AIAA 93-0050 Experimental Investigation of Vortex-Fin Interaction

Anthony E. Washburn VIGYAN, Inc Hampton, VA

Luther N. Jenkins NASA Langley Hampton, VA

Marty A. Ferman McDonnell Douglas Government Aerospace St. Louis, MO

31 st Aerospace Sciences Meeting & Exhibit

January 11-14,1993 / Reno, NV

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

AIAA-93-0050

EXPERIMENTAL INVESTIGATION OF VORTEX-FIN INTERACTION

d Anthony E. Washhum'

VIGYAN. Inc. Ifompion, Virginia

Luther N. Jenkins** NASA h g l q

Hompron, Virginin

Marty A. Fermant McDoMell Douglas GovernmeM Aerospace

SI. Louis, MO

Abstract An experimental investigation has been conducted to

examine the mechanisms of vortex-fin interaction on a twin- fin configuration. The investigation included a parametric study of the effect of tail location. The vortices were generated by a 76' sharpedged delta wing with vertical tails mounted behind the wing. The model included both a dynamically-scaled flexible tail and a pressure instrumented rigid tail. Surface oil-flow patterns, off-body laser light sheet visualizations, aerodynamic load measurements, mean and unsteady flexible tail response, and unsteady tail surface pressure measurementS were obtained. The results show that the tail location did not affect the upstream trajectory of the delta wing vortex. The tail location did affect the location of vortex breakdown, the global structure of the flow field, the aerodynamic loads, and the fin buffeting levels. The buffeting levels were reduced as the fins were moved laterally toward the vortex core trajectory. Two distinct peaks were observed in the pressure excitation spectra in the post-breakdown flow. Finally, the presence of the flexible tail opposite the rigid pressure tail altered the pressure measurements at one angle of attack.

Nomenclature A R b C

Aspect ratio span Root chord Lift coefficient Lift coefficient slope Frequency Contribution to power spectrum of p' / qt in a frequency band dn Uncorrected auto spectral density Corrected auto spectral density Root bending moment of tail

~~

Research Engineer, Member AIAA. Aerospace Engineer.

t Currently Associate Professor, Parks College, Cahokia, IL, A$sociate Fellow AlAA

**

?his paper i s &Id a work ofthc U.S. Govcmrnent and is not subjcct 10 copyrighl protection in lhc United States.

d

MRB

n m P Q-

Re S f U ,

AP

x . v. 2

Non-dimensional root bending moment of tail, M a a l q ~ b , b Nondimensional frequency, fblU,

Nondimensional pressure fluctuations Surface pressure Free-stream dynamic pressure Buffet pressure Reynolds number based on chord length Planform area Thickness Free-stream velocity Model coordinates measured from amx . ,.

a Angle of attack ate

Eb Bias e m r E Random error y2(f)

RMS Root mean square

Angle of attack when breakdown crosses [railing edge

Coherence between tail pressures and free- smam pressures

Subscripts: f Refers to tail

Note: Bar over symbol means RMS value

Introduction Future combat aircraft will require agility at high angles of

attack. Separated vortical flow conditions are characteristic of aircraft in this flight regime. The vortices are generated by highly swept wings, leading-edge extensions (LEX), and chined forebodies. These vortices may impinge on the aircraft empennages and cause severe buffeting. This type of buffeting has been studied extensively on the F/A-18,1-4 and the most severe buffeting has been attributed to the highly turbulent flow resulting from breakdown of the LEX vortices upstream of the vertical tails at high angles of attack. The vortex breakdown causes a peak in the pressure spectrum that can tune to different structural modes depending on angle of attack and velocity. The pressure peak shifts to a lower frequency as the angle of attack is increased.4

1

Aerodynamic control of the position and structure of the vortices is one method for tin buffet reduction. There have been several attempls to reduce the buffeting on the F/A-I8 by controlling the LEX vortices. These include LEX deflections, dorsal fin extensions, and the addition of an indusny-developed fence on the upper surface of the F/A-I 8 LEX. The LEX deflection and dorsal tin extension studied by Rao et a15 reduced the fin buffet levels, but also diminished performance by reducing the lift. The most successful buffet attenuation technique for the F/A-l8 to dale has been the LEX fence. Lee and Brown2 found that the LEX fences had little effect on the aircraft pitching moment and lift. Martin and Thompson3 found that the LEX fence had little effect on the longitudinal position of vortex breakdown, although the vertical location and general shape of the region were altered. Both sets of investigators ObSeNed that the unsteady pressure fluctuations on the tin were reduced with the LEX fence installed, and Matin and Thompson noted that the energy in the flow was distributed over a much wider frequency band. The LEX fence method of buffet alleviation is, however, very geometry dependent and the effectiveness of the fence is very sensitive to small changes in position.

Research on fin buffeting has mostly been confined to buffet prcvention on existing airframes. In the futurc, the topic will become an integral part of combat aircraft design. Further research to investigate the mechanisms of vortex-fin interaction and the closely related phenomenon of vortex breakdown must be conducted to achieve this goal. Unlike wing buffet, vortex breakdown-induced empennage buffet is essentially an incompressible flow phenomenon since it occurs at high angles of attack that are only attainable on aircraft at low dynamic pressures (Zimmerman et a16). Due to the incompressible nature of the flow and the relative insensitivity of leading-edge vortices to Reynolds number, small-scale models can be used in atmospheric subsonic tunnels to investigate this behavior.

