National Aeronautics and Space Administration

NASA Technical Memorandum 104261

A Summary of the Forebody High-Angle-of-Attack Aerodynamics Research on the F-18 and the X-29A Aircraft

Lisa J. Bjarke, John H. Del Frate, and David F. Fisher

November 1992

National Aeronautics andSpace Administration

Dryden Flight Research Facility

Edwards, California 93523-0273

1992

NASA Technical Memorandum 104261

A Summary of the Forebody High-Angle-of-Attack Aerodynamics Research on the F-18 and the X-29A Aircraft

Lisa J. Bjarke, John H. Del Frate, and David F. FisherNASA Dryden Flight Research Facility, Edwards, California

A Summary of the Forebody High-Angle-of-AttackAerodynamics Research on the F-18 and the

X-29A Aircraft

Lisa J. Bjarke, John H. Del Frate, and David F. FisherNASA Dryden Flight Research Facility

ABSTRACT

High-angle-of-attack aerodynamic studies havebeen conducted on both the F-18 High Alpha Re-search Vehicle (HARV) and the X-29A aircraft. Dataobtained include on- and off-surface flow visualiza-tion and static pressure measurements on the fore-body. Comparisons of similar results are madebetween the two aircraft where possible. The fore-body shapes of the two aircraft are different and theX-29A forebody flow is affected by the addition ofnose strakes and a flight test noseboom. The fore-body flow field of the F-18 HARV is fairly symmetricat zero sideslip and has distinct, well-defined vorti-ces. The X-29A forebody vortices are more diffuseand are sometimes asymmetric at zero sideslip.These asymmetries correlate with observed zero-sideslip aircraft yawing moments.

INTRODUCTION

Personnel at NASA are currently involved in sev-eral high-angle-of-attack research programs, eitheras a part of the High Alpha Technology Program(HATP) or in joint research programs with other U.S.and international government agencies. The empha-sis on high-angle-of-attack research resulted fromthe philosophy that modern fighter aircraft should becapable of controlled flight at high angles of attack.Two of the flight research programs at the NASADryden Flight Research Facility utilize the F-18 HighAlpha Research Vehicle (HARV) and the X-29A air-craft. The F-18 HARV project is part of the HATP,which seeks to provide design guidelines and newconcepts for vortex control on advanced, highly ma-neuverable aircraft at high angle of attack. The F-18HARV serves as a validation and demonstrationtool, using results from wind-tunnel and flight

research to validate and update computational fluiddynamics (CFD) codes. The X-29A high-angle-of-attack program has been a joint program betweenthe U.S. Air Force (Wright Laboratories and FlightTest Center), NASA, and Grumman Aircraft. Themain emphasis of the X-29A high-angle-of-attackprogram has been in flight controls, handling quali-ties, and military utility and agility research.

Although the F-18 HARV and X-29A aircrafthave been used for high-angle-of-attack research,the projects were operated from different philoso-phies. From the beginning of the F-18 HARV projectthere were plans to use flow visualization and pres-sure measurements to help define the aerodynamicsof the aircraft at high angles of attack. Therefore, in-strumentation to accomplish these objectives wasincorporated early in the program and given a highpriority. Conversely, on the X-29A project, flow visu-alization and pressure measurements were per-formed as part of a follow-on program. This follow-onprogram was initiated because some of the X-29Ahigh-angle-of-attack flight characteristics were quitedifferent than predicted.1 It was anticipated that abetter understanding of the forebody aerodynamicscould help explain the differences, given the successof the F-18 HARV experiments.2,3,4

The results from the F-18 HARV program includeboth on- and off-surface flow visualization andpressure measurements for the forebody and theleading-edge extension (LEX). In addition, surfaceflow visualization of the fuselage aft of the canopy,wing, and vertical tails are included. Flow visualiza-tion results are documented in Refs. 2 and 3and pressure distribution results are found in Refs. 3and 4. Results from the X-29A follow-on program in-clude off-surface flow visualization and pressuremeasurements for the forebody and surface flow vi-sualization of the wing and vertical tail.5,6 This paper

will summarize the forebody aerodynamics researchdone on both aircraft and compare results wherepossible.

AIRCRAFT DESCRIPTION

F-18 HARV - The NASA HARV (Fig. 1) is asingle-place preproduction F-18 aircraft built by theMcDonnell Douglas (St. Louis, Missouri) andNorthrop (Newbury Park, California) corporations. Itis powered by two GE (General Electric, Lynn, –Massachusetts) F404-GE-400 afterburning turbo-fan engines. The aircraft features a midwing withleading- and trailing-edge flaps. Leading-edge ex-tensions (LEXs) are mounted on each side of the fu-selage from the wing roots to just forward of thewindscreen. The aircraft has twin vertical stabilizerscanted out to 20° from the vertical and differential all-moving horizontal tails.

