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Unsteady Aerodynamic Characteristics Of A Fighter Model

Undergoing Large-Amplitude Pitching Motions- At

High Angles Of Attack

Jay M. Brandon & Gautam H. Shah NASA Langley Research Center

Hampton, VA 23665-5225

28th Aerospace Sciences Meeting January 8-1 1, 19901Ren0, Nevada

For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics 370 L'Enfant Promenade, S.W., Washington, D.C. 20024

UNSTEADY AERODYNAMIC CHARACTERISTICS OF A FIGHTER MODEL UNDERGOING LARGE-AMPLITUDE PITCHING MOTIONS

AT HIGH ANGLES OF ATTACK

Jay M. Brandon and Gautam H. Shah NASA Langley Research Center #yd 42 6~ 32

ABSTRACT

A low-speed wind tunnel study was conducted to explore the effects of large- amplitude pitching motions on the aerodynamic characteristics of a modem fighter airplane configuration. Tests were conducted using a computer- controlled hydraulically-driven dynamic model support system. Data for oscillatory motions were obtained for comparison with results previously measured for a flat-plate delta wing. Additionally, ramp motions over an angle of attack range from 0" to 75" were used to investigate the effects of reduced pitch rate and motion time history, and to determine the persistence of unsteady effects. Force and moment data were obtained using a 6-component internal strain-gage balance. To aid in the interpretation of the results, flow visualization was conducted during dynamic test conditions using smoke injected near the apex of the wing leading-edge extension.

Results of this study indicate that lags in the vortical flow development and breakdown promoted large force overshoots and hysteresis. The aerodynamic characteristics were a strong function of pitch rate and angle of attack. Large effects of motion time history were also observed. Persistence of the dynamic effects was investigated and found to be highly dependent on the motion variables. Maximum persistence values of approximately 50 convective time units were measured. A preliminary evaluation of the impact of dynamic effects on maneuver capability

indicates very minor impact on flight path turn performance, but a potentially large effect on pitch rate capability.

NOMENCLATURE

mean aerodynamic chord, ft lift coefficient pitching moment coefficient normal force coefficient frequency, Hz

reduced frequency, v

pitch rate, radlsec

non-dimensional pitch rate,= 2v

time, sec convective time unit free stream velocity, ftlsec angle of attack. deg induced angle of attack, deg increment in lift coefficient due

to dynamic effects leading-edge extension

INTRODUCTION

At high angles of attack, unsteady aerodynamic effects may have a major impact on the maneuverability and controllability of an airplane. With current emphasis on aggressive maneuvering capability near or beyond the stall angle of attack for future airplanes. research into the effects of large-amplitude unsteady motions at high angles of attack on stability, control and performance is needed. Recent research (ref 1) has discussed the potential tactical benefits of unsteady

This paper is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

aerodynamic characteristics; however, little research has been conducted on realistic three-dimensional configurations to quantify unsteady effects or develop an understanding of the flow mechanisms involved. Additionally, the issues of impact on airplane flight dynamics and practical means of utilizing these unsteady effects have to be addressed (ref 2).

A significant amount of research has been conducted in the area of unsteady aerodynamics at high angles of attack. However, most of the work to date has been on 2-D airfoils. These results show substantial force overshoots for pitching and plunging airfoils (refs 3-5). Current design trends for advanced airplane configurations show a clear need for extending the research to 3-D airplane configurations. Wind tunnel experiments have been conducted using delta wings undergoing pitching and plunging oscillations to determine vortical flow characteristics during large-amplitude motions (refs 6-8). The primary focus of the studies was the dynamics of the leading-edge vortex while the wing was undergoing motion. Under these conditions, the vortex burst point location was observed to lag the static location during pitch-up and pitch-down motions. Water tunnel results (ref 9) have shown vortex lag time of up to 30 convective times units (time required for an air particle to travel across the wing), which is a much slower response than that normally associated with 2-TI unsteady aerodynamic phc:lornena. More recently, research was conducted on a series of wings of different sweep angles which showed that the magnitude of the unsteady aerodynamic effects decreased as the sweep angle increased (ref 10).

