximately 30 ° out from vertical and retards the flow on the right forebody below the strake.

Greater than ambient pressures can be seen on the outboard strake surface at the intersection with

the forebody, y/b = 0.2. The combination of the right and left vortices accelerate the flow around

the left forebody as compared to the lower strake deflection angles. The right strake suction peak

is a maximum at 0 = 160% F.S. 85, and quickly diminishes moving aft as the right forebody/strake

Left Right

'/! !', : : i /

- 1 ........... J ..... ' .......... F.S. 184

t p , t

F.S. 142

Cp ,

F.S. 107

F.S. 85

F.S. 70

360 300 240 180 120 60 0

O, deg

(d) 8s, L/R = 0°/30°"

Figure 12. Continued.

-2

-1

0

_.5

Nots: Cp offset

Open symbols -

outboard surface

Closed symbols -

Inboard surface

F.S. 107 : I

: t

!iiii 

F.S. 70 i

i

I I

0 .5 1.0

y/o, right strake

960631

19
Left Right

-2 I I 1 I I

o ...... :...... _...... ' .....

Cp

360 300 240 180 120 60

e, deg

F.S. 184

F.S. 142

F.S. 107

F.S. 85

F.S. 70

-1

0

0 -.5

(e) 6s, L/R = 0°/60°.

Figure 12. Continued.

Note: Cp offset

Open symbolsoutboard surface

Closed symbols -

Inboard surface

1 I

...... I '-At

I 1

0 .5 1.0

y/b, fight strake

960632

vortex lifts from the surface of the radome. The left forebody vortex grows in strength, stays close

to the surface, and causes the flow to accelerate around the left side of the forebody. This vortex

action results in lower pressures on the left side, hence a yawing moment to the left. The left vortex

suction peak is a maximum at 0 = 200 ° at F.S. 107 and moves inboard progressing aft, so this peak

is at 0 = 180 ° at F.S. 184. On the strake, pressures on the outboard surface are greater than those

on the inboard surface, resulting in a yawing moment to the left. The suction peak (Cp = -2.0) on

the outboard strake surface at F.S. 70, y/b = 0.85, is almost the same as for the right vortex suction

peak (Cp = -2.1) at F.S. 85, suggesting that the forebody/strake vortex passes close to that orifice.

The sum of large pressure differences on the forebody and strake results in a large yawing moment

to the left, (figs. 9(c) and 1 l(a)).

When the right strake is fully extended to 90 ° the strake retards the flow below the strake even

more than for 8s = 60 ° (fig. 12(f)). The suction peaks at F.S. 70, F.S. 85, and F.S. 107 and 0 = 60 °

to 120 ° are significantly lower than for _s = 60° (fig" 12(e)). Pressures on the left side of the fuselage are approximately the same as for 6s = 60°, (fig' 12(e)). On the strakes at F.S. 85, the difference

between the inboard and outboard surface pressures is significantly larger than for _5s = 60 °. Like

the _is = 60 ° case, the pressure distributions indicate that the right vortex lifts from the surface after

F.S. 85 while the left vortex stays close to the surface and accelerates the flow around the left side.

The suction peak (Cp = -1.9) on the inboard strake surface at F.S. 85, y/b = 0.65, is almost the same

as on the fuselage at the same fuselage station. As a result of the lower suction peaks on the forebody and the outboard and inboard surface pressures on the strake at F.S. 85, the yawing moment

to the left was even larger than for _5s = 60 °, (figs. 9(c) and 11 (a)).

20
-2

-1

0

0

Cp

0

O'

0

1

36O

Left Right

I I I I I

Note: Cp offset

F.S. 184 Open symbols -

outboard surface

Closed symbols -

inboard surface

F.S. 142

o

1

300 240 180 120 60 -.5 0 .5 I.::

e, deg y/b, right atrake

960633

(f) ks, L/R = 0°/90°"

Figure 12. Concluded.