The aircraft is flown in the fighter escort configu-ration without stores. The production LEX fenceshave been removed from the aircraft. The aircraftcarries no missiles and the wingtip Sidewindermissile launch racks have been replaced with spe-cial camera pods and wingtip airdata booms.7 Theflight test noseboom has been removed from the air-craft and a NASA flush airdata system8 has beeninstalled.

X-29A AIRCRAFT - The X-29A aircraft (Fig. 2(a))is a technology demonstrator built by the GrummanAircraft Corp. (Bethpage, New York). It is poweredby one General Electric F404-GE-400 afterburningturbofan engine. The aircraft features a forward-swept wing, close-coupled canards, aft bodystrakes, and relaxed static stability.1,9,10 The wingincorporates double-hinged trailing-edge flaps thatare scheduled as a function of free-stream Machnumber, pressure altitude, and angle of attack(α).The aircraft has one vertical stabilizer and the aftbody strakes incorporate flaps, which generally mir-ror the canard deflection. The all-movable canardsdeflect symmetrically and are scheduled as a func-tion of free-stream Mach number, pressure altitude,and angle of attack. The X-29A aircraft uses an F-5Aforebody that was modified by shortening it by 11 in.and adding a nose strake and a flight test noseboomat the apex.9,10 The noseboom and strakes are indi-cated in Fig. 2(b).

EXPERIMENTAL SETUP

F-18 HARV - The off-surface flow visualizationused a smoke generation system11,12 which ducted

smoke to the forebody vortex cores at high angles ofattack. The smoke was generated by pyrotechniccartridges located inside the forebody. Twelve car-tridges were carried on board. The number of car-tridges ignited at one time could be varied, buttypically four cartridges were used for each testpoint, resulting in three test points for each flight.Data were obtained at steady-state and dynamicflight conditions. Time-correlated onboard video andstill cameras were used to document the off-surfaceflow visualization data. The camera locations andsmoke generator system locations are indicated inFig. 1. The smoke ports were located symmetricallyon both sides of the aircraft near the nose and at theLEX apex, which is also indicated in Fig. 1.

The on-surface flow visualization utilized theemitted fluid technique.13–15 The emitted fluid tech-nique used a small quantity (approximately 1 qt) of asolvent, propylene glycol monomethyl ether(PGME), and a toluene-based red dye. This fluidwas emitted slowly from five circumferential rings onthe F-18 HARV forebody (Fig. 3) while the aircraftwas stabilized at the flight test conditions. As the flu-id flowed back along the surface, the PGME evapo-rated, leaving the dye to mark the surfacestreamlines. This technique required the pilot to sta-bilize at the test conditions for 75 to 90 sec while thePGME evaporated and the dye was set. The result-ing dye traces were photographed on the groundpostflight, allowing one test point to be obtained foreach flight.

Pressure measurements were made on the F-18HARV forebody at the same five fuselage stationsused for PGME visualization, forward of the canopyusing rings of static pressure orifices at nondimen-sional length (x/l) = 0.015, 0.038, 0.071, 0.126, and0.190 (Fig. 3). Details about the number of orifices ineach ring can be found in Ref. 4. This reference alsocontains details about the discontinuities and protru-sions present on the F-18 HARV forebody.

X-29A AIRCRAFT - The X-29A forebody vorti-ces were visualized with smoke using the samemethod employed on the F-18 HARV. The smoke-generating system was located in the X-29A fore-body. However, since space was limited only fourcartridges could be carried on board. A flexible ductrouted the smoke from the cartridges to a “Y” whichdiverted smoke to an exhaust port on each side ofthe aircraft (Fig. 2(b)). All four cartridges were re-quired for adequate smoke density, resulting in onesmoke test point for each flight. The right side of theforebody was painted flat black to provide the maxi-mum contrast between the white smoke and the

2

Fig. 1 F-18 HARV smoke generator system and camera locations.

background when viewed by the wingtip cameras.The camera locations are indicated in Fig. 2(a).

Pressure measurements were made on theX-29A aircraft using circumferential rings of staticpressure orifices. Four rings were installed ahead ofthe cockpit at x/l = 0.026, 0.056, 0.136, and 0.201(Fig. 4) and 202 orifices were installed. Gaps in theorifice distribution were caused by internal structureor lack of internal access. The X-29A forebodysurface was considered to be smooth and free ofprotuberances typically found on operational aircraft.

Fig. 3 F-18 HARV forebody pressure measurementstations.

(a) Overall view.

Fig. 2 X-29A aircraft.

LEX

Nose static pressure rings

x/ = 0.015

x/ = 0.038

x/ = 0.071

x/ = 0.126x/ = 0.190

920627

(b) Closeup of nose apex, side view.

Fig. 2 Concluded.