This paper presents results of low-speed wind-tunnel tests of a model of a modern fighter configuration undergoing large- amplitude pitching motions for an angle of attack range of 0 to 75 degrees. Various types of motion were tested including harmonic oscillations and

constant pitch rate ramps. The data obtained consisted of aerodynamic forces and moments acting on the model during and subsequent to the motion. Persistence of the unsteady effects was measured. and a preliminary analysis was conducted to assess the potential impact of the dynamic characteristics on the maneuver performance of a full-scale airplane.

MODELS AND TEST APPARATUS

A sketch of the model used for the current study is shown in figure 1. The test configuration incorporated a moderately swept wing with a highly swept leading-edge extension (LEX). All data presented herein will be with all control surfaces set to 0" except the leading-edge flaps which were deflected 34". The model was mounted to the dynamic test rig through a six- component strain-gage balance. The dynamic test rig was a computer controlled, hydraulically-actuated system which was sting-mounted on a C- strut support system in the NASA Langley 1 2-Foot Low-Speed Tunnel. A sketch of the installation is shown in figure 2. The mounting arrangement rotated the model about a point 3.5 inches below the reference cg location of 24%c and provided an angle of attack range from -10" to 80". The maximum capability of the test rig was 2290 degrees/second2 pitch acceleration, and 260 degreeslsecond pitch rate. In addition to harmonic oscillations, the apparatus could produce nearly constant-rate ramp model motions through a large range of angle of attack. The combination of model pitch rates and tunnel speeds allowed for a realistic representation of full-scale maneuvering conditions within the capabilities of current and projected future high performance airplanes.

Force tests were conducted at a dynamic pressure of 4 psf resulting in a Reynolds number of 0 . 4 ~ 1 0 ~ per foot. Static data were obtained over an angle of attack

range of -10" to 80". Dynamic tests were conducted using various initial and final angles of attack and pitch rates. Oscillatory data were obtained at a reduced frequency of k=0.02, and ramp data were measured at non-dimensional

pitch rates varying from :=0.0061 to 0.0182. Gravity and inertia tare values were determined by measuring the forces and moments during wind-off test runs. These data were subtracted from the wind-on dynamic data. Because of the relatively low test frequencies and small model weight, inertia tare corrections were small.

Data recorded during test runs included a linear velocityldisplacement transducer reading from the pitch rig to determine the pitch angle. 6-component force and moment data from the strain- gage balance, and tunnel dynamic pressure. These inputs were obtained at rates between 125 and 250 sampleslsecond per channel, depending on the test condition. All channels were electronically filtered by a 100 Hz low- pass filter. Further digital filtering was applied during post-test data analysis. All moment data were referenced to the 24%; location.

Flow visualization data were obtained by injecting smoke near the apex of the LEX. The smoke was illuminated by a helium- neon laser and the results were recorded with videotape and still photographs. The laser was pulsed at selected angles of attack and pitch motions to illuminate the LEX flow field.

RESULTS AND DISCUSSION - Sta t ic Resul ts

The flow field over the model at high angles of attack was dominated by a strong vortex system generated by the LEX. This vortex system contributed both to increased lift and reduced pitch stability. Development of the LEX vortex system started at very moderate angles of attack. At these conditions, the vortex

system trailed over the wing and passed just outboard of the vertical tails. With increasing angle of attack, breakdown of the vortices progressed forward such that the burst point moved over the wing and LEX. In addition, a second vortex system developed on the forebody which interacted with the LEX vortices. A detailed discussion of the high-a vortical flow phenomena at static conditions for this configuration is contained in reference 11. Static longitudinal force and moment data measured in the current tests are presented in figures 3 and 4. The lift results show a fairly linear variation with angle of attack up to about 20'. At this point, a reduction in lift- curve slope was caused by progression of the LEX vortex burst point onto the wing. However, lift continued to increase until a maximum value was reached at about a=3S0. Figure 4 shows pitching moment characteristics. The results indicate near neutral stability for angles of attack up to 40". This characteristic was due in large part to the lift produced on the LEX ahead of the moment reference center. Above a=4g0, which corresponds to the condition at which the vortex burst point reached the LEX apex, the data show a substantial increase in pitch stability.