Figure 13 provides supporting evidence of this interpretation, as shown in the wingtip photograph ct = 50 °, [5 ---0 °, and ks, LIR - 0°/90°" The right forebody/strake vortex can be seen very high

off the forebody, departing the surface at the strake. A weaker left forebody vortex can be seen

lower and near the surface of the forebody.

Figure 13. Forebody/strake vortex flow visualization for ot = 50 ° and _is, LIR = 0°/90°"

21
20° SymmetricStrakeDeployment

Symmetricstrakedeploymentswereusedathighanglesof attackto overcometheproblemof

yawcontrolreversalat smallstrakedeflections.At o_= 50°, the20° symmetricstrakedeployment

yieldedthebestcontrollinearity characteristics(figs. 9(c)and1l(b)).

Figure14(a)showsthepressuresdistributionsfor _Ss, L/R = 20°/20°" Compared to the 0 ° strake

deflection case (fig. 12(a)), the suction peaks from the forebody/strakes are slightly more pronounced for F.S. 70, F.S. 85, and F.S. 107. A slight asymmetry exists in the pressure distributions,

especially at F.S. 184, with slightly lower pressures on the right side, resulting in a slight yawing

moment to the right. No significant differences are seen in the strake pressures from left to right.

For 5s, L/R = 15°/25° (fig' 14(b)), the right forebody/strake vortex suction peak is greater at

F.S. 85 than the left but lifts off from the surface sooner, as can be seen by the vortex suction peak

pressures at F.S. 142. Except for the suction peaks from the forebody/strake, the forebody pressure

distributions from F.S. 70 to F.S. 142 are generally symmetric left to right. At F.S. 184, the pressures on the left side are generally less than those on the right side, resulting in a small yawing moment to the left as desired. The pressures on the outboard side of the strake are generally less than

those on the inboard side with the largest difference on the left strake. Again, this difference results

in a small net yawing moment to the left.

Figure 14(c) shows a similar trend for 5s,/_JR = 10°/30°" With the exception of the suction peaks

from the forebody/strake vortices, the pressure distributions at F.S. 70 and F.S. 85 are nearly symmetric. As the right forebody/strake vortex suction peak diminishes and the vortex lifts from the

surface, the left forebody/strake vortex suction peak increases and the vortex accelerates the flow

around the left side of the fuselage resulting in lower pressures on the left side at F.S. 107 to

Left Right

-1 _._ , .... : .... _........ F.S. 184

1 F.S. 107

o - ' .od

"°1 F.S. 86

..... i

F.S. 70

ol ; ', -

1.0 .5 0 - .5 360 300 240 180 120 60 0

y/b, left strake

0, deg.

(a) _is,/_JR = 20°/20°"

Note: Cp offset

Open symbols -

outboard surface

Closed symbols -

inboard surface

2:

i : :

! F.S._07 :

i i

I I

/ :l :J I

o

- .5 0 .5 1.0

y/o, right strske

960635

Figure 14. Forebody and strake pressure distributions with the ANSER radome at a = 50 ° for the

20 ° symmetric strake deployment, pilots view.

22
Cp

_2 ! !

• _ Fis10,

0 .... _ P .....

_ F.S. 85

y/b, left strake

-2

_1L0 bo-_

0 _-

l

0_---

1/

360

Left Right

I t I I I

-_---

300 240 180 120 60

0, deg

(b) 5s, L/R = 15°/25°"

F.S. 184

F.S. 142

F.S. 107

F.S. 85

F.S. 70

Note: Cp offset

Open symbols -

outboard surface

Closed symbols -

inboard surface

'| I I

y/b, right atrake

960636

Cp

Left Right

- I I I i 1 __tF.S"

0 --- :.... "..... " ....

! !, I . F.S. 142

/ : FJS. 107 I v

/ ad¢_._,,_ / o - F.S. 107

: ',F.S. 85 0

to s 01 li -- 360 300 240 180 120 60 0

y/b, left strake 0, deg

(C) _)s,LIR = 10°/30°"

Figure 14. Continued.

184

Note: Cp offset

Open symbols -

outboard surface

Closed symbols -

Inboard surface

-2

I

* )

i p

0 ..... " ....