3

INSTRUMENTATION

PRESSURE MEASUREMENTS - The instru-mentation used for the F-18 HARV and the X-29Aaircraft was quite similar. Each orifice on the fore-body was connected to temperature-controlledelectronic scanning pressure modules with 6 ft of0.062 in. i.d. pneumatic tubing. It was previously de-termined that 8 ft of 0.062 in. flexible tubing wouldhave a pneumatic lag of approximately 10 msec atan altitude of approximately 20,000 ft,8 which wasacceptable. Reference pressure was supplied by areference tank vented to the F-18 HARV forwardfuselage4 and by a small reference manifold ventedinside the X-29A forebody.5 The reference pressurewas monitored by a high-resolution digital absolutepressure transducer. The pressure transducerswith each module were scanned sequentially25 samples/sec and outputs were sampled by a10-bit pulse code modulation (PCM) data system.In-flight zero differential pressure readings were tak-en before each test point and were used postflight tocorrect the data for calibration offsets. The forebodypressures were measured with approximately216 Ib/ft2 differential range pressure transducerswith an estimated accuracy of approximately 1 Ib/ft2.

FREE-STREAM AIRSPEED AND ALTITUDE -Airspeed and altitude were measured on bothaircraft using a specially designed swivel probewhich self aligned with the local flow. A swivel probewas mounted on the left wingtip of both the F-18HARV and X-29A aircraft. The probes were calibrat-ed for Mach number and altitude.7

FREE-STREAM FLOW ANGLES - The F-18HARV flow angle measurements were taken fromthe two wingtip booms.7 Angle of attack was mea-sured by using a vane on the right wingtip boom. Themeasurement was then corrected for upwash andboom bending. Angle of sideslip was determined byaveraging the left- and right-wingtip boom sideslipvane measurements corrected for angle of attack.

On the X-29A aircraft, angle of attack was aflight-critical input parameter to the triple-redundantflight-control system. Therefore, three independentangle-of-attack vanes were mounted on the nose-boom. For high angles of attack, the vanes were cal-ibrated using the aircraft inertial navigation systemand meteorological analysis of rawinsonde balloondata.16,17 A single vane mounted on the noseboomwas used to determine angle of sideslip.

TEST CONDITIONS

F-18 HARV - The on- and off-surface flow visu-alization data were obtained during 1-g flight condi-tions. The nominal altitudes were between 20,000and 30,000 ft and the Mach numbers varied from ap-proximately 0.2 to 0.4. Angles of attack ranged from10.0° to approximately 54.0° over the course of theflight program. This paper presents F-18 HARV flowvisualization results only for α = 26.0° to 47.7°.

Surface pressure data presented were obtainedin quasi-stabilized, 1-g flight maneuvers. Datawere obtained at nominal altitudes of 20,000 and45,000 ft. The data presented are for α = 10.0° to50.0° and 0° angle of sideslip (β).

X-29A AIRCRAFT - The off-surface flow visual-ization data were obtained during 1-g flight. Theoff-surface flow visualization data presented rangefrom α = 25.5° to 50.5°. These test points were flownat altitudes between 17,000 and 30,000 ft.

Pressure distributions on the forebody were ob-tained at angles of attack from 15.0° to 50.0° during1-g quasi-steady-state flight conditions at nominalaltitudes of 20,000 and 40,000 ft. Pressure distribu-tions at α > 55.0° were obtained on a single flightduring a pullup-pushover maneuver of which 6.5 secwere at α ≥ 50.0°. As mentioned in the Instrumenta-tion section, there was little lag in the pneumatic tub-ing between the orifice and the pressure transducer.At α ≤ 55.0° data from this dynamic maneuver wereconsistent with similar data from stabilized testpoints on other flights.

RESULTS

F-18 HARV OFF-SURFACE FLOW VISUAL-IZATION - Figure 5 shows wingtip view photographsof the F-18 HARV forebody vortices at two angles ofattack. At α = 29.5° and β = 0.4°, the forebody vortexcores stay quite close to the fuselage, pass over thecanopy, and continue straight aft. At α = 47.0° andβ = 0.7°, the forebody vortex cores are farther awayfrom the surface, arch higher over the canopy, and

x/ = 0.201920628

x/ = 0.136x/ = 0.056x/ = 0.026

Fig. 4 Location of X-29A forebody pressure orifices.

4

then are pulled down into the LEX vortex.2,3 At side-slip, the forebody vortex core position changes as il-lustrated in Fig. 6. This wingtip view photograph (andaccompanying illustration) show the position of theforebody vortex cores at α = 45.1° and β = –5.5°. Thewindward (left) forebody vortex core shifts to theright and away from the surface, travels over thecanopy and straight aft. The leeward (right) forebodyvortex core also shifts to the right, but this shift bringsit into close proximity to the LEX vortex and draws itinto the LEX vortex.2,3