Oscillatory Motion Results

Past studies have documented the hysteresis and overshoot characteristics of the aerodynamics of flat-plate delta wings undergoing large-amplitude oscillations (refs 6-10). Results of these studies showed that lags in the leading- edge vortex system development and breakdown contributed to significant force and moment increments during the large-amplitude model motions at high angles of attack. The LEX vortex system on the current model was very similar to the leading-edge vortex system on highly swept delta wings. However, the overall flowfield over the subject model was substantially more complex, encompassing not only the LEX vortex system, but also the forebody, wing, and tail flowfields and their interactions. As

a first step in investigating the behavior of this complex flowfield under dynamic conditions, oscillatory tests were conducted so that the results could be compared to existing flat-plate delta wing data. The testing consisted of constant frequency harmonic oscillations about a mean angle of attack of 32" at various amplitudes.

Figure 5 shows the effect of oscillation amplitude on lift coefficient. The data exhibit large hysteresis loops, particularly for the larger amplitude cases. Flow visualization tests suggest that a key mechanism responsible for these characteristics was the lag in breakdown and reformation of the LEX vortices during an oscillation cycle. These results were obtained using a laser light sheet which was pulsed to illuminate a cross-section of the smoke flow above the model at a given angle of attack. Figure 6 presents a comparison of flow above the model taken at a=27"

during an upward cycle (&0) and a downward cycle (Ck0). The photographs show the relative size of the vortex during model motion. During pitch-up, the vortex appears to be smaller indicating a more stable vortex. This effect was a result of the lag in breakdown of the vortex. During pitch- up, bursting occurred at a point further downstream than would occur statically, resulting in a vortex system which is equivalent to a static system at a reduced angle of attack. During pitch-down, the vorte.: s:re is larger indicating a less stabie \ X ?ex. Again, due to vortex flow laf 1.9c dynamic vortex system can be v - XI as equivalent to the static system at angle of attack earlier in the cycle which, for pitch-down motion, is a higher value than that for the dynamic case. These characteristics are qualitatively very similar to the behavior of the leading-edge vortices observed in earlier studies of highly-swept delta wings. Returning to figure 5 , the results show that increasing the amplitude of the oscillation substantially enlarges the hysteresis loop. Two factors are believed

to be primarily responsible for this characteristic. First, the larger amplitude resulted in higher pitch rates at a given angle of attack. Secondly, the larger amplitude oscillation encompassed a broader range of flow conditions thus tending to promote larger flow lags associated with the separationlreattachment and vortex breakdownlreformation processes. Again, these characteristics are generally very similar to those observed previously in delta-wing studies.

Because the dynamic effects altered the force distribution on the model due to flow lags in both the LEX vortex flow development and the wing flow separationlreattachment, significant dynamic effects are also evident in pitching moment. Figure 7 shows the effect of oscillation amplitude on pitching moment characteristics. For the smallest oscillation amplitude of 10". the data show a fairly conventional loop with the largest nose-down CM occurring

during the positive dl portion of the cycle and the largest nose-up CM during the

negative 01 portion. It is likely that a primary mechanism involved in this case is the conventional pitch damping provided by the horizontal tails. The larger amplitude data (10" and 20°), however, exhibit some highly nonlinear behavior that indicates that the dynamic vortical flow effects discussed earlier were dominant. For these cases, for angles of attack below about 48". a nose- down pitching moment was generated during the upward cycle (&0) because the LEX vortex burst point was aft of the location for static conditions. Conversely, during pitch-down, the lag in the flow delayed the loading of the aft portions of the model which produced a nose-up increment compared to static results. Above a=48O, which is approximately where the LEX vortex breakdown reached the LEX apex under static conditions and resulted in a strong stable break in CM (figure 4). the data show an apparent reversal in characteristics. During the upward cycle

