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DYNAMIC STALL EFFECTS AND APPLICATIONSTO HIGH PERFORMANCE AIRCRAFT

byJay M. Brandon

NASA Langley Research CenterMail Stop 355

Hampton, Virginia 23665

SUMMARY

This paper highlights recent research conducted at theNASA Langley Research Center on the effects of large-amplitude pitching motions on the aerodynamiccharacteristics of modern fighter airplaneconfigurations. Wind tunnel tests were conducted onsimple flat-plate wings to gain understanding of thecomplex flow phenomena during unsteady motions athigh angles of attack. Studies then progressed to arepresentative modern fighter configuration. Using acomputer controlled dynamic apparatus, tests wereconducted to investigate effects of pitch rate andmotion time history, and to determine the persistenceof unsteady effects. Data were also obtained insideslip and with control surface deflections toinvestigate dynamic effects on lateral stability andavailable control power. Force and moment data wereobtained using a 6-component internal strain-gagebalance. To aid in the interpretation of the results,flow visualization using a laser light-sheet system wasalso obtained. Results of these tests will be discussed,along with their implications on the maneuverabilityof future advanced airplanes designed to operate in thehighly dynamic, high angle-of-attack environment.

SYMBOLS

c mean aerodynamic chord, ftCD drag coefficientCl rolling moment coefficientCL lift coefficientCM pitching moment coefficientCN normal force coefficient

k reduced frequency, ωc2V

L/D lift-to-drag ratio, CL/CDq pitch rate, rad/sec

q non-dimensional pitch rate,

qc

2Vt time, sectc convective time unitV free stream velocity, ft/secα angle of attack, deg

α0 mean angle of attack

ω oscillation frequency, rad/secLEX leading-edge extension

INTRODUCTION

At high angles of attack, unsteady aerodynamic effectsmay have a major impact on the maneuverability andcontrollability of an airplane. Currently, somemodern fighter airplanes are capable of performingtransient maneuvers involving high pitch rates toextreme angles of attack; this has been graphicallydemonstrated in the so-called "Cobra maneuver" flownby Soviet aircraft at recent airshows. The advent ofinnovative high-α control effectors such as thrustvectoring and forebody controls will enable evergreater capability to effectively exploit a

substantially enlarged envelope for air combat. Withthe emphasis on aggressive maneuvering capabilitynear or beyond the stall angle of attack for futureairplanes, research is needed to understand the effectsof large-amplitude unsteady motions at high angles ofattack on stability, control and performance.

A recent paper (ref 1) discussed the potential tacticalbenefits of unsteady aerodynamic phenomena;however, little research has been conducted onrealistic three-dimensional configurations to quantifyunsteady effects or to develop an understanding of theflow mechanisms involved. Additionally, the impactof unsteady aerodynamic effects on airplane flightdynamics (including stability and control effectsduring rapid high angle-of-attack maneuvers), and apractical means of utilizing these unsteady effectshave to be addressed (ref 2).

A significant amount of research has been conducted inthe area of unsteady aerodynamics at high angles ofattack. However, most of the work to date has been ontwo-dimensional airfoils. These results showsubstantial force overshoots for pitching andplunging airfoils (refs 3-5). Current design trends foradvanced airplane configurations show a clear need forextending the research to three-dimensional airplaneconfigurations. Wind tunnel experiments have beenconducted using delta wings undergoing pitching andplunging oscillations to determine vortical flowcharacteristics during large-amplitude motions (refs 6-8). The primary focus of the studies was the dynamicsof the leading-edge vortex while the wing wasundergoing motion. Under these conditions, thevortex burst point location was observed to lag thestatic location during pitch-up and pitch-downmotions. Water tunnel results (ref 9) have shownvortex lag times of up to 30 convective time units(time required for an air particle to travel across thewing), which is a much slower response than thatnormally associated with two-dimensional dynamicstall phenomena. More recently, research wasconducted on a series of wings of different sweepangles which showed that the magnitude of theunsteady aerodynamic effects decreased as the sweepangle increased (ref 10). Additional tests of semi-spanmodels with various wing sweeps were reported inreference 11. Reference 12 extended this work to arepresentative three-dimensional configuration, testedat zero sideslip and with no control deflections, whichshowed significant force and moment increments dueto dynamic stall effects. Measured persistence timesof the dynamic effects were also presented. Reference13 presented detailed wing surface pressuremeasurements of a straked-wing fighter modelundergoing pitch oscillations. Force and momentmeasurements from water tunnel studies with a seriesof straked-wing models have been reported in reference14.

This paper presents a summary of low-speed wind-tunnel tests of flat-plate wings and of a modern fighterconfiguration undergoing large-amplitude pitching

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motions over an angle-of-attack range from 0° to 75°.Fundamental flow characteristics associated withdynamic stall are most evident in the simple wingmodels; however, those phenomena are also observedin the more complex actual aircraft geometry. Thisresearch has provided fundamental information ondynamic stall aerodynamics and potential impact onmaneuvering flight dynamics of modern fighteraircraft. Much additional work is needed to enableaccurate prediction of these effects, to provide reliabledesign methods, and to develop flow control conceptsto exploit these dynamic phenomena for maneuveringenhancement.

FLAT-PLATE WING PLANFORMS

Low-speed wind tunnel tests were conducted on threeflat-plate wing planforms to study basic 3-D dynamicstall flow mechanisms, and the effect of wing sweepon these phenomena. The three models, with wingsweeps of 70°, 45° and 0°, were tested on a pitchoscillation rig which oscillated the model sinusoidallyat various reduced frequencies and about various meanangles of attack. A photograph of the models isshown in figure 1. The amplitude of the oscillationswas held constant for these tests at 18°. The mountingsystem oscillated the models about their 40%clocation. The tests were conducted in the Langley 12-Foot Low-Speed Wind Tunnel at a dynamic pressure of4 psf resulting in a Reynolds number of 0.4x106 perfoot.