X-29A OFF-SURFACE FLOW VISUALIZA-TION - Figure 7 presents wingtip photographs of theX-29A forebody vortices at α = 25.0° and 50.5°. Onthe X-29A forebody, the smoke entrained into thevortices is rather diffuse; the cores do not appear astight and distinct as those seen on the F-18 HARVforebody. As angle of attack increases from 25.0° to50.5°, the vortices lift farther away from the surfaceaft of the canopy. Figure 8 is a photograph showing

a vertical tail-view at α = 33.2° and β = 1.0°. Thecores appear as two white lobes over the canopy.These lobes shift right and left as a pair with sideslip,and the windward vortex core shifts away from theaircraft surface. The lobes are separated by a dark“midplane” region. This midplane was considered tobe representative of the angular position of the vor-tex system over the canopy. To analyze the behaviorof this vortex system, the angular position was mea-sured from the video images from the tail. Figure 9 il-lustrates how this angle, θv, is defined. θv is the anglebetween the midplane and the vertical plane. It is de-fined to be positive to the right, as seen by the tailcamera, and negative to the left. At each angle of at-tack investigated, θv was plotted as a function ofsideslip. Although there was some scatter in the da-ta, the relationship was fairly linear and a linear ap-proximation was sketched through the data set.Figure 10 shows these linear approximations at an-gles of attack ranging from α = 25.5° to 50.5°. All the

5

(a) α = 29.5°, β = 0.4°.

(b) α = 47.7°, β = 0.7°.

Fig. 5 Wingtip view of F-18 HARV forebody vortices.

6

Fig. 6 Wingtip view of F-18 HARV forebody vortices in sideslip, α = 45.1°, β = –5.5°.

EC91 498-46 EC91 536-25

(a) α = 25.0°, β = 2.8°. (b) α = 50.5°, β = 3.3°.Fig. 7 Wingtip view of X-29A forebody vortices.

0° +10°-10°-20° +20°

-30° +30°

-40° +40°

θv

Forebody vortex pair

Canopy outline

920308

Fig. 8 Tail camera view of X-29A forebody vortices,α = 33.2°, β = 1.0°.

Fig. 9 X-29A forebody-vortex measurementtechnique.

slopes of these lines are similar, but the bias is notalways zero at zero sideslip. When the vortex pair isshifted in one direction, for example to the left, thereis more attached flow on the right side of the fore-body. This would produce lower pressure on theright side than the left, causing a net force to theright. Conversely, when the vortices are shifted tothe right, a nose-left force results. This hypothesis issupported by comparing the X-29A yawing momentat β = 0° ( ) with the vortex angular position at β =0° as shown in Fig. 11.

The F-18 HARV and the X-29A aircraft have dif-ferent forebody shapes. The apex of the F-18 HARVforebody has a circular cross section which transi-tions to an elliptical cross section with the major axisalong the vertical. The X-29A forebody is actually amodified F-5A forebody. The cross section is also el-liptical; however, the major axis is along the horizon-tal. Further aft on the forebody, this elliptical crosssection becomes squared at the major axis.

The F-18 HARV and X-29A forebody vortices donot behave in the same manner at high angles of at-tack. The F-18 HARV forebody vortices have fairlywell-defined cores, which arch over the canopy andget pulled down into the LEX vortices at the higherangles of attack. The X-29A forebody vortices aremore diffuse (as visualized by the smoke generatingsystem) with no well-defined cores visible. TheX-29A vortex path is fairly straight aft of the canopy.

In sideslip, the F-18 HARV windward vortex shiftsaway from the surface and the leeward vortex shiftstoward the surface and interacts with the LEX vor-tex.2,3 With sideslip, the X-29A forebody vortexcores generally shift left and right as a pair and overthe forebody; there are no major shifts in the positionvertically. The respective forebody cross-sectionaldifferences between the F-18 HARV and X-29A air-craft may be a cause for the differences observed.However, the noseboom and nose strakes on theX-29A forebody have an effect as well. On the X-29Aforebody there is no additional strong vortex system(similar to the F-18 HARV LEX vortices) to interactwith the forebody vortices and affect their verticalposition.

SURFACE FLOW VISUALIZATION - Two formsof surface flow visualization were used during theF-18 HARV program. The first was the emitted fluidtechnique (on the forebody) and the second was flowcones and tufts (on the wing, fuselage, and verticaltails). The emitted fluid technique was not used dur-ing the X-29A program because of lack of space andelectrical concerns. However, flow cones and tuftswere used on the canard, wing, aft fuselage, andvertical tail. This section will only discuss F-18 HARVsurface flow visualization results obtained using theemitted fluid technique. Results from the F-18 HARVand X-29A flow cone and tuft observations can befound in Refs. 2 (F-18 HARV) and 6 (X-29A).

Cn0

7

Fig. 10 X-29A approximated forebody-vortex system position as a function of sideslip.

Angle of attack25.5°35.0°41.3°46.0°50.5°

1050β, deg

-5-10-30

-20

-10

0

10

20

30

40

θv,

deg

920309

Fig. 11 Comparison of X-29A yawing moment at β = 0° and forebody vortex system position (θv).