( 6 0 ) . the delay in the vortex burst reaching the apex produced an incremental nose-up moment whereas during the downward cycle (&co), the delay in reformation of the LEX vortex at the apex resulted in an incremental nose-down moment. The consequence of this phenomenon is a clockwise hysteresis loop in contrast to the counterclockwise loop exhibited below a=48".

Ramp Motion Results

A series of tests using ramp input motions were conducted to investigate the effects of pitch rate and motion time history, and to determine the persistence of the unsteady aerodynamic effects. Pitch motions were generated starting from various angles of attack and at non- dimensional pitch rates up to

approximately c=0.0226. Most testing

was conducted at a pitch rate of <=0.0182 which corresponds to a pitch rate of approximately 40 deglsec for the full- scale airplane at an altitude of 20000 feet and an airspeed of 125 kts. Tests were conducted with both positive and negative pitch rates.

Effect of p i u L m e

To assess the effects of pitch rate on the unsteady aerodynamics, the model was pitched at constant rates over an angle of attack range of 0" to 75" and also from 75" to 0". The effect of positive pitch race on lift coefficient is shown in figure 8. The data show an increase in lift coefficient due to pitch rate over the entire range of motion. A maximum increment of roughly 0.6 was observed for the higher pitch rate case as compared to the static values. Additionally, as pitch rate was increased. the increment in CL increased. Although the lift enhancement would be expected to peak at some value of pitch rate, the range of rates tested did not allow determination of the pitch rate beyond which no further increase in lift

would occur. Figure 9 shows the effect of positive pitch rate on pitching moment. A number of factors must be considered in analyzing these data. The first is the angle of attack induced along the length of the model by the constant pitch rate.

For the higher pitch rate case, 6=0.0 182, a maximum Aai n 4" was generated at the horizontal tail and Aai n -6" was induced at the nose. The second primary factor is the flow lag associated first with flow separation and vortex formation at the lower angles of attack, and then with vortex breakdown at the higher angles of attack. The data presented in figure 9 show that for angles of attack up to 50". the ramp motions produced large nose- down CM increments compared to the static values. At the lower angles of attack, this effect is believed to be due primarily to the induced angle of attack distribution discussed earlier which tended to increase the loads at the aft end of the model while reducing the loads forward of the moment center. At the higher angles of attack, however, the lag in the breakdown of the LEX vortex system became dominant. In particular above about a=55", the dynamic data show a nose-up increment compared to the static data. Under static conditions, the LEX vortex burst point reached the apex of the LEX at about a=48", reducing lift on the foreword region of the model which resulted in a net pitch-down increment. During the pitch-up motion, the vortex flow remained established to higher angles of attack before bursting reaches the LEX apex and therefore the sharp pitching moment break exhibited statically above 48" angle of attack was substantially delayed.

Figure 10 shows the effect of negative pitch rate on lift while the model was pitched from 75" to 0" angle of attack. At a=75", the flow condition on the model was characterized by complete flow separation and breakdown. During the pitch-down motion, flow lags in the formation of the LEX vortex, and in the reattachment of the wing flow resulted in the large decreases in lift compared to

static values as shown in the figure. At

the faster pitch rate of i=0.0182, the flow was not completely reestablished until the motion reached a=OO. At the slower pitch rate, the flow lag was much less such that by a=30° the level of lift was very close to that corresponding to the static conditions. The effect of negative pitch rates on pitching moment is presented in figure 11. This data show similar trends caused by flow lags. During the initial pitch down, a negative increment in CM was produced compared to the static data due to the delay in the development of the LEX vortex system and hence the development of lift forward of the moment reference center. Further in the pitch-down cycle, the LEX vortex system developed. However the burst location was much further forward than the static condition, and the wing flowfield remained stalled. This condition resulted in the positive increment in CM observed at the mid angle of attack range for the high pitch rate case. The effect of pitch rate on the delay of flowfield establishment is very

A

evident in comparing the q=-0.0061 data n

with the q=-0.0182 data. At the lower pitch rate, the pitching moment characteristics converged to the static values by approximately a=30° in contrast to nearly a=OO at the higher rate.