Data recorded during test runs included a potentiometerreading of the oscillation rig pitch angle, 6-component force and moment data from the strain-gage balance, and tunnel dynamic pressure. Theseinputs were obtained at a rate of 100 samples/sec perchannel. All input signals were electronically filteredby a 10 Hz low-pass filter. After data acquisition,weight and inertia tares were subtracted and the datawere reduced to coefficient form.

70° Delta Wing

It is well known that vortical flow plays a dominantrole in the high angle-of-attack characteristics ofmodern fighter aircraft. Since the flow field over a 70°delta wing is dominated by the presence of a strongleading-edge vortex system, the flat plate wing is aneffective tool for the study of unsteady effects onvortical flows.

Longitudinal Characteristics- As angle of attack wasincreased from 0°, the leading-edge vortex systembegan to develop on the wing. As angle of attack wasfurther increased, the vortex burst point traveledforward onto the wing. Figure 2 shows the effect ofangle of attack and pitch rate on vortex burst locationas determined by flow visualization tests. Statically,at α=22.5°, the burst point of the leading-edgevortices reached the trailing edge of the wing. Asangle of attack was further increased, the loss invortex lift over the rear of the wing due to the burstvortices resulted in a decrease in the CL-α slope and an

increase in the nose-up pitching moment. At α=42°,the static vortex burst point reached the apex of thewing, and the flow field over the wing wascharacterized by fully separated flow.

Numerous studies have documented the hysteresis andovershoot characteristics exhibited by delta wingsundergoing oscillations. Figure 2 shows the flowmechanism responsible for observed dynamic effects.The effect of pitch rate can be seen by comparing thedynamic vortex burst point positions with the staticlocation. At a reduced frequency of k=.0426, duringpitch-up motion, the dynamic burst point positionlagged the static location by approximately 24%chord at 32° angle of attack. As reduced frequency wasincreased, the lag in the burst point locationincreased. As the model motion reversed direction andstarted to pitch down, the vortex burst point againlagged the static value for the remainder of the pitchoscillation cycle. At the faster reduced frequency, itwas seen that the vortex burst point never reached theapex of the wing.

The vortex flow lags, and their resultant impact onaerodynamic forces and moments have been observedto be dependent on a number of factors including pitchrate, angle-of-attack range, and motion time history.Figure 3 shows a comparison of normal forcecoefficient for oscillations about various mean anglesof attack from 22° to 37°. These data show that as themodel was oscillated at high angles of attack, theovershoot and hysteresis loop was enlarged. Theincrease in overshoot due to the increase in meanangle of attack is due, in part, to the increase of pitchrate which occurred closer to the maximum static CN asthe mean angle of attack was increased. The hysteresisloop widened as the angle-of-attack range extended tohigher values, due to more complete flow separationon the upper surface of the wing and the lagsassociated with the separation and reattachmentprocess. As oscillation frequency was increased, thehysteresis loops and overshoot also increased due tothe increased lag in vortex breakdown andredevelopment. A summary of the normal forceovershoot for a range of reduced frequencies is shownin figure 4.

Lateral Characteristics- The vortex flow over themodel also dominated the lateral stabilitycharacteristics. Figure 5 shows the burst locations ofthe vortices on the leeward and windward sides of themodel under static conditions. As angle of attack wasincreased over 20°, the windward vortex burst pointprogressed forward onto the wing surface resulting in aloss of vortex lift on the windward side of the wing.This asymmetric loss of lift resulted in destabilizingincrements to Cl as angle of attack increased. At

α=30°, the leeward vortex burst point moved onto thewing and began a progression towards the apex asangle of attack was further increased. Again, thisresulted in an asymmetric loss of vortex lift whichgave a stabilizing increment to Cl. Above α=40°,very little vortex lift remained on either the windwardor leeward sides of the wing and Cl assumed anessentially constant value. Dynamic lateral stabilitycharacteristics obtained by oscillating the model atβ=5° are shown in figure 6. The effect of theoscillatory motion on roll stability was to delay thestatic characteristics which, as previously seen, werelargely determined by the vortex burst pointlocations. Figure 6 shows the effect of frequency onlateral stability at a mean angle of attack of 27° and at5° sideslip. The data show a delay in the reduction ofstability compared to the static data at the mid angle-of-attack range; however, the dynamic data was much

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less stable than the static data at the higher angles ofattack. During the pitch-down phase, the dynamic datashowed little effect of the asymmetric vortexbreakdown progression as was discussed for the staticand pitch-up data. The absence of variation of rollingmoment during the pitch-down cycle was due to thedelay in vortex formation on the upper surface of thewing. During the pitch-down cycle, the vortex burstpoints were much nearer the wing apex than for theupstroke or static cases, resulting in reduced forces andmoments from asymmetric flow conditions. As thepositive pitch rate was increased, the lateral stabilityimproved progressively over the mid angle-of-attackrange and the unstable increment at high angles ofattack decreased.

Flow visualization provided information on the extentof the lag of the vortex burst point locations insideslip. A comparison between figures 7 and 2 showsthe effects of pitch rate on the asymmetric burst pointlocations in sideslip. As would be expected, for bothstatic and dynamic conditions, the windward vortexburst at a point forward of the leeward vortex burstpoint. During oscillatory motion, the lag in thevortex burst point was evident in both windward andleeward vortex systems.

Effect of wing sweep

Modern fighter aircraft geometries consist of highlyswept strakes, forebodies, and leading-edge wingextensions which generate concentrated vortical flowfields similar to the previously discussed 70° deltawing. In addition to the vortical flow fields, theseaircraft can also incorporate moderately swept wingsand tail surfaces which create a hybrid flow systemover the configuration. To further understand thedynamic effects on wings with lower sweep angles,tests were also conducted on flat-plate wings with 45°and 0° leading-edge sweep angles. A photographcomparing the three wings tested was shown in figure1. The two delta wings had the same wing area, and therectangular wing had an aspect ratio of 1.46 whichcorresponds to the aspect ratio for the 70° delta wing.