–4

0

4

706050403020100–.04

–.02

0

.02

.04

Cn0

α, deg920630

θv

θv,deg

Nose right

Nose left

Cn0

Results from the emitted fluid technique werephotographed postflight, thus only one test pointcould be obtained during each flight. This techniquewas used for surface flow visualization on the fore-body and the LEX. The emitted fluid techniquemarks surface streamlines with the red dye left be-hind after the PGME evaporated during the testpoint. Where the flow streamlines merge, lines offlow separation are defined. Conversely, where thestreamlines diverge, lines of reattachment are de-fined. Since the fluid flows away from the reattach-ment line, reattachment is visible only in the dye neara source of fluid.

Figure 12 shows two photographs of the fore-body taken after the emitted fluid technique wasused at α = 26.0°. Both the primary and secondaryforebody separation lines are visible and are nearlysymmetric. Only the primary vortex had been visibleduring smoke flow visualization. This may be be-cause the secondary vortices form farther aft on theforebody than the smoke ports and are weaker. Theemitted fluid results do not indicate the presence of avortex until approximately x/l = 0.126 at this angle ofattack. There are some small kinks or curves instreamlines, which indicate a laminar separationbubble (LSB). This will be discussed in more detaillater.

Figure 13 shows an example of surface flow vi-sualization on the forebody at α = 47.0°. The primarydifferences between α = 47.0° and α = 26.0° are that

the streamlines are more smeared at the higher an-gles of attack and that the secondary vortex separa-tion lines have moved forward. At α = 47.0°, theyappear at approximately x/l = 0.038 as opposed tox/l = 0.126 at α = 26.0°. The smearing of the stream-lines is simply because the flight conditions weremore difficult to hold steady at the higher angles ofattack. Although the separation lines are smeared,they are nearly symmetric.

Further and more definitive indications ofboundary-layer transition on the forebody were evi-dent at α = 47.0° (Fig. 13). The effect of theboundary-layer transition is seen in the closeup viewin Fig. 14. A large dye puddle is noted extending in-termittently from θ = 240° at x/l = 0.015 to approxi-mately x/l = 0.075 and θ = 247°. Though notpresented, symmetric results were obtained onthe left side at θ = 129° and 113° at x/l = 0.015and 0.075, respectively. It is believed that these pud-dles are the result of an LSB with boundary-layertransition occurring downstream. The dye puddle didnot occur at the screwhead protuberances aroundthe plugged smoker port. (These screwhead protu-berances would cause premature transition.) In ad-dition, the fluid windward of the LSB flowed towardthe LSB and the fluid leeward of the LSB flowedaway. This indicates that the flow reattached turbu-lently past the very localized LSB and that this is notthe primary vortex separation line.

8

(a) 1/4 view.

(b) Head-on view.

Fig. 12 Surface flow visualization on F-18 HARVforebody, α = 26.0°.

(a) 1/4 view.

(b) Head-on view.

Fig. 13 Surface flow visualization on F-18 HARVforebody, α = 47.0°.

9

Fig. 14 Closeup of nose cone of F-18 HARV, α = 47°.

F-18 HARV FOREBODY PRESSURES -Figure 15 shows the F-18 HARV forebody pressuredistributions over an angle-of-attack rangefrom 10.0° to 50.0°. On the forward three rows (x/l =0.015, 0.038, and 0.071) the flow acceleratingaround the forebody produces a pair of maximumsuction peaks starting at α = 19.7° (Figs. 15(a)–(c)).The angular location of these peaks were at θ ≈ 100°to 120° and θ ≈ 240° to 260°. These suction peaksbecame more pronounced with increasing angle ofattack. The “footprints” of the primary vortex coresare first visible at x/l = 0.038 and 0.071 at α = 34.3°(Figs. 15(b)–(c)). The footprints are indicated by thesuction peaks at θ = 168° and 192°. As angle of at-tack increases, these peaks become more negative.The pressure distributions for the three forward rowsare symmetric at β = 0° at θ ≈ 180° up to α = 50.0°(Figs. 15(a)–(c)).

As shown in Fig. 15(d), at x/l = 0.126, the maxi-mum suction peaks are indicated at θ = 70°and 290°. The sharp peaks in the pressure distribu-tion appearing at α ≥ 19.7° at θ = 90° to 110° and θ =270° to 250° are caused by local separation behindantenna covers. The angular location of these pointsmoves up as angle of attack increases. The forebodyvortex core footprints are first indicated at α = 25.8°at θ = 160° and 200°. These footprint peaks becomeincreasingly negative as angle of attack increasesto 45.4°, then diminish at α = 50.0° indicating the

vortices are lifting away from the surface. Aside fromthe differences caused by local protuberances ordiscontinuities, the pressure distributions are gener-ally symmetric at θ ≈ 180°.