A cooperative research effort was initiated between the NASA Langley Research Center and McDonnell Douglas Government Aerospace to study these phenomena. The objectives of the program were to increase the understanding of the vortex-fin interaction process and to expand the data base necessary to further validate the buffet response prediction methods developed by Ferman et The investigation was conducted using the flow generated by a 76' delta wing with twin vertical tails placed aft of the wing on tail support booms. This geometrically simple configuration generated the pertinent flow-field physics necessary to study the relationship between buffeting and the vortical flow field over the wing.

Model Description The model consisted of a 76" (AR = 1) delta wing with

twin vertical tails mounted aft of the wing (Figure I ) . The delta wing was identical to the geometry used by Humme17 and had a sharp leading edge, a flat upper surface, and a triangular cross section. The root chord was 1.5 feet (0.46 m] and the maximum thickness was 2.1% of the rmt chord at x/c = 0.9. The leading edge radius was 0.005 inches [.13

2

mml and roughness on the upper and lower surfaces was 32 p-in [0.8 pm] and 64 p-in [1.6 pml respectively. A schematic of the delta wing model is presented in Figure 2. The tail support booms were attached to a t-plate under the delta wing and could be mounted at three spanwise positions corresponding to 33%. 56% and 78% of the wing semispan. Holes were drilled along the length of the tail support booms (Figure 1) to allow chordwise variation of the tail locations. The model was tested with the tails at three chordwise locations. These locations corresponded to positions of x / c = 0.95, 1.06, and 1.14 (referenced from the apex of the delta wing to the tail leading edge). The vertical tails were oriented normal to the upper surface of the delta wing and had a centerline sweep of 53.5'. The tails had a span of bL = 4 inches tO.10 m] and a mt chord of cI = 4.55 inches t0.16 m] with a taper ratio of 0.23. Figure 3 shows a schematic of the tail with dimensions. Three tail models were used in the investigation; a dynamically scaled flexible tail, a non- instrumented rigid tail, and an instrumented rigid tail. The flexible tail was instrumented with two strain gages oriented to measure root bending and torsional buffet response. The instrumented rigid tail had ten dynamic response pressure transducers, five per side as shown in Figure 4.

Figure 1. 76' delta wing with tails

Y .

2. Schematic of 76' delta wing

Test Facility and Conditions The investigation was conducted in the NASA Langley

Research Center Basic Aerodynamic Research Tunnel (BART). The BART* is an open-return wind tunnel with a closed rectangular test section 28 inches r0.71 ml high, 40 inches [LO2 m] wide, and 10 feet r3.05 ml long. The maximum flow velocity in the t a t section is 220 ft/sec, which yields aRe@ of 1.4 million. The airflow entering the test section is conditioned by a honeycomb, four anti- turbulence screens and an 11:l contraction ratio.

1st bending - 33 Hz 2nd bending - 180 Hz 1 si torsion - 307 Hz

1 .OS H

'd

I k0.384 c- Ail dimensions in inches 4.55' * 20" Bevel

Model COnStN&'Ofl Single Aluminum Spar Balsa Wood Covering Ballast Weights @ 15% and 85% Chord

Figure 3. Schematic of flexible tail

! & a t b n c a Q I s ! ~ 1 50% w/.

4 33% 50% I//-

tail - Figure 4. Schematic of instrumented rigid tail

The flow conditioners provide a low-turbulence, uniform flow in the test section. The turbulence intensity ranges from 0.03% to 0.09% as the dynamic pressure ranges from 3 lb/ft* to 30 1b/ft2 (50 5 U, 5 160 ft/sec).

The model was supported by a short sting mounted on a post support. The sting was attached to the underside of the model to reduce mounting interference on the flat upper surface of the delta wing. The post-monnt height was fixed .-/

for the investigation, so the model moved off the tunnel centerline as the angle of auack changed.

Experimental Techniques

Surface Flow Visualization

Surface flow visualizations were obtained on the vertical tails atRe of 0.75 x 106 and angles of attack of 20". 30", and 32'. A mixture of titanium dioxide (TiO2) suspended in kerosene, with a small amount of oleic acid added as an anticoagulant, was spread on the tails. The tunnel was quickly brought to test conditions and maintained until the kerosene had evaporated and left the Ti& residue. The tails were removed from the model and photographed for a permanent record. This technique provides the fine-grain detail necessary to identify features such as separation and attachment lines on the model surface.

Off-body Flow Visualization

Laser light sheet flow visnalizations were obtained at Re of 0.5 x 106 for 20" 5 a 5 40". The light source was a 6-W argon ion laser. The light sheets were produced by two methods; a scanning twin-mirrored galvanometer9 and a cylindrical lens. The smoke for the light sheet was produced by vaporizing propylene glycol at 380°F and introduced into the tunnel circuit upstream of the honeycomb.

The light sheet flow visualizations were used to identify global features of the flow field, map the vortex trajectories, and determine the location of vortex breakdown. Vortex trajectories were mapped by digitally processing video frames and displaying them with flow analysis software in three-dimensional space.