As sweep angle is reduced, the flow field over the wingbecomes more 2-dimensional in nature. Previousairfoil studies have documented the dynamic stallprocess for 2-dimensional configurations. Thelongitudinal force results for the airfoils showed trendssimilar to those observed for the 70° delta wingincluding large hysteresis loops and overshoots;however the mechanism responsible is quite different.Figure 8 shows a comparison of the two dynamic stallmechanisms. For the 2-D airfoil, as the airfoil flowseparated during pitch-up, a dynamic stall vortexforms near the leading edge. This vortex grows asangle of attack increases, and then separates from theleading edge and convects downstream. A trailing-edge vortex also develops momentarily at high anglesof attack. It is this dynamic stall vortex system whichis responsible for the force overshoots and hysteresisseen for airfoils. This mechanism is contrasted to thatdiscussed previously for the 70° wing. The 2-Ddynamic stall characteristics are dominated by thepresence of vorticity developed by flow separating atthe leading edge and convecting downstream over thesurface of the airfoil. For the 70° delta wing, theprimary mechanism responsible for the dynamic stalleffects was lags in the burst point location of theestablished leading-edge vortex system on the wing.

For 3-dimensional, rectangular wing configurations,reference 15 reported that the inboard section of thewing produced flow which was essentially 2-D innature; however, near the wing tip the dynamic stallvortex interacted with the tip vortex which behavedsimilarly to the observed leading edge vortex systemon the 70° delta wing. The presence of the two typesof flowfields provides a further challenge to theunderstanding of the unsteady aerodynamic effectsassociated with dynamic stall.

Longitudinal Characteristics- A comparison of thestatic normal force results between the three wings isshown in figure 9. As sweep angle was decreased, themaximum normal force coefficient and stall angle ofattack were reduced due to the well known vortex liftwhich occurs on swept wings at high angles of attack.Dynamic normal force characteristics for the 45° and0° sweep configurations ares shown in figures 10 and11. These data show similar qualitative results tothose seen previously for the 70° delta wing. Figure12 shows a comparison of the maximum overshoot innormal force due to dynamic effects for the three flat-plate wings. The rectangular wing exhibited thelargest overshoot and hysteresis effects due to pitchoscillation. This overshoot was due to the delay inflow separation and reattachment on the upper surfaceof the wing and the dynamic stall vortex, similar tothe flow discussed earlier for airfoil dynamic stall.Because of the lower sweep angle, the 45° delta wingdeveloped a less concentrated leading-edge vortexsystem which broke down at substantially lowerangles of attack compared to the 70° wing.Additionally, the 45° sweep angle prevented large 2-Dtype of dynamic vortex lift to develop resulting in ahybrid flow situation somewhere in between theleading-edge dominated flow field of the 70° wing, andthe quasi 2-D flow of the rectangular wing. The resultsclearly show that the dynamic stall effects decreasewith increasing wing leading-edge sweep.

Lateral Characteristics- The effect of wing sweep onlateral stability is illustrated by comparing figures 5,13, and 14. As noted previously, the asymmetric burstof the leading-edge vortices is the primary factor inthe static lateral stability of wings with substantialsweep angle. Figure 13 shows that the 45° swept wingwas fairly insensitive to dynamic effects on lateralstability. Because the moderate sweep angle does notdevelop the concentrated leading-edge vortex system,the asymmetric vortex structure over the wing and thelag in the vortex burst point locations during pitchingoscillations, which were seen to cause the largechanges in lateral stability for the 70° wingconfiguration, did not develop to a significant degreeon the 45° wing. Figure 15 shows a sketch of therectangular wing flow field in sideslip. The windwardedge of the rectangular wing became a highly sweptsurface in sideslip, generating a vortex which passedover the wing. Figure 14 compares the static anddynamic results of the rectangular wing at 5° sideslip.The dynamic data exhibits a very large amount ofscatter denoted by the shaded area. The inconsistencyof the data for this wing is interesting because the dataindicate large deviations from static values during theoscillations. The data trend suggests that underdynamic conditions the windward side of the wingtends to stay attached to higher angles of attack thusproviding increases in stability. As discussedpreviously, a vortex is probably present on thewindward side due to the effective 85° sweep angle of

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the windward tip. These results indicate thesensitivity of the unsteady aerodynamic effects toplanform shape. Realistic airplane configurationswill likely exhibit several of these basic types ofvortical flow fields which can interact to producehighly complex dynamic stall phenomena.

REALISTIC AIRCRAFT CONFIGURATION

Low-speed wind tunnel studies were conducted toexplore the effects of large-amplitude pitchingmotions on the aerodynamic characteristics of amodern fighter airplane configuration. A sketch of theF-18 configuration tested is shown in figure 16. Thetest configuration incorporated a moderately sweptwing with a highly swept leading-edge extension(LEX). All data presented herein will be with theleading-edge flaps deflected to 34° and trailing-edgeflaps set to 0°. Tests were conducted at 0° and 10°sideslip, and with deflections of ailerons andhorizontal tails. The model was mounted on thedynamic test rig through a six-component strain-gagebalance. The dynamic test rig was a computercontrolled, hydraulically-actuated system which wassting-mounted on a C-strut support system in theNASA Langley 12-Foot Low-Speed Wind Tunnel. Asketch of the installation is shown in figure 17. Themounting arrangement rotated the model about thereference center of gravity location of 24%c andprovided an angle-of-attack range from 5° to 75°. Thecombination of model pitch rates and tunnel speedsallowed for a realistic representation of full-scalemaneuvering conditions within the capabilities ofcurrent and projected future high performanceairplanes. Figure 18 compares non-dimensional pitchrate to equivalent full-scale pitch rate for arepresentative fighter airplane.

Force tests were conducted at a dynamic pressure of 4psf resulting in a Reynolds number of 0.4x106 perfoot. Static data were obtained over an angle-of-attackrange of 5° to 75°. Dynamic data were obtained forboth oscillatory and ramp motions; however, thecurrent discussion will focus on constant pitch-rateramp motions. Tests were conducted using a range ofinitial and final angles of attack and pitch rates.Dynamic data were measured at non-dimensional pitch

rates varying from q=0.0087 to 0.0260. Gravity andinertia tare values were determined by measuring theforces and moments during wind-off test runs. Thesedata were subtracted from the wind-on dynamic data.Because of the relatively low test frequencies and smallmodel weight, inertia tare corrections were small.