To obtain some insight into the performance implications of unsteady aerodynamic effects during large- amplitudc maneuvers, drag polar plots are shown in figure 12 for static and both pitck-up and pitch-down conditions. These data show that during the pitch-up maneuver, equivalent values of lift may be obtained with significantly lower drag values than for the static case. It is also noted that at the conditions where the highest values of dynamic lift occur, correspondingly high values of drag are produced. An interesting characteristic of the pitch-down motion is the relative insensitivity of lift as angle of attack is reduced (see figure 10). Thus, as indicated in figure 12. the pitch-down

motion is essentially modulating drag while maintaining approximately the same level of lift. Further research involving piloted simulation studies are planned to determine if these characteristics can be exploited for maneuvering advantages.

Effect of motion time history

The effects of motion time history were investigated using ramp model motions beginning and ending at various angles of attack. This approach allows comparison of aerodynamic characteristics at equivalent angles of attack and pitch rate with the only difference being the initial angle of attack.

The effects of motion history on lift, normal force, and pitching moment during constant rate pitch-up ramp motions are shown in figures 13-15. The data were obtained by pitching the model from an initial angle of attack of 0". 20". or 40" to a maximum of 75" at a pitch rate

of c=0.0182. Figure 13 shows the measured lift characteristics for the three conditions. The data indicate that the initial flow field is important in determining the dynamic characteristics. This effect can be seen more clearly in the normal force data shown in figure 14. Motion initiated at each of the angles of attack produced the characteristic overshoots in the force data as discussed earlier. Comparing the three sets of data however, reveals the additional effect of motion history. For the case starting at a=OO, the flow at the initiation of the motion was attached. As the angle of attack was increased, lags in the flow development occurred, as previously discussed, creating a LEX vortex flowfield which was equivalent to a lower angle of attack static condition. As the angle of attack was further increased, the vortex system began to break down. Again, flow lags delayed the breakdown and resulted in higher values of normal force than for the static case. For model motions initiated at 20" angle of attack, the LEX vortex flowfield was fully

developed. As the model pitched up, the delay in breakdown of these vortices were primarily responsible for the increase in forces. Note that from an initial angle of attack of 20". the forces generated at higher angles of attack were greater than those resulting from the motion initiated at a=OO. This result demonstrates the importance of the lag in vortex formation as well as the lag in breakdown in determining the dynamic effects. For model motions initiated at 40" angle of attack, the wings were stalled, and the LEX vortex burst point was close to the LEX apex. Starting the motion from this condition, the vortex burst point quickly moved to the LEX apex, and therefore did not generate as large an increment in normal force as the condition starting at a=20°. Pitching moment data, figure 15, show similar results. At the very high angles of attack, the data from an initial angle of attack of 20" show a much larger nose-up increment relative to the static data than the other two cases. As discussed earlier, this effect was caused by the greater lag in vortex breakdown that occurred because of the stronger initial LEX vortex system when the motion started at a=20°. It is important to note that for identical instantaneous conditions of angle of attack and pitch rate, markedly different values of pitching moment were obtained depending on the previous time history of the motion.