Aerodynamic Load and Moment Measurements

The mean aerodynamic loads were measured with a six- component, internal strain gage balance. The load data were obtained at Re of 0.5, 0.75, 1.0, and 1.25 x 106 for -6'5 a 5 40". The analog signal was low-pass filtered at 10 Hz, and digitally sampled at a rate of 100 samples per second. The mean was obtained by averaging 512 measurements. The load results were corrected for second-order interactions, sting and balance deflections, and tunnel upflow. The longitudinal aerodynamic coefficients are presentd in the stability axis system. The moment reference center was located on the wing centerline of the flat upper surface at 0.5~. This location corresponds to 25% of the mean aerodynamic chord. No wall interference corrections were applied to the data.

The mean and unsteady root bending and root torsion moments of the flexible tail were obtained using two strain gages mounted in the tail. The flexible tail moments were obtained at Re of 0.5, 0.58, and 0.75 x 106 for 0" 2 a 5 40". Figure 5 illustrates the signal conditioning performed on the strain gage outputs. The signals were amplified and split in two. The signals used to determine the mean moments were low pass filtered at 10 Hz. To obtain the best Dossible signal-to-noise ratio. the lines used to determine the - unsteady moments were conditioned in the following

3

sequence: 1 Hz high pass filter, amplification, 800 Hz low pass filter, and further amplification. The mean moments were obtained with an integrating voltmeter sampling at 40 Hz. The fluctuating signals were digitized with 12-bit transient data recorders sampling at 3900 Hz.

ifferential mplifier

Figure 5. Data acquisition system

The strain gage outputs were calibrated by hanging weights on the flexible tail for both bending and torsion. A ground vibration test determined frequencies of the first 3 bending and first 2 torsional natural modes of the flexible tail. These results are tabulated in figure 3.

The digitized unsteady time history results were used to calculate the total RMS moments and the power spectral density estimates. The power spectral density estimates were calculated using fast fourier transforms and the Hanning window. The RMS moments, nRE, due to a specific natural bending mode were calculated by integrating the power spectral density estimates in the frequency band pertaining to the mode of interest.

Tail Pressure Measurements

Mean and unsleady pressure measurements were obtained using 11 dynamic surface pressure transducers -- IO mounted flush on a rigid vertical tail and 1 mounted in a free stream static pressure probe. The surface pressures were obtained at Re of 0.5, 0.58,0.75, and 1.0 x 106 for 0' 5 a S 40". The signal conditioning technique was the same as that used for the flexible tail s!nin gages illustrated in Figure 5. The mean pressure data were obtained in the same manner as the flexible tail bending moments. The fluctuating data were recorded ci~ FM tape with a bandwidth of 1 Hz to 5000 Hz. The pressure data resulk were digitally sampled from the tape with 12-bit transient data recorders at 3900 Hz. The analog pressure data were low pass filtered at 750 Hz. The data were used to calculate the RMS pressures, power spectral densities, cross spectral densities, and histograms. The coherent acoustic noise present in both the free stream and the tail pressures was removed from the tail pressure spectra using ~ ( f ) = (I - 7'(f))c'(f) as suggested by Bendat and Piersollo. The surface pressure excitation spectra are presented as the non-dimensional RMS pressure parameter .\lnFo as a function of n. The use of this parameter was recommended by Mabey" for comparison of buffet test results and is defmed as

Measurement Uncertainty

The uncertainty in the vortex core trajectories from the laser light sheet results is estimated to be within +1/8 inches

for the angle of attack at which breakdown crosses the wing Wailing edge are expected to be within k0.25'.

Uncenainty in the mean aerodynamic loads, the mean and RMS flexible tail dynamic loads, and the RMS tail pressures were determined using the analysis described in the ASME Performance Test Codes supplement on Measurement Uncenainty.'2 The total uncertainty number was determined by the root-sum-square method, which gives a 95% confidence interval

or f1.4% of the wing span. The laser light sheet estimates u'

111 E = [E,' + (t&,)']

where &b is the bias limit, cr is the precision index of the mean, and f is the 95th percentile point for the two-tailed Student t distribution. There are no bias errors inherent in the estimate of mean and mean square values. The effect of precision error is reduced by averaging several measurements and using the average. The precision error in the mean value of a series of measurements is given by

where u is the standard deviation of the measurements and N is the number of measurements. The precision error in the RMS value of a series of measurements is given by

The uncertainty is propagated from the measurements to the result by calculating the sensitivity so that if a quantity P = f ( s , . s , , ..., s,), then the uncertainty bias limit or the urecision index of a result can be given as

L

The bias limit uncertainties for the lift, drag, and pitching moment coefficients at Re = 1.0 x I@ are 0.04, 0.03, and 0.004 respectively. There are no estimates for the precision index of the aerodynamic load data because the necessary statistical data were not stored. The precision error in the mean measurement is however, expected to be very small.

The bias limit uncertainties for the mean and RMS non- dimensional flexible tail bending moments are 3%. The precision index of the mean and RMS non-dimensional moments are 0.8% and 0.5% respectively.

The bias limit uncertainty for the non-dimensional pressure measurements was 10%. The precision index of the RMS non-dimensional pressures was 1.1%.

The statistical errors in the computation of the spectral quantities are described in Bendat and Piersol.10 The normalized bias and random errors for the power spectral densities, G(fJ, are

respectively. The variable E , is the spectral resolution bandwidth and nd is the number of ensemble averaged

records. The second derivatives, G"(f), were computed by using a second order cenual difference.