F.S. 85

o..... i-_-

,F.S. 7O ',

' :

i i

0.5 0 .5 1.0

y/b, right atrake

960637

F.S. 184. Note the strength of the left vortex footprint at 0 = 200 ° at F.S. 142 and the left and right

side pressures as compared to the same for the 0 ° symmetric strake deployment, Gs, L/R = 0°/20°

case (fig. 12(c)). For the same differential strake deflection of 20 °, the left forebody vortex had little beneficial effect. In that case, the dominant right forebody/strake vortex caused the flow to accelerate around the right side of the forebody for all stations aft of F.S. 85, increasing those suction

pressures and resulting in a yawing moment to the right. On the strakes for Gs, LJR = 10°/30° the

suction peak on the inboard right strake surface at F.S. 85, y/b = 0.85, is caused by the proximity

of the forebody/strake vortex as was seen for _s, LIR = 0°/30° in figure 12(d). Again the outboard

strake surface pressures are generally less than the inboard surface pressures with the largest

23
Left Right

-2 I I I I 1

-1 ......, , '

-1 '°°,.__ /I--- - -- - !

',

_

F•S•',107

.... t o

%oL .:to o 6._ :.....:i::s.-7oI o

0 / / l 1

1.0 .5 0 - .5 360 300 240 180 120 60 0

y/b, left strake e, deg

(d) ks, UR = 5°/35°"

Figure 14. Concluded.

F.S. 184

F.S. 142

F.S. 107

F.S. 85

F.S. 70

Note: Cp offset

Open symbols -

outboard surface

Closed symbolsInboard surface

-21 I,, I

-1 F'S'1107 i

• • Ij ,

o _.s__o:....:.....

i i

0.5 0 .5 1.0

y/b, right strake

960638

difference on the left side• Both the forebody and strakes pressure integrations result in yawing

moments to the left, (figs. 9(c) and 11 (b)).

This same trend continues on for _Ss,L/R = 5°/35° (fig" 14(d)). The result of the small left strake

deflection is an increase in the strength of the left forebody/strake vortex. For example, the vortex

footprint at F.S. 142, 0 = 200 °, for this strake deflection is much larger than the same differential

strake deflection for the 0 ° symmetric strake deployment, _is, L/R = 0°/30° (fig" 12(d)). As a result,

the stronger left forebody/strake vortex generates lower pressures on the left side and a much larger

yawing moment to the left (fig. 9(c)).

As supporting evidence of this analysis, for o_ = 47 °, 13= 0.2 °, and ks, L/R = 5°/35°, nearly the

same conditions as figure 14(d), figure 15(a) shows the left and right forebody/strake vortices in

the wingtip photograph. The right forebody/strake vortex can be seen higher and lifting off the surface sooner. The right vortex stays much closer to the surface and provides the suction needed to

accelerate the flow around the left side of the forward fuselage aft of the radome. For comparison,

smoke flow visualization is shown in figure 15(b) for the 0 ° symmetric strake deployment case at

_is,/JR = 0°/27°, o_ = 50 °, and 13= -1.4 °. For nearly the same differential strake position, only the

right forebody/strake vortex close to the surface can be seen. The left vortex does not appear to

have a core that can be identified.

Comparison with Wind Tunnel Results

A similar radome with conformal strakes was tested on an F-18 airplane in the NASA Ames

Research Center, Moffett Field, California, 80- by 120-Ft Wind Tunnel. 23 This full-scale model

had pressure orifices at the same fuselage stations as the HARV. The conformal strakes on this

radome, however, did not have the cutouts for the bulkheads that were on the HARV, (fig. 6).