Similar data for pitch-down motions are shown in figures 16-18. The data were obtained at initial angles of attack of 75" and 50". Figures 16 and 17 show the lift and normal force characteristics, respectively. This data again show the importance of the initial flowfield on the overall aerodynamics during a large- amplitude pitch maneuver. The lag in the flowfield development resulted in considerably reduced values of CL and CN for the maneuver initiated at a=75" due to the initial extreme flow breakdown condition. Pitching moment data, figure 18, show the characteristic

lag in the flow development during the pitch-down motion. The results for the 75" to 0" angle of attack case show initially the lag in the formation of the LEX vortices which resulted in a nose- down increment compared to static data. At angles of attack below approximately 35". the lag in the vortex burst point position caused the center of pressure to be located further forward than the static case, and produced a nose-up increment in pitching moment. Initiating the pitch- down motion from 50" angle of attack resulted in a positive increment in CM over the entire range of motion. This result again illustrates the importance of motion time history on the transient aerodynamic forces and moments. This dependence of the dynamic effects on initial flowfield condition greatly increases the difficulty in formulating accurate mathematical aerodynamic models for use in aircraft motion simulation studies.

Of primary importance in the evaluation of the impact of unsteady aerodynamic effects is the persistence of the phenomena. For convenience, the measure of persistence was chosen to be non-dimensionalized in terms of convective time units, where one convective time unit corresponds to the time required for an air particle in the free-stream to travel a distance equal to the mean aerodynamic chord of the model.

Ramp motions were used to study the persistence of the dynamic effects. The persistence time was defined as the time increment from cessation of the model motion to when the measured lift reached its steady-state value.

Time history data of two representative persistence runs are shown in figures 19 and 20. Figure 19 shows results for a ramp from 0" to 75". and figure 20 shows similar data for a ramp from 0" to 50". Data from figure 19 show that the motion

stopped well beyond the peak lift value. The dynamic lift increment remaining when the motion stopped dissipated in about 40 convective time units. Figure 20 represents a case where the model motion stopped closer to the maximum lift angle of attack. This resulted in a larger dynamic CL increment which dissipated in approximately 50 convective time units. Persistence was found to be highly dependent on a number of motion variables including pitch rate and initial and final angles of attack. Figure 21 shows the effect of pitch rate and final angle of attack on persistence for pitching motions initiated at 0' angle of attack. The data indicate that an increase in pitch rate enhanced persistence. Also, stopping the motion at a=50° resulted in greater persistence compared to going to a=7S0. Similar data are shown for negative pitch rates in figure 22. The results indicate that initiating the maneuver at a lower, less separated flow condition resulted in greater persistence.

Assessment of Maneuvering Impact

A preliminary assessment of the potential flight dynamics impact of the unsteady aerodynamic effects discussed above was made by considering two fundamental high-a maneuvers. These maneuvers are nose pointing and velocity vector turning and are illustrated in figure 23 (see reference 2 for additional discussion).

A rapid large-amplitude pitch-up maneuver to point the nose of the airplane at an adversary can be very effective in close-in air-to-air combat. The primary desired airplane characteristics to effectively perform this type of maneuver are the ability to generate large pitch accelerations and rates, and adequate levels of stability at the resulting high angles of attack. In addition, excessive build-up of drag is undesirable due to the increased energy

loss during the maneuver. The large- amplitude ramp motions explored in this study are representative of a nose- pointing maneuver. As discussed earlier, the data measured for this type of motion show a large increase in lift due to unsteady effects, and at the higher angles of attack, a marked increase in drag compared to static values (see figures 8 and 12). Of even greater importance perhaps, is the unsteady pitching moment characteristics. Figure 24 shows a comparison of pitching moment measured during a ramp motion from 0" to 75" angle of attack with predicted results based on conventional small-amplitude forced oscillation tests of the same configuration. The data suggests that the effective pitch rate damping due to unsteady dynamic effects can be much greater than that measured in conventional wind tunnel testing involving small-amplitude motions. This result implies that the dynamic effects could inhibit the ability to generate the high pitch rates and accelerations desired for effective nose pointing, and that these effects should be included in simulation studies for accurate prediction of these types of motions.