Data recorded during test runs included a linear variabledifferential transformer reading from the pitch rig todetermine the pitch angle, 6-component force andmoment data from the strain-gage balance, and tunneldynamic pressure. These inputs were obtained at ratesbetween 100 and 200 samples/second per channel,depending on the test condition. All channels wereelectronically filtered by a 100 Hz low-pass filter.Further digital filtering was applied during post-testdata analysis. All moment data were referenced to the24%c location.

Longitudinal Characteristics

Static results- The flow field over the model at highangles of attack was dominated by a strong vortexsystem generated by the LEX. This vortex system

contributed both to increased lift and reduced pitchstability. Development of the LEX vortex systemstarted at very moderate angles of attack. At theseconditions, the vortex system trailed over the wingand passed just outboard of the vertical tails. Withincreasing angle of attack, breakdown of the vorticesprogressed forward such that the burst point movedover the wing and LEX. In addition, a second vortexsystem developed on the forebody which interactedwith the LEX vortices. A detailed discussion of thehigh-α vortical flow phenomena at static conditionsfor this configuration is contained in reference 16.Static longitudinal force and moment data measured inthe current tests are presented in figures 19 and 20.Figure 19 shows a linear variation in lift with angle ofattack up to about 20°. At this point, a reduction inlift-curve slope was caused by progression of the LEXvortex burst point onto the wing, however, liftcontinued to increase until a maximum value wasreached at about α=35°. Figure 20 shows pitchingmoment characteristics. The results indicate nearneutral stability for angles of attack up to 40°. Thischaracteristic was due in large part to the lift producedon the LEX ahead of the moment reference center.Above α=48°, which corresponds to the condition atwhich the vortex burst point reached the LEX apex, thedata show a substantial increase in pitch stability.

Dynamic results- A series of tests using constant pitchrate motions were conducted to investigate the effectsof pitch rate and motion time history, and to determinethe persistence of the unsteady aerodynamic effects.Pitch motions were generated starting from variousangles of attack and at non-dimensional pitch rates

from q=0.0087 to q=0.0260. Tests were conducted atboth positive and negative pitch rates.

The effects of positive pitch rate on longitudinalcharacteristics are shown in figures 21-23. The modelwas pitched at constant pitch rates over an angle-of-attack range from 5° to 75°. Figure 21 shows anincrease in lift coefficient due to pitch rate over theentire range of motion. A maximum increment ofroughly 0.6 in CL was observed for the highest pitchrate case as compared with static values, and themaximum increment occurred beyond the angle ofattack for maximum lift. The figure shows that aspitch rate was increased, the increment in CLincreased. Although the lift enhancement would beexpected to peak at some value of pitch rate, the rangeof rates tested did not determine the pitch rate beyondwhich no further increase in lift occurred. Figure 22shows the effect of pitch rate on the drag polar of themodel. These data show substantial decreases in dragfor a constant lift coefficient at angles of attack belowmaximum lift. It is also noted that beyond maximumlift, in addition to the increased values of liftcoefficient, dynamic effects produced correspondinglyhigh values of drag. Figure 23 shows the effect ofpositive pitch rate on pitching moment. As discussedin reference 12, a number of factors must be consideredwhen analyzing these data including the distributionof induced angle of attack generated over the model dueto the pitch rate, and the flow lag associated first withflow separation and vortex formation at the lowerangles of attack, and then with vortex breakdown atthe higher angles of attack. The data presented infigure 23 show that for all angles of attack up to 50°,the ramp motions produced large nose-down CMincrements compared to the static values. At the lower

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angles of attack, this effect was found to be largely dueto the induced angle-of-attack distribution discussedearlier which tended to increase loads on thehorizontal tails producing nose-down CM increments.At the higher angles of attack, however, the lag in thebreakdown of the LEX vortex system becamedominant. The lag in the LEX generated flow fieldresulted in the nose-up pitching moment incrementsseen above about α=55°. Under static conditions, theLEX vortex burst point reached the apex of the LEX atabout α=48°, reducing lift on the forward region of themodel which resulted in a net pitch-down increment.During pitch-up motion, the established vortex flowremained intact to higher angles of attack andtherefore the sharp pitching moment break exhibitedstatically above α=48° was substantially delayed.

The effects of negative pitch rate on longitudinalcharacteristics while the model was pitched from 75°to 5° angle of attack are shown in figures 24-26. Atα=75°, the flow condition on the model wascharacterized by complete flow separation andbreakdown. During the pitch-down motion, theformation of the LEX vortex system lags incomparison to the static case and the reattachment ofthe wing flow is delayed. This resulted in largeincrements in the forces and moments compared tostatic values. Figure 24 shows the effect of negativepitch rate on lift coefficient. Large decreases in liftcoefficient were evident due to the motion. As pitchrate was increased, the loss in lift increased due to thelags in vortex development and wing flowreattachment. Figure 25 shows the effect of negativepitch rate on the drag polar of the model. The flowcharacteristics discussed above resulted in littleincrease in lift during the pitch-down motion, and invery low L/D values compared to static data. Similartrends were seen in pitching moment characteristicsduring pitch-down motions. Figure 26 shows that thedelay in the vortex formation on the LEX (forward ofthe moment reference center) resulted in a negativeincrement in CM from 60° to 40° angle of attack.Further in the pitch-down cycle, the LEX vortexsystem developed; however, the burst location wasmuch further forward than for the static condition. Theforward location of the LEX vortex system resulted inthe positive increment in CM seen from 40° to 0°angle of attack.