24
(a)_Ss,/.JR= 5°/35°"

(b) 8s,L/R = 0°/27°"

Figure 15. Forebody/strake vortex flow visualization for o_ = 50 ° and _s,d -- 30°"

Pressure distributions from flight and wind tunnel at tx = 50 ° and ks, L/R = 0°/90° are compared

in figure 16(a). The most obvious difference between flight and the wind-tunnel results is the deficit in the suction peak on the left side (0 -- 240 ° to 280 °) starting at F.S. 70 and progressing aft for

the flight data. The wind-tunnel data suction peak is much smoother and fuller, similar to previous

HARV radome results from flight (fig. 12(a)). 17 The major difference between the wind-tunnel

model radome and the ANSER radome was that the large cutouts for the flight ANSER radome

were not present on the wind-tunnel model. These cutouts were sealed from the interior of the

25
radome but not sealed between other cutouts. Axial flow from a region of higher pressure to a region of lower pressure could have existed, thereby affecting the suction peaks at F,S. 70, F.S. 85,

and F.S. 107.

On the right side of the fuselage, the pressure distributions from flight and wind tunnel showed

good agreement at F.S. 70, F.S. 85, and F.S. 107. At F.S. 142 and F.S. 184, however, the suction

pressures on the right side from flight are higher than those from the wind tunnel. The net result is

that the wind tunnel has higher forebody yawing moment coefficients for this test condition (fig. 7).

In figure 16(b), flight and wind-tunnel results are compared for ct = 50 ° and _Ss,L/R = 0°/20°, a

condition for which a large yawing moment reversal was noted in flight (fig. 9(c)). Again, at

F.S. 70, F.S. 85, and F.S. 107 on the left side, 0 -- 240 ° to 280 °, the deficit in the flight pressure

distributions can be seen compared to the wind-tunnel data. At F.S. 142 and F.S. 184, the

wind-tunnel pressures on the left in the region of 0 = 240 ° to 300 ° are approximately the same as

the corresponding pressures on the right; whereas, those from flight are lower on the right. The net

result was a slightly favorable yawing moment to the left for the wind-tunnel data and an unfavorable yawing moment to the right for the flight data, (figs. 9(c) and 10).

Figure 16(c) shows the comparison between flight and wind tunnel data at _t = 50 ° and 8s, L/R =

10o/30 °, a case where both flight and wind-tunnel data have favorable yawing moments. The flight

and wind-tunnel data compare well with the exception of F.S. 142 and F.S. 184 on the right side

where the flight pressures are slightly lower. At F.S. 107 on the left strake, the wind-tunnel pressures

were slightly lower. No data were obtained at F.S. 85 on the right strake for the wind tunnel.

-2

-1

0

Cp

Left Right

I I

-1

I

F.S. 184

F.S. 142

F.S. 107

F.S. 85

F.S. 70

6O 0

1 1

360 300 240 180 120 -.5

0, deg

Note: Cp offset

Solid curve, symbols

wind tunnel, ref. 23

Dashed curve, flight

Open symbols -

outboard surface

Closed symbols -

Inboard surface

iz,

...... i

"'i

,

F.S., 70

J

i 0

t

I I

0 .5 1.0

y/b, right strake

960641

(a) 8s, £/R = 0°/90°"

Figure 16. Comparison of flight and wind tunnel pressure distributions, ot = 50 °.

26
Left Right Note: Cp offset

-2 I I

: ,_i I: I Solid wind curve, tunnel, symbols ref. 23

, , ', ._. .'. L_ t_ , , __ _z.:, J F.S. 184 Dashed curve, flight

- 1 _ --;- - - ?_..-._ _;;; Open symbols -

- - - -t -,-_ ........... , .... i" "- ./_ -

, , _ • , outboard surface

.... "' Closed symbols -

0 ..... ,....... , ..... _ 7 _ .... inboard surface 2

I ' F.S. 142 ,, ,,

c,o -'F....

o .............. i .s.,o I

360 300 240 180 120 60 0 -.5 0 .5 1.0

O, deg y/b, right etrake

960642

(b) _is, UR = 0°/20°"

Left Right

i i i _ i

_ i n a

, _ , , ,,.... F.S.

Solid curve, symbols ' " ....

wind tunnel, ref. 23 -_* _r_- _;Y ' _' _.;

Dashed curve, flight ...... -_ - "_- - _ _....