As shown in Fig. 15(e), at x/l = 0.190, the maxi-mum suction peaks have moved up to θ = 120°and 240° because of the influence of the LEX. TheLEX apex is located only 13 in. aft of this fuselagestation at θ = 123° and 237°. The maximum suctionpeaks are diminished from those seen at x/l = 0.126and forward. The primary vortex footprints at x/l =0.190 are indicated at θ = 165° and 195° at α > 25.8°,but are smaller in magnitude than those at x/l =0.126, indicating they are even farther off the sur-face. The pressure peaks on x/l = 0.190 at θ ≈ 48°to 60° and θ ≈ 300° to 312° for α ≥ 34.3° are causedby local separation behind the aircraft productionpitot-static probes.

X-29A FOREBODY PRESSURES - Figure 16shows the X-29A forebody pressure distributionsover an angle-of-attack range from 14.9° to 66.2°. Aschematic of the forebody cross section is alsoshown. The pressure distributions at x/l = 0.026(Fig. 16(a)) are different from those seen on the F-18HARV (Fig. 15) in that the maximum suction peak iscaused by the nose strake vortex rather than wherethe flow accelerates around the forebody. Thesesuction peaks are at θ ≈ 108° and 252° and generallyincrease in magnitude with angle of attack. The

10

11

(a) x/l = 0.015. (b) x/l = 0.038.

Fig. 15 Effect of angle of attack on F-18 HARV forebody pressures.

1

1

1

1

1

1

1

1

1

Cp

0

-1

-2

0 60 120 180 240 300 360

θ, deg

10.0°

50.0°

45.4°

PortStarboard

α =

15.2°

19.7°

25.8°

30.0°

34.3°

39.3°

0 60 120 180 240 300 360

θ, deg

10.0°

50.0°

45.4°

PortStarboard

α =

15.2°

19.7°

25.8°

30.0°

34.3°

39.3°

θ = 0°

90°

180°

270°

θ = 0°

90°

180°

270°

920636

12

(c) x/l = 0.071. (d) x/l = 0.126.

Fig. 15 Continued.

1

1

1

1

1

1

1

1

1

Cp

0

-1

-2

0 60 120 180 240 300 360

θ, deg

10.0°

50.0°

45.4°

PortStarboard

α =

15.2°

19.7°

25.8°

30.0°

34.3°

39.3°

0 60 120 180 240 300 360

θ, deg

10.0°

50.0°

45.4°

PortStarboard

α =

15.2°

19.7°

25.8°

30.0°

34.3°

39.3°

180°

90° 270°

θ = 0°

180°

90° 270°

θ = 0°

920637

suction peaks are symmetric up to α = 30.1°, at α >30.1° asymmetries develop (Fig. 16(a)). The magni-tude of the port suction peak is greater than the star-board suction peak, indicating the port vortex iscloser to the surface. This asymmetry switches tostarboard at 59.1° ≤ α ≤ 66.2°.

The three aft forebody pressure rows (x/l =0.056, 0.136, and 0.201) (Figs.16(b)–(d)) are behindthe nose strakes and therefore have the maximumsuction peaks caused by the flow acceleratingaround the forebody. The suction peaks caused bythe nose strake vortices diminish in magnitude thefarther aft the measurement location is. The angularlocation of the vortex footprints is θ ≈ 140° and 220°at x/l = 0.056 (Fig. 16(b)) and θ ≈ 160° and 200° atx/l = 0.136 (Fig. 16(c)) and 0.201 (Fig.16(d)). The re-duction in magnitude of the peaks is caused by thevortex lifting away from the surface. The onset ofasymmetries in the pressure distribution is also de-layed as the measurement location moves fartheraft. At x/l = 0.056 (Fig.16(b)), asymmetries appear atα = 49.7° with the higher magnitudes on the portside. At α = 66.2°, the pressure distribution is nearlysymmetric again. At x/l = 0.136 (Fig. 16(c)), the portasymmetries start at α = 54.7° and switch tostarboard at α = 66.2°.

To determine if the asymmetries seen in thepressure distributions contributed to the total aircraftyawing moment at zero sideslip, the pressure distri-butions were integrated over the projected side area.The resultant forebody yawing moment coefficient,

was plotted as a function of sideslip. A linewas faired through the data and the intercept,

was determined. Figure 17 shows the totalaircraft yawing moment coefficient1 and the fore-body yawing moment coefficient plotted as a func-tion of angle of attack. The large right aircraft yawingmoment at zero sideslip at α = 45.0° did not correlatewith the forebody pressures. However, there is agood correlation between total aircraft and forebodyyawing moments at α ≥ 50.0°.