Two hundred ensemble averaged records were used to determine the frequency domain quantities with Be = 1.91 Hz. The normalized errors for the auto spectral density estimates were found to be €6 < 3.6% and cr = 7% for the tip transducers for angles of attack between 28" and 40". These errors propagated to values of €b < 3.1% and cr = 3.5% in the buffet excitation parameter zjnF(n).

d

All of the data acquired were reptable.

Results and Discussion

Parametric Study

C h o r h Forward center

AA

'J

Inboard Midspan Outboard 29.6' 3 . 2 0 24.10 32.2' 27.9' 27.10 31.2O 30.6O 29.2"

A parametric survey was conducted as the first phase of the investigation with the tails located at nine positions: the 3 spanwise tail boom positions, and the 3 chordwise positions. In the f is t phase, laser light sheet visualizations were used to characterize the effect of the vertical tail position on the vortex structure and the position of vortex breakdown.

To determine the position of vortex breakdown, the light sheet was oriented to illuminate a vertical plane along the vortex trajectory. The breakdown location was defined by the sudden expansion of the vortex core region as shown in Figure 6. Table 1 shows the angle of attack at which the vortex breakdown crossed the trailing edge of the wing as the model was pitched slowly upward. For comparison, the ale was 32.5' for the wing and t-plate alone, without the tails or support booms. At all 3 tail chord positions, ale was reduced as the tails were separated laterally. The

and held that position within 10% chord. The same abrupt jump to 70% was observed when the tails were placed at the inboard positions and at the most aft chord positions. However, when the tails were located in the middle and outer positions in the forward and center chord positions, the character of the phenomenon was altered. At these positions, the breakdown did not loiter aft of the trailing edge as distinctly as before. When are was reached, the breakdown slowly crept over the sailing edge to 90% chord or greater. In at least one instance the breakdown oscillated over the trailing edge. The breakdown location slowly crossing the trailing edge was believed to be due the increased adverse pressure created by the tails W i g placed in the vortex path and altering the conditions at the trailing edge.

Table 1. Incidence at which breakdown crosses trailin edge r\ Soanl I I

The global structure of the flow was also observed by orienting the light sheet to illuminate both cross-sectional planes normal to the wing and planes parallel to the model down the length of the vortex. The primary vortex cores traveled outboard and above the tails when the tails were placed at the inboard and midspan positions. The primary vortex core traveled nearly directly above the vertical tail when the tails were in the outboard positions.

A vortex was observed to form along the leading edge of the tail as illustrated by Figure 7 at both the midspan and .~

greatest reduction in ate occurred between the inboard and midspan tail positions. In general, as the tails were moved forward toward the delta wing, the ui, was further reduced. The dependence of the vortex breakdown position on tail location was surmised to be due to an adverse pressure gradient induced by the presence of the tails.

outboard tail positions. Figure 7 shows the vortex s-mcture from above and in front of the model. In this figure, the P b W vortex core is directly above the vertical tail and the tail vortex is on the outer side. The leading-edge tail vortex was formed by the outward spanwise flow from the underside of the primary vortex impinging on the tail at an effective angle of attack. The spanwise flow separated at the

Figure 6. Vortex Breakdown, a = 30, outboard center tails

The manner in which the vortex breakdown crossed the trailing edge was also a function of tail location. When the tails were removed, the breakdown loitered just aft of the trailing edge as the angle of attack was increased. At a, the breakdown abruptly jumped to approximately 70% chord

5

tail leading edge to form a vo&x on the outer surface of the tail. This tail vortex was observed to be stable at all angles of attack, even when the primary vortex breakdown occurred upstream of the tails. No evidence of a tail vortex was observed when the tails were at any of the inboard locations.

A third vortex was seen below and slightly outboard of the tail vortex. This third vortex was a tip vortex from the t- plate under the model. It was observed for all nine tail positions and for the delta wing and t-plate alone. The thud vonex was not present when the t-plate was removed.

The effects of the tails on the model lift coefficients are presented in Figures 8 and 9. Figure 8 presents the lift coefficient for the midspan tail locations as a function of chordwise placement and compares them with tail-off data. The maximum lift coefficient was reduced when the tails were installed on the model and was further reduced as the tails were moved forward toward the wing. This trend was consistent at all the spanwise tail locations. Figure 9 shows that the lift coefficient was not as sensitive to the spanwise

tail location, and the maximum lift coefficient was unaffected. The angles of attack between which a reduction in C L ~ were F i t observed are shown in Table 2 for all tail positions. The trends and angles in Table 2 are similar to those shown in Table 1. Hence it a p p s that the reduction in C L ~ is related to the Occurrence of vortex breakdown crossing the wing trailing edge. Similar trends were observed in the pitching moment curves, where the presence of the tails decreased the pitch stability at high angles of attack. As the tails were moved forward, the decrease in stability occurred at lower angles of attack. No significant changes in the pitching moments were observed when the tails were moved in the spanwise direction.

Figure 7. Vortex Smctures, a = 20, outboard center tails, xlc = 1.3, front view

1.4 F

1.2

1 .o 0.8

0.6

0.4

negative at low angles of attack, become positive by a = lo", reach a maximum value between 20' and 303 and are reduced at higher angles of attack. Two possible reasons for the outward bending of the tails can be interpreted from the laser light sheet results. The suction pressure caused by the primary vortex passing outboard of the tail would create an outward bending moment for the inboard and midspan tail positions. The effective angle of attack induced by the outward spanwise flov from the primary vortex on the tail also contributes to the outward bending moment. The maximum root bending moment occurred at the center, inboard tail location. The mean moments were independent of Reynolds number for the range tested.