/ : F:S.1071 0 i'_

-1_ .... -_ <_/_____. .... ..__;_ ,i_ i__i_ __

Cp 0 8 0 ....

n _o_.._ ..... _ .... -.1 '_ __

" F.S. 70 I 0 -

i i I I l I i

01.0 .5 0 -.5 360 300 240 180 120 60

y/b, left strake 0, dsg

(c) 6s, UR = 10°130°"

Figure 16. Concluded.

184

F.S. 142

F.S. 107

F.S. 85

;.S. 70

0

-2

-1

Note: Cp offset

Open symbols -

outboard surface

Closed symbolsinboard surface

I I

 .s.11o7i

..... : ..... _.- -

t_ _ a

t

i" :i'::.

.....

 's'7°i /

2 i i /

- .5 0 .5 1.0

y/b, right strake

960643

27
CONCLUDING REMARKS

In-flight pressure distributions have been reported at an angle of attack (_) of 50 ° at five fuselage stations on the F-18 High-Alpha Research Vehicle (HARV) forebody with actuated forebody

strakes. Forebody yawing moments integrated from the forebody pressures were presented at angles

of attack of 20°-65 °. The results have been correlated with in-flight off-surface flow visualization

and wind-tunnel data from the NASA Ames Research Center 80-Ft by 120-Ft Wind Tunnel and the

NASA Langley Research Center 30-Ft by 60-Ft Wind Tunnel.

Forebody strakes can generate large yawing moments at high angles of attack. When one strake

is deployed fully open, the maximum forebody yawing moment from flight is approximately onehalf of the maximum aircraft yawing moment from the wind tunnels. At ct = 50 °, the flight forebody

yawing moment is only approximately 80 percent of the wind-tunnel forebody yawing moment. The

peak forebody yawing moment from flight was at t_ = 55 ° and at ot = 59 ° (maximum test ct) for the

wind tunnel. Peak forebody yawing moment for the whole airplane from the wind tunnel was at 50 °.

At ct = 59 ° for the wind tunnel and 65 ° for flight, the forebody and total aircraft yawing moments

from the wind tunnel were approximately equivalent.

At angles of attack of 40 ° and greater, deflecting one strake at a time resulted in a forebody yawing moment control reversal for small strake deflection angles. At tx = 40 ° and 50 °, deflecting the

strakes differentially about a 20 ° symmetric strake deployment eliminated the control reversal and

produced a nearly linear variation of forebody yawing moment with differential strake deflection.

When the forebody yawing moments are broken down into radome and forward fuselage yawing moments at ot = 50 °, the major forebody yawing moment is not generated at the region where

the strakes are located but rather at the forebody region behind the strakes. At I_)s,dl < 30 °, the

forward fuselage accounts for practically all of the forebody yawing moment.

With the right strake deflections of 10° and 20 ° and left strake retracted at ot = 50 °, the pressure

distributions show that the right forebody/strake vortex was close to the surface. This configuration

caused the flow to accelerate around the forward fuselage behind the strake, which resulted in lower

pressures on the fight side and a forebody yawing moment to the fight. For fight strake deflections

of 60 °, the fight forebody/strake vortex lifted from the surface of the radome and the left forebody

vortex grew in strength and stayed close to the surface. This vortex action caused the flow to accelerate around the left side of the forebody, resulting in lower pressures on the left side and hence a

yawing moment to the left.

At o_= 50 ° using the 20 ° symmetric strake deployment and right strake deflections between 25 °

and 35 °, the deflected left strake increased the strength of the left forebody/strake vortex. This vortex was also closer to the surface than the right forebody/strake vortex and accelerated the flow

around the left side, resulting in lower pressures on the left side for all conditions and a forebody

yawing moment to the left with no forebody yawing moment control reversal.

A deficit in the suction peaks of the flight pressure distributions was noted just below the undeflected strake as compared to the original HARV radome data and to the wind-tunnel radome data.

This deficit was possibly caused by the cutouts in the strakes required for the fuselage bulkheads on

the flight hardware.