A comparison of effective pitch dampingcharacteristics was made between current results anddata from conventional small-amplitude forcedoscillation tests. Figure 27 summarizes these results.In order to directly compare results from the twodifferent test techniques, a pitch damping derivativewas calculated from the current data by subtractingstatic values from the dynamic results. This differencewas then non-dimensionalized to obtain a pitchdamping derivative. A comparison was made betweenthe two test techniques for both pitch-up and pitch-down motions. Because of the assumed linearity ofpitch damping with pitch rate, inherent in theconventional forced oscillation results, the predictedpitch damping derivative from this test is identical forboth positive and negative pitch rate motions. Thecomparison shows that during pitch-up motions, largeamplitude and conventional results agreed closelybelow α=50°. Above α=50°, the large amplituderesults indicate slightly less damping than predictedby conventional tests. During pitch-down motions,

the current results show large differences in dampingcompared to the forced oscillation results. Thisdifference reflects the importance of motion timehistory on dynamic stall phenomena.

To aid in the interpretation of the force and momentresults, flow visualization was conducted using apulsed laser light sheet system. Figure 28 shows acomparison of static and pitching motions at 30°angle of attack at a station corresponding to thejunction of the wing leading edge and the fuselage.The pictures show that compared to the static case, theupward motion developed a vortex core which wasslightly closer to the model surface (fig 28 b), and theburst point of the LEX vortex was seen to besubstantially more aft. Conversely for the pitch-downcase, the vortex core has burst at this location (fig 28c).

In addition to the lag and overshoot characteristics,two other important attributes of dynamic stall need tobe considered. The first is the previously mentionedimpact of motion time history on the dynamic effects.Because dynamic stall phenomena are essentially flowlag effects, the past time history of the motion canplay an important role in determining thecharacteristics. As an example, figure 29 showsnormal force coefficient for a series of three constantpitch rate motions starting at different angles ofattack. At a given angle of attack, all three dynamicmotions had identical values of pitch rate; however, ascan be seen, substantial differences in forces existedbetween the motions. This is due to the differences inthe initial flow field condition over the model beforethe motion began. These types of effects are notcurrently captured in the aerodynamic mathematicalmodels used in aircraft simulations. The secondimportant attribute of dynamic stall effects is thepersistence of the effects. The identification ofpersistence times is an essential ingredient inpredicting the importance of dynamic stall effects onairplane capabilities. Obviously, in order to have animpact on the flight dynamics of the airplane, theforce and moment increments induced by dynamic stalleffects must persist long enough to integrate intodisplacement of velocity or rotational vectors.Persistence is usually measured in convective timeunits. One convective time unit is defined to be equalto the amount of time it takes for an air particle totravel across the mean aerodynamic chord of the wing.Figure 30 shows the relationship between convectivetime units and actual time for a full-scale F-18airplane. Tests with the current configuration haveshown persistence to be a function of pitch rate,angle-of-attack range, and motion time history.Figure 31 shows the persistence of the lift overshootfor two pitch rates and two final angles of attack. Thedata show that the maximum persistence timesachieved were near 80 convective time units. Thelargest persistence values were obtained when themodel motion was stopped at an angle of attack of 45°.This corresponds to the point at which the largestdynamic increment was seen (figure 21). Preliminaryanalysis (ref 12) which took into account only liftcharacteristics, has shown that persistence valuesobtained in these tests were too short to significantlyenhance maneuver capability of this airplane.

Lateral Characteristics

In addition to longitudinal stability and performanceaspects, the impact of dynamic stall effects on lateral

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stability may also be important. It is well known thatthe forebody and LEX flow fields play a dominant rolein the lateral stability characteristics of modernfighter configurations at high angles of attack;therefore, the vortex lag effects seen previously wouldbe expected to influence lateral stability as well aslongitudinal stability.

Static results- Figure 32 shows the static rollingmoment coefficient for sideslip values of 0° and 10°.Focusing on the 10° sideslip case, the progression ofLEX vortex burst over the wing can be seen. Typicalof swept wing configurations, as angle of attack wasincreased from 5° to 15° an increase in lateral stabilitywas seen. As angle of attack was further increased, thewindward vortex system burst point reached thetrailing edge of the wing and progressed forward,reducing lift on the windward side and therebydecreasing the rolling moment coefficient. Beyond35° angle of attack, the leeward LEX vortex burst pointmoved forward onto the wing reducing lift on theleeward side which resulted in an increase in static rollstability.

Dynamic results- To assess the effects of pitch rate onthe lateral stability characteristics, the model waspitched at constant rates over an angle-of-attack rangefrom 5° to 75° and 75° to 5° at several values ofsideslip. The data exhibited large differences inrolling moment due to pitch rate. The dynamic rollingmoment characteristics reflected the effect of flow lagssimilar to what was discussed earlier for the 70° deltawing. Figure 33 shows the effect of positive pitch rateon lateral stability at 10° sideslip. As previouslydiscussed, during pitch-up motions vortex burstingand wing flow separation is delayed to higher anglesof attack compared to static conditions. At sideslip,this resulted in a delay of windward wing lift losswhich, in turn increased stable dihedral effect(negative Cl), and also extended the conventional lowangle-of-attack range where lateral stability increasedwith increasing angle of attack. As pitch rate wasincreased, the magnitude of the dynamic increments inrolling moment increased.

In addition to lags in the windward vortex, delay of theleeward vortex burst was also evident. This effect canmost easily be seen in the data at the lowest pitch rate,q=0.0087. The data show that over the angle-of-attackrange from 30° to 50° the delay in the burst of theleeward vortex system resulted in a destabilizingrolling moment contribution.

At negative pitch rates (figure 34), the model motionwas initiated at conditions of fully separated flow. Asthe model was pitched towards lower angles of attack,lag in the redevelopment of vortical flow over themodel resulted in large variations in rolling momentcompared to static results. The data show that over amoderate angle-of-attack range from 25° to 40°, themotion produced substantial increments in negativeCl. This result suggests a relatively long delay in thereformation of the leeward vortex system.