Velocitv Vector Turning

As shown in figure 23, the other class of maneuver assessed was velocity vector turning. The ability to perform rapid, low radius turns is very advantageous in close-in air-to-air combat. The primary aerodynamic factor in turning performance is the level of lift that can be generated. Because unsteady aerodynamic effects can produce large increments in lift over the static values, there has been much speculation about the benefits in maneuverability which may be achievable utilizing these effects. Assuming an increment in lift of 0.5, which about the maximum measured due to unsteady effects during the current study, potential turn performance improvements can be assessed. Figure 25

shows the instantaneous turn rate achievable for a baseline configuration (no unsteady aerodynamic effects) and for a configuration with an assumed 0.5 increment in CL due to unsteady aerodynamics for a 32000 lb airplane at 10000 feet altitude. The upper boundary represents comer speed as dictated by limit load factor. The data show a significant increase in instantaneous turn rate due to the unsteady aerodynamic effects, and a reduction in comer speed. These improvements, however, are meaningful only if the persistence of the dynamic lift is sufficient to allow the forces to integrate into significantly improved turning angles compared to the baseline performance capability. A preliminary assessment of the required persistence was made using a very simplified analysis. The experimental data (figures 19 and 20, for example) indicate that the lift bleedoff following motion cessation is approximately exponential. Thus this behavior was modeled as follows:

K = initial A C L ~ = 0.5

a = 3.22V/tc;

where "a" was chosen such that at the end of the persistence time, tc, only 4% of the initial lift increment remained.

Using this model and a simple point mass analysis method, the effect of persistence on the turn performance of the airplane was addressed. Figure 26 shows the increased turn angle capability due to dynamic effects for various persistence times. These results suggest that for the dynamic lift effects to become significant in terms of a tuning advantage, the required persistence of the unsteady effects would probably need to be in excess of 100 convective time units (incremental turn angle > 4"). As discussed earlier, maximum values of only about 50 convective time units were seen during the experiments. Thus it appears that the natural persistence of

the dynamic lift overshoots is too low to provide significant turn benefits. Work at Langley Research Center is now focusing on active flow control concepts to obtain dynamic characteristics with desired levels of magnitude and persistence.

CONCLUDING REMARKS

A low-speed wind-tunnel study has been conducted to examine the effects of large-amplitude pitching motions on aerodynamic characteristics of a modem fighter airplane configuration. Data for oscillatory motions were obtained for comparison with results previously measured for a flat-plate delta wing. Additionally, ramp motions over an angle of attack range from 0" to 75" were used to investigate the effects of reduced pitch rate and motion time history, and to determine the persistence of the unsteady effects. Results of the investigation can be summarized as follows:

1. Substantial increments in lift, drag, and pitching moment due to unsteady aerodynamic effects were measured for the rapid, large-amplitude pitch motions.

2. The characteristics of the observed increments in aerodynamic coefficients due to unsteady effects were qualitatively similar to those seen previously for highly swept flat-plate delta wings. The basic mechanism observed during the delta wing studies - the lag in leading-edge vortex breakdown and reformation - was also seen in the tests of the present configuration for the LEX vortex system.

3. The unsteady aerodynamic effects were seen to be highly dependent on pitch rate and motion time history. At identical instantaneous conditions of angle of attack and pitch rate, large variations in forces and moments were measured due to differences in previous motion time history. These effects were caused by lags in both flowfield formation and flowfield breakdown

during the large-amplitude pitch motions.

4. Persistence of the dynamic effects was a strong function of rate and motion time history. Maximum persistence of the lift increments measured was approximately 50 convective time units, which corresponds to approximately 2 seconds for a full-scale airplane in low-speed maneuvering flight. This level of persistence is insufficient to significantly improve airplane turn performance.

5. Unsteady effects on pitching moment characteristics, if unaccounted for, may lead to erroneous predictions of airplane pitch rate capability during large- amplitude maneuvers.