28
REFERENCES

1Chambers,JosephR., High-Angle-of-Attack Technology: Progress and Challenges, NASACP-3149, vol. 1, pt. 2, 1990, pp. 1-22.

2Bowers, Albion H. and Pahle, Joseph, "The High Alpha Research Vehicle," High Angle of

Attack Conference, NASA Langley Research Center, Sept. 1996.

3Chambers, Joseph R., Gilbert, William P., and Nguyen, Luat T., ed, High-Angle-of-Attack

Technology, vol. 1, pt. 1-3, NASA-CP-3149, vol. 1, 1990.

4Matheny, Neil W., comp., High-Angle-of-Attack Projects and Technology Conference, NASACP-3137, vol. 1, 1992.

5Fourth High Alpha Conference, NASA-CP-10143, vol. 1, July 1994.

6Asbury, Scott C. and Capone, Francis J., Multiaxis Thrust-Vectoring Characteristics of a

Model Representative of the F-18 High Alpha Research Vehicle at Angles of Attack from 0 ° to 70 °,

NASA TP-3531, 1995.

7Murri, 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, 1990, pp. 335-380.

8Biedron, Robert T. and Thomas, James L., Navier-Stokes Computations for an F-18 Forebody

with Actuated Control Strake, NASA-CP-3149, vol. 1, pt. 1, 1990, pp. 481-506.

9Lanser, Wendy R. and Meyn, Larry A., "Forebody Flow Control on a Full-Scale F/A-18

Aircraft," J. ofAircrafi, vol. 31, no. 6, 1994, pp. 1365-1371.

10Lanser, Wendy R. and Murri. Daniel G., "Wind Tunnel Measurements on a Full-Scale F/A- 18

with Forebody Slot Blowing or Forebody Strakes," AIAA 93-1018, 1993.

11pahle, Joseph, Foster, John, Bundick, Tom, and Wichman, Keith, "An Overview of Controls

and Flying Qualities Technology on the F/A-18 High Alpha Research Vehicle," High Angle of

Attack Conference, Sept. 1996.

12Hancock, Regis and Fullerton, Gordon, "X-29 Vortex Flow Control Tests," Society of

Experimental Test Pilots 1992 Report to the Aerospace Profession, 36th Symposium Proceedings,

1992, pp. 209-219.

13Walchli, Lawrence A., "Flight Evaluation of Pneumatic Forebody Vortex Control in PostStall Flight, In NASA. Dryden Flight Research Center" Fourth High Alpha Conference, NASA

CP-10143, vol. 3, 1994.

14Murri, Daniel G, and Rao, Dhanvada M., "Exploratory Studies of Actuated Forebody Strakes

for Yaw Control at High Angles of Attack," AIAA 87-2557, 1987.

15Murri, Daniel G., Shah, Gautam H., DiCarlo, Daniel J., and Trilling, Todd W., "Actuated

Forebody Strake Controls for the F- 18 High-Alpha Research Vehicle," J. of Aircraft, vol. 32, no. 3,

1995, pp. 555-562.

16Moes, Timothy R. and Whitmore, Stephen A., A Preliminary Look at Techniques Used to

Obtain Airdata From Flight at High Angles of Attack, NASA TM-101729, 1990.

29
17Fisher, David E, Banks, Daniel W., and Richwine, David M., F-18 High Alpha Research

Vehicle Surface Pressures: Initial In-Flight Results and Correlation with Flow Visualization and

Wind-Tunnel Data, NASA TM-101724, 1990.

18Curry, Robert E. and Richwine, David M., "An Airborne System for Vortex Flow

Visualization on the F-18 High-Alpha Research Vehicle," AIAA/NASA/AFWAL Sensors and

Measurements Technology Conference, 88-4671-CE Atlanta, Georgia, Sept. 7-9, 1988.

19Fisher, David E, Del Frate, John H., and Richwine, David M., In-Flight Flow Visualization

Characteristics of the NASA F-18 High Alpha Research Vehicle at High Angles of Attack, NASA

TM-4193, 1990.