Control Characteristics

Adequate control effectiveness must be maintained toachieve effective maneuver performance in the highangle-of-attack flight regime. As an airplane isconducting large-amplitude, high rate maneuvers athigh angles of attack, dynamic stall effects may

impact control effectiveness. Accurate predictions ofthe high-α agility of an airplane require understandingof dynamic stall effects on control characteristics.Tests were conducted to obtain an indication ofdynamic effects on control effectiveness in the pitchand roll axes.

Pitch control power- A comparison of total availablepitch control power between static and dynamicconditions is shown in figure 35. The differencebetween data obtained with full nose-up control (δh=-

24) and that obtained with δh=0 resulted in the ∆CMvalues shown in the figure. Values of ∆CM werecomputed for both static and dynamic test conditions.The results showed that a slight increase in horizontaltail effectiveness was generated due to positive pitchrate at angles of attack below 62°. Part of the observedeffect was due to the improved flow field conditions atthe aft end of the model caused by flow lags previouslydiscussed. Additionally, the change in the horizontaltail deflection slightly altered the pitch dampingcharacteristics, which are an integral part of thedynamic data set shown in the figure.

Roll control power- Figure 36 shows a comparison ofthe static and dynamic rolling moment for a right rollcontrol input (δa=-25°). The static data show a largeroll control moment at angles of attack below 20°.Above α=20°, the roll control power rapidly declineddue to the stall of the wing and breakdown of the LEXvortices above the wing. The dynamic data againshowed the impact of flow lags. During pitch-upmotion, the breakdown of the LEX vortex system overthe wing was delayed. This resulted in a wing flowfield condition corresponding to a lower angle ofattack. Because of the delay in the wing flow fieldbreakdown, aileron effectiveness remained at the low-α control levels to 30° angle of attack beforedeclining. As angle of attack was further increased,the dynamic roll control effectiveness decreased, dueto the vortex and wing flow field breakdown; however,the lags in the breakdown compared to staticconditions resulted in a higher level of available rollcontrol over most of the angle-of-attack range.During pitch down, the model motion was initiated ina region of fully separated flow. The pitch-downmotion delayed the reformation of the flow field overthe wing which resulted in a delayed increase in rollcontrol power as angle of attack was reduced comparedto static data.

Impact On Maneuvering Capability

A preliminary assessment of the potential flightdynamics impact of the unsteady aerodynamic effectsdiscussed above was made by considering twofundamental high-α maneuvers. These maneuvers arenose pointing and velocity vector turning and areillustrated in figure 37 (see reference 2 for additionaldiscussion).

Nose Pointing Capability- A rapid large-amplitudepitch-up maneuver to point the nose of the airplane atan adversary can be very effective in close-in air-to-aircombat. The primary desired airplane characteristicsto effectively perform this type of maneuver are theability to generate large pitch accelerations and rates,and adequate levels of stability at the resulting highangles of attack. In addition, excessive build-up of

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drag is undesirable due to the increased energy lossduring the maneuver. The large-amplitude rampmotions explored in this study are representative of anose-pointing maneuver. As discussed earlier, the datameasured for this type of motion show a large increasein lift due to unsteady effects, and at the higher anglesof attack, a marked increase in drag compared to staticvalues (see figures 22 and 25). Of even greaterimportance perhaps, is the unsteady pitching momentcharacteristics. The data suggests that the effectivepitch rate damping due to unsteady dynamic effects canvary widely from that measured in conventional windtunnel testing involving small-amplitude motions.This result implies that the dynamic effects couldinhibit the ability to generate the high controllablepitch rates and accelerations desired for effective nosepointing, or adversely affect the stability of theconfiguration during air combat maneuvering. Theseeffects may have to be included in simulation studiesfor accurate prediction of these types of motions.

Velocity Vector Turning- As shown in figure 37, theother class of maneuver assessed was velocity vectorturning. The ability to perform rapid, low radius turnsis very advantageous in close-in air-to-air combat.The primary aerodynamic factor in turningperformance is the level of lift that can be generated.Because unsteady aerodynamic effects can producelarge increments in lift over the static values, therehas been much speculation about the benefits inmaneuverability which may be achievable utilizingthese effects. Assuming an increment in lift of 0.5,approximately the maximum measured due to unsteadyeffects, potential turn performance improvements canbe assessed. Figure 38 shows the instantaneous turnrate achievable for a baseline configuration (nounsteady aerodynamic effects) and for a configurationwith an assumed 0.5 increment in CL due to unsteadyaerodynamics for a 32000 lb airplane at 10000 feetaltitude. The upper boundary represents corner speedas dictated by limit load factor. The data show asignificant increase in instantaneous turn rate due tothe unsteady aerodynamic effects, and a reduction incorner speed. These improvements, however, aremeaningful only if the persistence of the dynamic liftis sufficient to allow the forces to integrate intosignificantly improved turning angles compared to thebaseline performance capability. A preliminaryassessment of the required persistence was made usinga very simplified analysis. The experimental dataindicated that the lift bleedoff following motioncessation is approximately exponential. Thus thisbehavior was modeled as an exponential function asdiscussed in reference 12.

Using this model and a simple point mass analysismethod, the effect of persistence on the turnperformance of the airplane was addressed. Figure 39shows the increased turn angle capability due todynamic effects for various persistence times. Theseresults suggest that for the dynamic lift effects tobecome significant in terms of a turning advantage,the required persistence of the unsteady effects wouldprobably need to be in excess of 100 convective timeunits (incremental turn angle > 4°). As discussedearlier, maximum values of only about 80 convectivetime units were seen during the experiments. Thus itappears that the natural persistence of the dynamic liftovershoots is too low to provide significant turnbenefits.

CONCLUDING REMARKS

Design trends for future fighter aircraft indicatecontinued emphasis on maneuver capability in thestall/post-stall high angle-of-attack flow regime.Aggressive exploitation of the high angle-of-attackenvelope will likely encounter large unsteadyaerodynamic effects. These effects need to beidentified and analyzed in order to ensure that adequatepredictions of airplane flight dynamics can beobtained.