'4

& Center, outboard tails -e- Center, midspan tails

0.4 -4- Center, inboard tails

0.2 1 0 15 20 25 30 35 40 45

Figure 9. Effect of spanwise tail placement on lift curve a (deg)

0.20 0.25 E

0.2 10 15 20 25 30 35 40 45

a (deg) Figure 8. Effect of chordwise tail placement on lift curve

The mean root bending moments from the flexiblc tail are presented in Figure 10 as a function of angle of attack for all nine fail positions. Positive moments correspond to an outward force on the tails. The moment curves exhibit similar trends for all nine tail positions; the moments sWt

6

a (deg) a) forward tail positions

0.25

0.20 F

+ midspan +inboard

-0.05-5 O.OO 5 L 15 25 35 4

45 a (deg)

b) center tail positions

Figure 10. Mean bending moment on the flexible tail W

4 45

a (deg) c) aft tail positions

Figure 10. Mean bending moment on the flexible tail (cont.)

The total (all structural modes combined) RMS root bending moments from the flexible tail are presented in Figure 11 as a function of angle of attack for all nine tail positions. Most of the flexible tail buffet response occurred in the first bending mode as determined by integrating the power spectral density estimates. The second bending and first torsion buffet response curves were similar in shape to the first bending response curves in the high angle of attack region. A large increase in buffet response (buffet onset) occurred in all cases near a = 30". Flow visualization has shown that the angle of attack for buffet onset is similar to the at,. The buffet loads were greatest with the tails at the inboard positions, regardless of the chordwise placement. The buffet loads were lowest with the tails at the o u h a r d locations. Of the nine tail positions, the center inboard position had the greatest buffet loads.

Detailed Study

The resulcs of the parametric survey were used to select two tail positions for further detailed study, which included the measurement of unsteady tail surface pressures. The center inboard position was selected as one condition since it exhibited the largest buffet response throughout the angle of attack range. The center outboard position was selected as the second condition since it exhibited one of the lower buffet response levels, completely different global flow field features, and required movement of the tail in the spanwise direction only. These two tail positions will hereafter be referred to as inboard and outboard respectively.

Surface flow features were investigated on the tails for a = 20", 30'. and 32". An example of the surface patterns for the inboard case at a = 32' is shown in Figure 12. Figures 13 and 14 are sketches comparing the surface patterns at a= 20' and 32" for both the inboard and outboard tail positions respectively. In the four cases, on both inner and outer sides, the flow separated along the tail booms just below the bolt holes near the maximum width.

The inner surface patterns are similar for the two angles of attack at both tail locations. The flow on the inner side attaches on the kve l near the leading edge when the tails are at the inboard location. The flow on the inner surface attaches at the juncture between the bevel and side when tails are at the outboard location. This indicates that the there was greater cross flow on the tail in thc outboard position.

.../

J

7

-B- inboard

-5 5 15 25 35 45 a (deg)

a) forward tail positions

S outboard

-5 5 15 25 35 45 a (deg)

b) center tail positions

+ inboard

-5 5 15 25 35 45 a (deal

c) aft tail positions

Figure 11. RMS bending moments on flexible tail

Figure 13 shows that the inboard tail position patterns on the outer surface are a function of angle of auack. However, the general features of the flow are independent of angle of attack. The flow was observed to attach on the inner surface and subsequently flow around the leading edge to the outer surface. On the outer surface, the flow separated at the bevel juncture, and midway between the leading edge and the juncture for a = 20" and 32" respectively. The light sheet results did not indicate separation at the tail leading edge either. A separated region was observed to emanate from the tail leading edge root at both angles of attack. The flow moved in essentially the chordwise direction in triangular regions above and below the separated region. The inclination of this structure referenced to the tail leading edge was a function of angle of attack. Due to these features, it

did not appear that this structure was a leading edge vortex for the inboard tail case. The light sheet results for the inboard tail case also showed no indication of a leading edge tail vortex.

the primary vortex breakdown location moved in front of the tails.

Inner side

Figure 12. Surface flow, a = 32", inboard center tails

Inner side

Outer side

a) a= 20'

b) a = 32'

Figure 13. Surface flow, center inboard tails

Figure 14 shows that the outboard tail position patterns on the outer side are not a function of angle of attack. The tail vortex secondary separation line was located along the bevel juncture at both angles of attack. The tail vonex- attachment line was located near and parallel to the tail boom along the bottom of the tail. The tail vortex encompassed a large portion of the outer tail surface and all the pressure transducer locations. The surface flow patterns were unchanged between these two angles of attack, even though

a) a = 20'

W b) a= 32" Figure 14. Surface flow, center outboard tails

Laser light sheet images were used to examine the vortex structures and the core trajectories for these two tail positions in greater detail. Figure 15 presents sketches of the flow features for both tail positions at a = 20" and x/c = 1.3 and 1.5 as observed from light sheet images normal to the model. The light sheet position at x/c = 1.3 is indicated in Figure 15. At x/c = 1.5, the light sheet is just behind the tail.