20Fisher, David E and Lanser, Wendy R., "Flight and Full-Scale Wind-Tunnel Comparison of

Pressure Distributions From an F-18 Aircraft at High Angles of Attack," Fourth High Alpha

Conference, NASA CP-10143, vol. 1, 1994.

21Whitmore, Stephen A. and Moes, Timothy R., "The Effects of Pressure Sensor Acoustics on

Airdata Derived from a High Angle-of-Attack Flush Airdata Sensing (HI-FADS) System,"

AIAA 91-0671, Jan. 1991.

22Murri, Daniel G., Fisher, David E, and Lanser, Wendy R., "Flight-Test Results of Actuated

Forebody Strake Controls on the F-18 High-Alpha Research Vehicle," High Angle of Attack

Conference, Sept. 1996.

23Lanser, Wendy R., Meyn, Larry A., Botha, Gavin A., James, Kevin D., Hall, Robert M., and

Murri, Daniel G., "Overview of Full-Scale F/A-18 Tests in the 80- by 120-ft Wind Tunnel with

Small-Scale Comparisons," NASA High Angle of Attack Conference held at Langley Research

Center, Sept. 17-19, 1996.

24Fisher, David F. and Cobleigh, Brent R., Controlling Forebody Asymmetries in

Flight-Experience with Boundary Layer Transition Strips, NASA TM-4595, 1994. (see also

AIAA-94-1826, July 1994.).

30

R EPORT DOCU M ENTATION PAG E FormApproved

OMB No. 0704-0188

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

October 1996 ] Technical Memorandum

4. TITLE AND SU BTITLE 5. FUNDING NUMBERS

Effect of Actuated Forebody Strakes on the Forebody Aerodynamics of the

NASA F- 18 HARV

6. AUTHOR(S)

David F. Fisher, Daniel G. Murri, Wendy R. Lanser

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

NASA Dryden Flight Research Center

P.O. Box 273

Edwards, California 93523-0273

9.SPONSORING/MONITORING AGENCYNAME(S)ANDADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

WU 5O5-68-30

8. PERFORMING ORGANIZATION

REPORT NUMBER

H-2136

10. SPONSORING/MONrrORING

AGENCY REPORT NUMBER

NASA TM-4774

11. SUPPLEMENTARY NOTES

Presented at the NASA Langley High Angle of Attack Technology Conference, Langley Research Center, Sept. 17-19,

1996. David F. Fisher, NASA Dryden Flight Research Center, Edwards, California; Daniel G. Murri, NASA Langley

Research Center, Hampton, Virginia; Wendy R. Lanser, Ames Research Center, Moffet Field, California.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified--Unlimited

Subject Category 02

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

Extensive pressure measurements and off-surface flow visualization were obtained on the forebody

and strakes of the NASA F-18 High Alpha Research Vehicle (HARV) equipped with actuated forebody strakes. Forebody yawing moments were obtained by integrating the circumferential pressures on

the forebody and strakes. Results show that large yawing moments can be generated with forebody

strakes. At angles of attack greater than 40 °, deflecting one strake at a time resulted in a forebody yawing moment control reversal for small strake deflection angles. At 0_ = 40 ° and 50 °, deflecting the

strakes differentially about a 20 ° symmetric strake deployment eliminated the control reversal and produced a near linear variation of forebody yawing moment with differential strake deflection. At o_= 50 °

and for 0 ° and 20 ° symmetric strake deployments, a larger forebody yawing moment was generated by

the forward fuselage (between the radome and the apex of the leading-edge extensions), than on the

radome where the actuated forebody strakes were located. Cutouts on the flight vehicle strakes that

were not on the wind tunnel models are believed to be responsible for deficits in the suction peaks on

the flight radome pressure distributions and differences in the forebody yawing moments.

14. SUBJECT TERMS

Angle-of-attack, F- 18 aircraft, Flight control, Flight test, Pressure distribution,

Vortices, Yawing moments

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

OF REPORT OF THIS PAGE OF ABSTRACT

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Unthicum Heights, MD 21090; (301)621-0390

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34

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AO3

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