This paper has summarized recent research at NASALangley Research Center on the effects of dynamicstall on airplane characteristics at high angles ofattack. Results have shown that fundamental flowfield mechanisms which occur on simple flat-platemodels are similar to those found on a realisticconfiguration. Substantial increments in lift, drag,and pitching moment due to dynamic effects weremeasured for the rapid large-amplitude pitchingmotions. Pitch rate was the dominant influence,however motion time history was also important.

The primary consideration of this research is theimpact of dynamic effects on airplane flight dynamics.Natural persistence of the dynamic lift effects wereseen to be too short to significantly affect turnperformance; however, longitudinal and lateralstability were significantly impacted by dynamiceffects. This latter characteristic may result intransient performance differences not currentlypredicted by conventional test techniques. In additionto force generation and stability impacts, dynamicstall effects also were apparent in controleffectiveness in both pitch and roll axes.

Additional research is needed to determine the impactof the dynamic stall effects on both stability andcontrol power and to explore flow control methods forexploiting these phenomena. Only by understandingthe complex flow phenomena at high angles of attack,and their effects on flight dynamics during high ratemaneuvering, can new aircraft be confidently designedto maneuver effectively and aggressively exploit thisexpanded maneuvering envelope.

REFERENCES

1 . Lang, J.D.; and Francis, M.S.: UnsteadyAerodynamics and Dynamic Aircraft Maneuverability.AGARD CP-386, May 1985.

2 . Nguyen, L.T.: Flight Dynamics Research forHighly Agile Aircraft. SAE paper 892235, September1989.

3 . Lorber, P.F.; and Carta, F.O.: Airfoil DynamicStall at Constant Pitch Rate and High ReynoldsNumber. AIAA 87-1329, June 1987.

4 . Galbraith, R.A.; Niven, A.J.; and Seto, L.Y.: Onthe Duration of Low Speed Dynamic Stall. ICAS-86-2.4.3.

5 . Francis, M.S. et al: An Investigation of AirfoilDynamic Stall Overshoot of Static AirfoilCharacteristics. AIAA-85-1773 August 1985.

6 . Wolffelt, K.W.: Investigation on the Movementof Vortex Burst Position With Dynamically ChangingAngle of Attack for a Schematic Deltawing in a

8

Watertunnel With Correlation to Similar Studies in aWind Tunnel. AGARD CP-413, October 1986.

7 . Lambourne, N.C. et al: The Behavior of theLeading-Edge Vortices Over a Delta Wing Following aSudden Change of Incidence. ARC-R/M-3645 1970.

8 . Gad-el-Hak, M.; and Ho, C.: The Pitching DeltaWing. AIAA Journal Vol. 23, No. 11, November1985.

9 . Reynolds, G.A.; and Abtahi, A.A.: Instabilitiesin Leading-Edge Vortex Development. AIAA-87-2424, August 1987.

10. Brandon, J.M.; and Shah, G.H.: Effect of LargeAmplitude Pitching Motions on the UnsteadyAerodynamic Characteristics of Flat-Plate Wings.AIAA-88-4331, August 1988.

11. David, M.: Water Tunnel Investigation of theVortex Dynamics of Periodically Pitched Wings. MSThesis, Air Force Inst. Technol., December 1988.(Available from DTIC as AD-A206 359).

12. Brandon, J.M.; and Shah, G.H.: UnsteadyAerodynamic Characteristics of a Fighter ModelUndergoing Large-Amplitude Pitching Motions atHigh Angles of Attack. AIAA-90-0309, January,1990.

13. Cunningham, A.M.; and den Boer, R.G.:Harmonic Analysis of Force and Pressure Data ResultsFor An Oscillating Straked Wing at High Angles.AIAA-87-2494, August, 1987.

14. Cunningham, A.M.; and Bushlow, T.: Steady AndUnsteady Force Testing Of Fighter Aircraft Models InA Water Tunnel. AIAA-90-2815, August, 1990.

15. Robinson, M.E.; and Wissler, J.B.: Pitch Rateand Reynolds Number Effects On A PitchingRectangular Wing. AIAA-88-2577, June, 1988.

16. Erickson, G.E.: Water Tunnel Flow Visualizationand Wind Tunnel Data Analysis of the F/A-18. NASACR-165859, May 1982.

Figure 1 - Photograph of flat-plate models

0

20

40

60

80

100

120

1400 10 20 30 40 50, degα

Burst pointlocation(% X/C)

Apex

Trailing-edge

kStatic0.04260.1346

Figure 2 - Effect of pitch rate on burst point location, 70° delta wing

2.0

1.5

1.0

.5

0 10 20 30 40 50 60, degα

C N

Static22273237

oα

Figure 3 - Effect of mean angle of attack on dynamic force characteristics, 70° wing,

k=0.0376

25

20

15

10

5

0 .01 .02 .03 .04Reduced frequency, k

CN%overshoot

9

Figure 4 - Effect of pitch rate on normal force overshoot, 70° wing

0

10

20

30

40

50

60

70

80

90

10010 20 30 40 50

, degα

Vortex burstlocation,

% c R

Apex

T.E.

Leeward vortexWindward vortex

Figure 5 - Vortex burst location, 70° wing, β=5°

0 10 20 30 40 50 60, degα

.02

.01

0

-.01

-.02

-.03

-.04

kStatic

0.0226

C

Figure 6 - Effect of pitch rate on lateral stability,70° wing, α0=27°, β=5°

0

10

20

30

40

50

60

70

80

90

10010 20 30 40 50

, degα

Vortex burstlocation,

% c R

Apex

T.E.