The forebody yawing moments at zero sideslipwere broken down further by individual orifice sta-tions to determine which regions contributed to theyawing moment. Figure 18 shows the yawing mo-ments at β = 0° for a unit length of fuselage at eachstation as a function of angle of attack. The effect forthe most forward row (x/l = 0.026) is small partly be-cause of its small minor diameter (height) and partlybecause of the nose strake. At α ≥ 55.0°, the secondand third forebody stations (x/l = 0.056 and 0.136)have the most effect on the forebody yawing mo-ment to the left. The most aft forebody station (x/l =0.201) has less effect at α ≥ 55.0°. However, there isa right (positive) yawing moment shown at x/l =0.201 at α = 45.0°. This suggests that the nose-rightyawing of the aircraft at α ≈ 45.0° is affected by pres-sures on a region aft of x/l = 0.201 (where there wasno instrumentation).5

1

1

1

1

1

1

1

1

1

Cp

0

-1

-2

0 60 120 180 240 300 360

θ, deg

10.0°

50.0°

45.4°

PortStarboard

α =

15.2°

19.7°

25.8°

30.0°

34.3°

39.3°

θ = 0°

90°

180°

270°

920638

(e) x/l = 0.190.

Fig. 15 Concluded.

Cn fb,

Cn0 fb,,

13

14

(a) x/l = 0.026. (b) x/l = 0.056.

Fig. 16 Effect of angle of attack on X-29A forebody pressures.

15

(c) x/l = 0.136. (d) x/l = 0.201.

Fig. 16 Concluded.

Fig. 17 X-29A aircraft and forebody yawing mo-ments at β = 0°.

Fig. 18 Effect of angle of attack on X-29A forebodystation yawing moment per unit length.

CONCLUDING REMARKS

Aerodynamic studies have been conducted athigh angles of attack on the F-18 High Alpha Re-search Vehicle (HARV) and X-29A aircraft. Data ob-tained include on- and off-surface flow visualizationand pressure measurements. These results can becorrelated with wind-tunnel and computational fluiddynamics (CFD) results. In the case of the X-29A air-craft, the forebody results correlate well with mea-sured aircraft results and help explain differencesfrom predictions. Some differences were observedin the forebody aerodynamics of the two aircraft.

The F-18 HARV pressure distributions weresymmetric at zero sideslip. This symmetry was alsoobserved in the surface flow visualization. On theother hand, the X-29A pressure distributions wereasymmetric at angles of attack (α) > 30°; this corre-lated with flight-measured yaw asymmetries.

The F-18 HARV forebody vortices visualizedwere fairly well defined with distinct cores. At nonze-ro sideslips, the windward vortex core lifted awayfrom the aircraft surface while the leeward vortexcore was drawn into the leading-edge extension(LEX) vortex. The X-29A forebody vortices weremore diffuse and nonzero sideslips tended to shift asa pair when viewed from the tail. The location of theX-29A forebody vortex cores at zero sideslip corre-lated well with flight-measured yawing momentasymmetries. The nose strakes and noseboom onthe X-29A forebody may be partly responsible for thediffusion of the forebody vortex cores.

NOMENCLATURE

Cn yawing moment coefficient (positive right)

forebody yawing moment coefficient determined from integration of forebody pressure over projected side area

yawing moment coefficient at zero sideslip

forebody yawing moment coefficient at zero sideslip, β = 0° intercept of as a function of angle-of-sideslip curve

Cp pressure coefficient

HARV High Alpha Research Vehicle

HATP High Alpha Technology Program

l aircraft length

LEX leading-edge extension

LSB laminar separation bubble

PGME propylene glycol monomethyl ether

x distance from nose apex along longitudinal axis of aircraft(positive aft)

α aircraft angle of attack, deg

β aircraft angle of sideslip, deg

θ forebody circumferential angle, deg (0° is bottom centerline, positive is clockwise as seen from a front view, 0° to 360°)

θv angular location of the midplane between the right- and left-forebody vortices, deg (0° is top center, positive right as viewed from the back of the aircraft)

–.0470

α, deg

–.03–.02–.01

0.01.02.03.04

605040302010

920654

Forebody, 40,000 ft

Aircraft, ref. 1

Forebody, 20,000 ft

Cn0

Cn fb

Cn0

Cn0 fb,Cn fb

16

REFERENCES

1Webster, Fredrick R. and Purifoy, Dana, X-29AHigh Angle of Attack Flying Qualities, AFFTC-TR-91-15, July 1991.

2Fisher, David F., Del Frate, John H., and Rich-wine, David M., In-Flight Flow Visualization Charac-teristics of the NASA F-18 High Alpha ResearchVehicle at High Angles of Attack, NASA TM-4193,1990. Also published as SAE 892222, Sept. 1989.

3Del Frate, John H., Fisher, David F., and Zuni-ga, Fanny A., In-Flight Flow Visualization with Pres-sure Measurements at Low Speeds on the NASAF-18 High Alpha Research Vehicle, NASATM-101726, 1990. Also published in Vortex FlowAerodynamics, AGARD-CP-494, paper no. 13.

4Fisher, David F., Banks, Daniel W., and Rich-wine, David M., F-18 High Alpha Research VehicleSurface Pressures: Initial In-Flight Results and Cor-relation with Flow Visualization and Wind-TunnelData, NASA TM-101724, 1990. Also published asAIAA-90-3018, Aug. 1990.