The light sheet images clearly revealed the tail vortex formed along the tail leading edge in the outboard tail case. The tail vortex was rotating in the opposite sense of the primary delta wing vortex, and formed on the outboard side of the tail. The tail vortex was formed due to the high sweep angle of the vertical tails and was similar to the vortex formed by the dorsal fm extensions on an FjA-I %like model studied by Rao et al.5 The tail vortex was stable at all angles of attack. The primary and tail vortices begin to merge near the tail tip where they began spiraling around one another as they travelled downstream.

The flow patterns observed with the tails at the inboard position did not show the same features as the outboard position results. Actually, the primary vortex shllcture was very similar in the inboard tail case to the primary vortex structure in the tail-off case. The inboard tail surface flow visualization showed a separated swucture on the outer side. This was not believed lo be a leading-edge vortex, otherwise entrainment of the smoke into that vortex would be

8

expected. The lack of a tail leading-edge vortex for the inboard tail case could be a result of the reduced sidewash angle in the flow impinging on the tail.

b) outbaud tails Figure 15. Vortex structures, inboard and outboard tails d

The off-body flow visualizations were also used to determine the innuence of the tail position on the vortex core trajectories. Figure 16 shows this effect in the x-z plane at 20" and 30° angle of attack for the two tail positions compared with the tail-off case. The presence of the tails in either position did not affect the vortex trajectory upstream of the tails. The primary vortex trajectory was only influenced with outboard tails where the interaction between primary and tail vortices began at X / G = 1.4 and continued on downstream. The presence of the tails did promote premature vortex breakdown as illustrated in Figure 16 at a = 30'. As the tails were moved toward the vortex core trajectory, the vortex breakdown position was translated upstream.

The tail surface pressure measurements have been used to calculate mean and RMS values and the buffet excitation spectra at each transducer location. The RMS buffet pressures were calculated from the pressure time histories on both sides of the tail. The buffet pressure is defined as the instantaneous differential pressure across the tail surface. This quantity was calculated by digitally differencing the pressure signals for the two identically positioned hansducers on each side of the tail. The digitally differenced results are then used to determine RMS and spectral characteristics. Buffet pressure is an important measurement for structural scaling law development4. The buffet pressures are better representations of the inputs for structural response prediction techniques than are the individual surface pressures.

zlc - Q - Tail vortex, Outboard tails

xlc a) a= 20"

+ Inboard tails

0.0 0.50 1 .o 1.5 2.0 xlc

b) a = 30' Figure 16. Vortcx core trajectories

9

-e- Re = 0.58 x lo6 - flex tail - Re = 0.75 x 106 -flex tail 0.35 -e- Re = 0.58 x 10;

0.30 [ - Re = 0.75 x 10

8 v la a

Figure 17. RMS buffet pressures, inboard tails, transducer location 4.

The inboard tail RMS buffet pressures are presented in Figure 17 as a function of angle of attack at transducer location 4. The RMS buffet pressures were a strong function of transducer position and location 4 yielded the greatest levels. The characteristics of the RMS buffet pressure curves compare very well to the RMS moment response measured by the flexible tail at this tail location (Figure Ilb). Good correlation between the RMS buffet pressures and the RMS moment response was also observed for the outboard tail case.

Ferman et ai4 showed that the buffet pressure due to vortex breakdown could be non-dimensionalized by the dynamic pressure if other parameters were held constant. These results support that observation. In Figure 17 the RMS buffet pressures at two different Reynolds numbers are presented after being non-dimensionalized by q. There was a slight Reynolds number dependence in the buffet pressures before the vortex breakdown moved upstream of the tails. When the tail buffeting was dominated by upstream breakdown, the buffet pressures are Reynolds number independent This ObSeNatiOn held true for both tail positions; for both

Generally when both a flexible tail and a rigid pressure tail are used in a twin-tail buffet experiment, they are tested in conjunction. Figure 17 illustrates that this can yield misleading results. When the tails were at the inboard position, the presence of the flexible tail altered the pressure measurements on the opposite side of the model at a= 30'. This occurred consistently for the three dynamic pressures tested. The phenomenon was very subtle since it occurred at one angle of attack only, and then not at an angle where maximum buffeting occurred. The uniqueness of this angle of attack was interesting. The only remarkable feature about this angle was that the vortex breakdown position was near the tails. However, when the breakdown was loitering at the trailing edge, for a = 32", no flexible tail influence was observed. Similarly, no influence was observed on the pressure measurements with the tails in the outboard position.

/ q versus a and .\lnFo versus n.

10

At low angles of attack the RMS buffet pressures are small. The buffet pressures increased sharply between 1 9 and 20" angle of attack. By a = ZOO, the primary vortices had grown large enough so that the shear layers were scrubbing the outboard surfaces of the tails. Between Z O O

and 32' angle of attack, the buffet pressures levelled off and maintained a nearly constant value (two rigid tails). At 32' angle of attack, the sharp increase in buffet pressures corresponded with the vortex breakdown position crossing the trailing edge (Table 1).

Figures 18 and 19 present the ten RMS surface pressures for the inboard and outboard tail positions. The results show that the RMS surface pressures increased in all instances when vortex breakdown moved upstream of the tails. However, the surface pressure fluctuations were sensitive to the tail position, the tail side, and the transducer location. The greatest overall buffet levels occurred on the inboard tails. In this case, the outer surface exhibited the largest RMS pressure levels. The increase in buffet pressures at a= 20" shown in Figure 17, was only observed on the outer side of the tail in Figure 18. This suppons the observation that the primary vortices grow and the shear layers begin to impinge on the outer tail surfaces. Unlie the inboard tail case, the inner side RMS pressure levels were the larger in the outboard tail case. This might be due to the stable vortex that was formed on the outer side of the tail. The lowest RMS pressure levels were consistently measured at transducer locations 2 and 5. These locations correspond to the transducers furthest from the tail leading edge.