Leeward vortexWindward vortex

Figure 7 - Dynamic vortex burst location atsideslip, 70° wing, α0=27°, β=5°, k=0.0426

<figure missing>

Figure 8 - Dynamic stall flow field mechanisms

, degα

CN

2.00

1.75

1.50

1.25

1.00

.75

.50

.25

0 10 20 30 40 50 60

SWEEP70450

oo

o

Figure 9 - Comparison of static normal forcecharacteristics for flat-plate wings

1.5

1.0

.5

0 10 20 30 40 50 60, degα

CN

Static0.00900.01360.0227

k

Figure 10 - Effect of reduced frequency on normalforce coefficient, 45° wing

, degα

CN

1.50

1.00

.50

0

-.50-20 -10 0 10 20 30 40

Static0.00750.02260.0376

k

Figure 11 - Effect of reduced frequency on normalforce coefficient, 0° wing

110

100

90

80

70

60

50

40

30

20

10

0 .01 .02 .03 .04k

0°45°70°

Wing sweep

%CNmax

Figure 12 - Comparison of maximum normal forceovershoot characteristics

10

.01

0

-.01

-.02

-.03

, degα0 10 20 30 40 50 60

C

StaticDynamic

Figure 13 - Effect of pitch oscillation on lateralstability, 45° wing, α0=15°, β=5°, k=0.0227

0 10 20 30 40, degα

.2

.1

0

-.1

-.2

-.3

-.4

-.5

C

Static data

Downstroke

Upstroke

Figure 14 - Effect of pitch oscillation on lateralstability, 0° wing, α0=12°, β=5°, k=0.0376

βV 8

Tip vortices

Figure 15 - Flow field in sideslip

Figure 16 - Sketch of F-18 configuration

Hydraulicactuator

Pivot

Balancemoment

referencecenter

Sting

Figure 17 - Model installation

11

150

100

50

0 .01 .02 .03 .04

Test range

Full-scalepitch rate,deg/sec

qc2V

q =^

Full-scale velocity

600 ft/sec 400 ft/sec

200 ft/sec

Figure 18 - Non-dimensional pitch rate

2.0

1.5

1.0

.5

0 10 20 30 40 50 60 70 80, degα

CL

Figure 19 - Static lift characteristics, F-18configuration

.2

.1

0

-.1

-.2

-.3

-.4

-.5

-.60 10 20 30 40 50 60 70 80

, degα

CM

Figure 20 - Static pitching moment characteristics,F-18 configuration

0

2.0

1.5

1.0

.5

C L

Static0.00870.01300.01730.0260

q̂

2.5

10 20 30 40 50 60 70 80, degα

Figure 21 - Effect of positive pitch rate on liftcoefficient

0

2.0

1.5

1.0

.5

C L

.5 1.0 1.5 2.0 2.5 3.0CD

Static0.00870.01300.01730.0260

q̂

2.5

Figure 22 - Effect of positive pitch rate on dragpolar

.2

.1

0

-.1

-.2

-.3

-.4

-.50 10 20 30 40 50 60 70 80

, degα

CM

Static0.00870.01300.01730.0260

q̂

Figure 23 - Effect of positive pitch rate on pitchingmoment coefficient

12

0 10 20 30 40 50 60 70 80, degα

Static-0.0087-0.0130-0.0173

q̂2.0

1.5

1.0

.5

C L

Figure 24 - Effect of negative pitch rate on liftcoefficient

0

Static-0.0087-0.0130-0.0173

q̂

2.0

1.5

1.0

.5

C L

.5 1.0 1.5 2.0 2.5 3.0CD

Figure 25 - Effect of negative pitch rate on dragpolar

.2

.1

0

-.1

-.2

-.3

-.4

-.5

-.60 10 20 30 40 50 60 70 80

, degα

CM

Static-0.0087-0.0130-0.0173

q̂

Figure 26 - Effect of negative pitch rate on pitchingmoment coefficient

20

10

0

-10

-200 10 20 30 40 50 60 70 80, degα

Pitchdamping

Forced oscillation^

^q = 0.0130

q = -0.0130

Figure 27 - Comparison of predicted pitch dampingbetween conventional results and large-amplitude data

Figure 28 - LEX vortex flow visualization, α=20°(a) Static

Figure 28 - Continued

(b) q=0.0260

13

Figure 28 - Concluded

(c) q=-0.0260

5

4

3

2

1

0

-1

CN

-10 0 10 20 30 40 50 60 70 80, degα

Static0° - 75°20° - 75°40° - 75°

Dynamic - rangeα

Figure 29 - Effect of motion time history

12

10

8

6

4

2

0 50 100 150 200Convective time units

Time,sec

Full-scale velocity200 ft/sec

400 ft/sec

600 ft/sec

Figure 30 - Convective time relationship

100

80

60

40

20

0

Convectivetime units

5 - 75 5 - 45, degα

0.0087

0.0173

q̂

Figure 31 - Persistence for pitch-up motions

.02

.01

0

-.01

-.02

-.03

-.04

-.05

-.060 10 20 30 40 50 60 70 80

, degα

C

= 0°β

= 10°β

Figure 32 - Static rolling moment characteristics

0

-.01

-.02

-.03

-.04

-.05

-.06

-.070 10 20 30 40 50 60 70 80

, degα

C

Static0.00870.01730.0260

q̂

Figure 33 - Effect of positive pitch rate on lateralstability, β=10°

-.01

-.02

-.03

-.04

-.05

-.06

-.070 10 20 30 40 50 60 70 80

, degα

CStatic0.00870.01730.0260

q̂

Figure 34 - Effect of negative pitch rate on lateralstability, β=10°

14

= -24

.5

.4

.3

.2

.1

0

-.10 10 20 30 40 50 60 70 80

, degα

CM∆δh

Static0.0173

q̂

Figure 35 - Nose-up control power

0 10 20 30 40 50 60 70 80, degα

C

.05

.04

.03

.02

.01

0

Figure 36 - Roll control characteristics, δa=-25°

Nose pointing

Velocity vector turning

Figure 37 - Fundamental high-α maneuvers

30

20

10

0 200 400 600 800V, ft/sec

Turn rate,deg/sec

BasicDynamic

Figure 38 - Dynamic effects on instantaneous turnrate

10

9

8

7

6

5

4

3

2

1

0 5 10 15 20t, sec

Increasedturn angle

2040100200

8

t c

Figure 39 - Effect of dynamic lift persistence onturn performance