5Fisher, David F., Richwine, David M., andLanders, Stephen, “Correlation of Forebody Pres-sures and Aircraft Yawing Moments on the X-29AAircraft at High Angles of Attack,” AIAA-92-4105,Aug. 1992.

6Del Frate, John H. and Saltzman, John A., “In-Flight Flow Visualization Results From the X-29AAircraft at High Angles of Attack,” AIAA-92-4102,Aug. 1992.

7Moes, Timothy R. and Whitmore, Stephen A., APreliminary Look at Techniques Used to Obtain Air-data From Flight at High Angles of Attack, NASATM-101729, 1990.

8Whitmore, Stephen A., Moes, Timothy R., andLarson, Terry J., Preliminary Results From a Sub-sonic High Angle-of-Attack Flush Airdata Sensing(HI-FADS) System: Design, Calibration, and FlightTest Evaluation, NASA TM-101713, 1990. Also pub-lished as AIAA-90-0232, Jan. 1990.

9Moore, M. and Frei, D., “X-29A Forward SweptWing Aerodynamic Overview,” AIAA-83-1834, July1983.

10Frei, Douglas, Garelick, Melvin, Hendrickson,Ronald, and Spacht, Glenn, Forward Swept WingStudy, AFFDL-TR-79-3151, Jan. 1980.

11Richwine, David M., Curry, Robert E., and Tra-cy, Gene V., A Smoke Generator System for Aero-dynamic Flight Research, NASA TM-4137, 1989.

12Curry, Robert E. and Richwine, David M., “AnAirborne System for Vortex Flow Visualization on theF-18 High-Alpha Research Vehicle,” AIAA-88-4671,Sept. 1988.

13Fisher, David F., Richwine, David M., andBanks, Daniel W., Surface Flow Visualization ofSeparated Flows on the Forebody of an F-18 Aircraftand Wind-Tunnel Model, NASA TM-100436,1988.Also published as AIAA-88-2112, May 1988.

14Belevtsov, N., Brumby, R.E., and Hughes,J.P., “In-Flight Flow Visualization, A Fluid Ap-proach,” 14th Annual Symposium Proceedings ofthe Society of Flight Test Engineers, Aug.1983, pp.4.2–1 to 4.2–7.

15Hughes, J.P., Brumby, R.E., and Belevtsov,N., “Flow Visualization From the Ground Up—for Air-craft Fuselages,” AIAA-83-2691, Nov. 1983.

16Rajczewski, David M., Capt., “X-29A HighAngle-of-Attack Flight Test: Air Data Comparisons ofan Inertial Navigation System and NoseboomProbe,” 21st Annual Symposium Proceedings of theSociety of Flight Test Engineers, Aug.1990, pp. 4.5–1 to 4.5–12.

17Pellicano, Paul, Krumenacker, Joseph, andVanhoy, David, “X-29A High Angle-of-Attack FlightTest Procedures, Results, and Lessons Learned,”21st Annual Symposium Proceedings of the Societyof Flight Test Engineers, Aug. 1990, pp. 2.4–1 to2.4–24.

17

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A Summary of the Forebody High-Angle-of-Attack AerodynamicsResearch on the F-18 and the X-29A Aircraft

WU 505-68-71WU 533-02-38

Lisa J. Bjarke, John H. Del Frate, and David F. Fisher

NASA Dryden Flight Research FacilityP.O. Box 273Edwards, California 93523-0273

H-1862

National Aeronautics and Space AdministrationWashington, DC 20546-0001 NASA TM-104261

High-angle-of-attack aerodynamic studies have been conducted on both the F-18 High Alpha ResearchVehicle (HARV) and the X-29A aircraft. Data obtained include on- and off-surface flow visualization andstatic pressure measurements on the forebody. Comparisons of similar results are made between the twoaircraft where possible. The forebody shapes of the two aircraft are different and the X-29A forebody flow isaffected by the addition of nose strakes and a flight test noseboom. The forebody flow field of the F-18 HARVis fairly symmetric at zero sideslip and has distinct, well-defined vortices. The X-29A forebody vortices aremore diffuse and are sometimes asymmetric at zero sideslip. These asymmetries correlate with observed zero-sideslip aircraft yawing moments.

F-18 HARV; Forebody pressure distribution; High angle of attack; In-flight flowvisualization; Vortex flow; X-29A aircraft

AO3

21

Unclassified Unclassified Unclassified Unlimited

November 1992 Technical Memorandum

Available from the NASA Center for AeroSpace Information, 800 Elkridge Landing Road, Linthicum Heights, MD 21090; (301)621-0390

Prepared as SAE 921996 for presentation at the Society of Automotive Engineers Aerotech ‘92 Conference,Anaheim, California, October 5–8, 1992.

Unclassified—UnlimitedSubject Category 02