8 v la

il

-5 5 15 25 35 45 a (deg)

a) outer surface

0.30

0.25

0.20

'0.15 0.10

0.05 0.00

-5 5 15 25 35 45 a (deg)

b) inner surface Figure 18. RMS pressures, inboard tails

'd

W

II

d

& Location 1 + Location 2 + Location 3

-0- Locatlon 5 +Location4

0.

-0- Location 5

-5 5 15 25 35 45 a (deg)

b) inner surface Figure 19. RMS pressures, outboard tails

Examples of the buffet excitation spectra are presented for the two tip transducers for both tail positions in Figures 20 and 21 for several angles of attack. The non-dimensional excitation spectra were independent of tunnel speed. The spectra are the average of 200 ensembles with a bandwidth resolution of n = 0.023. The solid cwes represent results with two rigid tails installed and the dotted curves represent the case when the flexible tail was installed opposite the pressure tail.

The figures show a large increase in buffet excitation when vortex breakdown moved forward past the tails as angle of attack was increased. The spectra show that the pressure field associated with the vortex breakdown phenomenon contains energy over a moderately narrow frequency band. There were two distinct frequency peaks in this band. These peaks represent coherent fluctuations in the flow at those frequencies. The frequency was reduced and h e magnitude increased for both of these peaks as the angle of attack was increased. The relative magnitude and frequency of the two also changed as angle of attack changed. When breakdown first influences the excitation spcctra, the lower frequency was dominant. As the angle of attack increased, the higher frequency peak grew in magnitude faster and the two peaks eventually merge into a single pcak near a = 40".

The general reduction in frequency of the vortex breakdown induced pressure field as angle of attack increases, has also been observed on U1eF/A-l8~-4and on a 76" delta wing.13 The F/A-18 tail pressure excitation spectra, however, generally exhibit only one peak.

i/

11

Likewise, in the investigation performed by Rediniotis et all3 only one frequency was observed in the spectra obtained by a hot-wire downstream of the wing. The surface pressure data obtained by Martin and Thompson3 for the F/A-18 are an exception. In their investigation, the addition of the LEX fence to the F/A-l8 reduced the magnitude of the original speck4 peak and created an additional spectral peak at a greater frequency. The overall excitation energy was reduced in that case.

Initially, it appeared that the frequency distribution downstream of vortex breakdown was solely a function of angle of atfack. However, the vortex breakdown position is also a function of angle of attack. Marlin and Thompson3 mounted two unsteady pressure transducers on the model wing and one on the vertical tail. Their results indicated that the peak frequency decreased with distance downstream of breakdown. .In the present investigation, the two dominant frequencies obtained for the inner surface tip transducer spectra for both tail positions are tabulated in Table 3 for several angles of attack. At 30' angle of attack the peak kequencies measured on the outer surface tail transducer are significantly larger than the inner surface tail transducer. The relative positions of vortex breakdown for each case are illustrated in Figure 16. The tail is fanher downstream of the breakdown in the outboard case. If the trend observed by Martin and Thompson was followed, the frequencies for the inboard tail case should be larger. However, that was not the case. The frequencies observed for the inboard tail position vary less over the angle of attack range. The measured peak frequencies converged to the same values as the angle of attack was increased.

Table 3. Peak frequencies in inboard tip swc!m I II Inboard tails I Outboard tails I

Figure 20 also illustrates the effect of the flexible tail on the rigid tail pressures in the frequency domain. As in the RMS buffet pressure curves shown in Figure 17, only at 30" angle of attack does the presence of the flexible tail significantly affect the excitation spectra. In this instance, the excitation spectra resemble the spectra obtained downstream of breakdown at other angles of attack. The presence of the flexible tail seems to amplify the breakdown induced pressure fluctuations on the rigid tail. The peaks seen in the pressure spectra with the flexible tail installed do not correspond to the structural bending frequencies of the flexible tail.

0.14 E a = 28" __ No flex tail

0.12 -0.10

0.08 c 0.06

0.04 0.02 0.00

... . . . . . ... ... With flex tail

0 1 2 3 4 5 6

0.14 0.12 - 0.10 0.08

0.04 0.02

O.O0O 1 2 3 4 5 6

0.14 0.12

0.04 0.02 0.00

a = 32" No flex tail

. . , .. . . . . . . . . . With Ilex tail

0 1 2 3 4 5 6

No flex tail 0.14 E a = 34" __ 0.12

0.04 0.02

O.O0O 1 2 3 4 5 6

0.25 a = 38" uter side __ No flex tail

. ..... . . . . . . . . With flex tail - O.O0O 1 2 3 4 5 6 - n = fblU

Figure 20. Buffet excitation, inboard tails, tip transduccrs

a = 28" 0.12

C 0.06 0.04 0.02 0.00

Outer side

0 1 2 3 4 5 6

a = 30" 0.12

a = 32" 0.12

0.14 E a = 340 0.12

-0.10 0.08 10.06 0.04 0.02 0.00

Inner side

~ Outer side

0 1 2 3 4 5 6

a = 38'