REFERENCES

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

Studies in a Wind Tunnel. AGARD CP- 413. October 1896.

7. Lambourne. N.C. et al: The Behavior of the Leading-Edge Vortices Over a Delta Wing Following a Sudden Change of Incidence. ARC-RIM-3645 1970.

8 . Gad-el-Hak. M.; and Ho, C.: The Pitching Delta Wing. AIAA Journal Vol. 23, No. 11. November 1985.

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

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

1 1. Erickson, G.E.: Water Tunnel Flow Visualization and Wind Tunnel Data Analysis of the F/A-18. NASA CR- 165859. May 1982.

2. Nguyen, L.T.: Flight Dynamics Research for Highly Agile Aircraft. SAE paper 892235, September 1989.

3 . Lorber, P.F.; and Carta. F.O.: Airfoil Dynamic Stall at Constant Pitch Rate and High Reynolds Number. AIAA 87-1329, June 1987.

4 . Galbraith, R.A.; Niven, A.J.; and Seto. I..)'.: On the Duration of Low Speed Dynamic Stall. ICAS-86-2.4.3.

5 . . Francis, M.S. et al: An investigation of Airfoil Dynamic Stall Overshoot of Static Airfoil Characteristics. AIAA-85-1773 August 1985.

Figure 1 - Sketch of model 6 . Wolffelt, K.W.: Investigation on the Movement of Vortex Burst Position With Dynamically Changing Angle of Attack for a Schematic Deltawing in a Watertunnel With Correlation to Similar

Figure 2 - Test set-up

Figure 4 - Static pitching moment characteristics

Figure 5 - Effect of pitch oscillation amplitude on lift coefficient k = 0.0226

Figure 3 - Static lift characteristics

Figure 6 - LEX vortex flow visualiza- tion, a = 27O

-3- Static ---- - 0.0061 - r0.0182

Figure 8 - Effect of positive pitch rate on lift coefficient

Amplitude

---- 30" -- 20" 1 0"

a, deg

Figure 9 - Effect of positive pitch rate

Figu. 7 - Effect of pitch oscillation amplitude on pitching moment coefficient. k = 0.0226

on pitching moment coefficient

Figure 10 - Effect of negative pitch rate

on lift coefficient

.3 r

Figure 11 - Effect of negative pitch rate on pitching moment coefficient

Static 0.0 182

Figure 12 - Effect on pitch rate on lift- drag polar

* Static

-- - 20" - 75" 40" - 75"

Figure 13 - Effect of motion history on n

lift coefficient, 9 = 0.0182

Figure 14 - Effect of motion history on normal force coefficient, A

9 = 0.0182

Figure 15 - Effect of motion history on pitching moment

n coefficient, 9 = 0.0182

2.5 - oynamk a - range

2.0 - --O-- Static

1.5 -

CL 1.0 -

.5 -

0 -

Figure 16 - Effect of motion history on n

lift coefficient. 9 = 0.0182

Figure 17 - Effect of motion history on normal force coefficient, A

9 -0.0182

Dynamic a - range * Static ---- - 75O - 0" 50" - 0"

Figure 18 - Effect of motion history on itching moment coefficient 9 = -0.0182

Figure 19 - Lift time history for ramp A

motion, 9 = 0.0182

Figure 20 - Lift time history for ramp A

motion, 9 = 0.0182

Figure 21 - Persistence for nose-up ramp motions

Figure 22 - Persistence for nose-down ramp motions

Figure 23 - Fundamental high-a maneuvers

.1

0

-. 1

Cm -.2

-3

-.4

-.5 Dynamic (experiment)

-'610 0 10 20 30 40 50 60 70 80 a, deg

Figure 24 - Comparison of dynamic pitching moment characteristics with predicted values

Turn rate, deglsec

0 Basic

Figure 25 - Dynamic effects on instantaneous turn rate

Figure 26 - Effect of dynamic lift persistence on turn performance