0.20

.

n = fblU m

LJ Figure 21. Buffet cxcilation, outboard tails, tip transduccrs

The current results indicate that buffet excitation levels were less when the tail was in the outboard location. This occurred even though the primary vortex core passed almost directly above the tail in this instance. It seems that the tail vortex formed at the tail leading edge might aerodynamically dampen the dynamic pressure fluctuations caused by the primary vortex breakdown. The tail vortex formed on the outer side of the tail (Figures 7 and 15) was always stable. The surface flow patterns in Figure 14 indicate that the tail vortex dominated the surface flow on the outer side of the tail. A similar result was observed by Rao et al5. When a dorsal fin was added in front of the vertical tail of an "F18- like" model, a vortex was formed on the outer side of the tail and the tail buffeting levels were reduced.

There is also some evidence that the reduction in the buffet excitation levels in the outboard tail position may be due the proximity of the tail to the vortex core path. Lee and Brown2 found that the lowest pressure fluctuation levels on the F/A-18 occurred at the vortex center and increased with distance from the center. They observed this trend both with and without the vertical tails in the flow field. Hence, lower buffet levels may be achieved by locating highly swept tails near the vortex path.

Further studies in spatially correlating the pressures on the tails and correlating the tail pressures with the vortical flow field are in progress. Future research will investigate the relationships between the angle of auack and breakdown characteristics such as breakdown position and breakdown frequency.

Concluding Remarks

An experimental investigation of the vortex-fin interaction process was conducted. The investigation used a 76" delta wing with twin vertical tails. The vertical tails were placed at nine positions aft of the delta wing. The data show that the experiment captures the pertinent physics associated with vortex induced buffet on complex configurations. The results obtained have shown that the aerodynamic loads are more sensitive to the chordwise tail location than the spanwise tail location. The vortex core trajectories upstream of the tails were not influenced by the tail location, however, the location of vortex breakdown was affected. The general character of the flow field was also greatly affected by the tail location, When the tails were in the midspan and outboard locations, a stable vofIex was formed along the tail leading edge. The buffeting on the vertical tail also showed a strong dependence on the tail location. As the tails were moved toward the vortex core, the buffet response and excitation were reduced. Two distinct frequencies were observed in the pressure field downstream of the vortex breakdown. The magnitudes and frequencies of these peaks were dependent on angle of auack.

Additionally, it was found that the presence of a flexible tail can affect the unsteady pressures on the rigid tail on the opposite side of the model.

Acknowledgments

This research was supported by the National Aeronautics and Space Administration under Conuacl No. NAS1-18585. Special thanks go to Dr. Ken Visser for his assistance with

the flow visualizations. Also, thanks go to Kathryn Stacey, Kurt Severance, and Brooks Childers for their help with the computer-aided vortex trajectory mapping.

References

'Wentz, W. H., Jr.: "Vorlex-Fin Interaction on a Fighter Aircraft", AIAA 87-247442P.

ZLee, B. H. K. and Brown, D.: "Wind-Tunnel Studies of F/A-18 Tail Buffet", J. Aircraft, Vol. 24, No. 1.

3Martin, C. A. and Thompson, D. H.: "Scale Model Measurements of Fin Buffet Due to Vortex Bursting on F/A- 1 8 , AGARD-CP-497, Paper no. 12, May 1991.

4Ferman, M. A,; Patel, S. R.; Zimmerman, N. H.; and Gerstenkorn, G.: "A Unified Approach to Buffet Response of Fighter Aircraft Empennage", AGARDmATO 70th Structures and Materials Meeting, Sorrento, Italy, April 1990.

5Rao, D. M.; Puram, C. K.; and Shah, G. H.: "Vortex Control for Tail Buffet Alleviation on a Twin-Tail Fighter Configuration", SAE 892221, Sept. 1989.

6Zimmerman, N. H.; Ferman, M. A,; Yurkovich, R. N.; and Gerstenkorn, G.: "Prediction of Tail Buffet Loads for Design Application". AIAA 89-1378.

'Hummel, D: "On the Vortex Formation Over a Slender Wing at Large Angles of Incidence" in High Angle of Aftack Aerodynamics, AGARD CP-247, 1979.

8Sellers, W. L. I11 and Kjelgaard, S . 0.: "The Basic Aerodynamics Research Tunnel - A Facility Dedicated to Code Validation", AIAA 88-1997, May 1988.

9Rhodes, D. B.; Franke. J. M.; Jones, S. B.; and Leighty, B. D.: "A Twin-Mirrored Galvanometer Laser Light Sheet Generator", NASA TM 100587, 1988.

'OBendat, J. S. and Piersol, A. G.: E n g i n e e r i n g Applications of Correlation and Spectral Analysis, Wiley- Interscience, 1980.

"Mabey, D. G.: "Some Aspects of Aircraft Dynamic Loads Due to Flow Separation", AGARD-R-750.

I2ASME Performance Test Codes Supplement on "Instruments and Apparatus, Part 1 - Measurement Uncertainty", ANSI/ASME F'TC 19.1-1985.

13Rediniotis. 0. K.; Telionis, D. P.; and Stapountzis, H.: "Periodic Vortex Shedding Over Delta Wings", AIAA 89- 1923. June 1989.

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