Document
AIAA-92-4101-CP
FLIGHT TEST RESULTS FROM A SUPERCRITICAL MISSION
ADAPTIVE WING WITH SMOOTH VARIABLE CAMBER
Sheryll Goecke Powers+ Lannie D. Webb* Edward L. Friend** William A. Lokos++ NASA Dryden Flight Research Facility Edwards, California advanced fighter technology integration Abstract normal acceleration at center of gravity, g The m1ss10n adaptive wing (MAW) consisted of leading- and trailing-edge variable-camber surfaces normal acceleration at cockpit, g that could be deflected in flight to provide a near-ideal normal acceleration at horizontal tail, g wing camber shape for any flight condition. These sur- normal acceleration at wingtip, g faces featured smooth, flexible upper surfaces and fully enclosed lower surfaces, distinguishing them from con- BIR buffet intensity rise ventional flaps that have discontinuous surfaces and b wing span, ft [56.55 ft] exposed or semiexposed mechanisms. Camber shape airplane normal-force coefficient, was controlled by either a manual or automatic flight (ancg W)/qooS control system. The wing and aircraft were extensively instrumented to evaluate the local flow characteristics pressure coefficient, (p - Poo)/qoo and the total aircraft performance. This paper dis- pressure coefficient on wing upper-surface at cusses the interrelationships between the wing pres- x/c = 0.96 sure, buffet, boundary-layer and flight deflection mea- C streamwise local chord, ft surement system analyses and describes the flight ma- neuvers used to obtain the data. The results are for mean geometric chord, ft, S/b [11.0 ft] Cav a wing sweep of 26°, a Mach number of 0.85, leading-
mean aerodynamic chord, ft, 2/ S J; c dy
CMAC and trailing-edge cambers (oLE/TE) of 0/2 and 5/10, [11.2 ft] and angles of attack from 3.0° to 14.0°. For the well- behaved flow of the OLE/TE = 0/2 camber, a typical section normal-force coefficient, J b.Cpdx/c Cn cruise camber shape, the local and global data are in FDMS flight deflection measurment system good agreement with respect to the flow properties of MAW mission adaptive wing
the wing. For the oLE/TE = 5/10 camber, a maneu-
vering camber shape, the local and global data have free-stream Mach number Moo similar trends and conclusions, but not the clear-cut PCM pulse code modulation agreement observed for cruise camber.
p local wing surface static pressure, lb/ft N omenc.lature · free-stream static pressure, lb/ft Poo free-stream dynamic pressure, lb/ft Reference values in brackets, [], based on a trape- zoidal planform at a leading-edge sweepback angle of rms root mean square 26° scaled up, from Ref. 17.
wing reference area, ft [622.0 ft ]
s
TACT transonic aircraft technology • Aerospace Engineer. Member AIAA.
local flow velocity in the boundary layer,
u
•• Aerospace Engineer.
ft/sec Copyright @1992 by the American Institute of Aeronau- tics and Astronautics, Inc. No copyright is asserted in the
local flow velocity at Y = 5 in., ft/sec
United States under Title 17, U.S. Code. The U.S. Govern- ment has a royalty-free license to exercise all rights under airplane weight, lb the copyright claimed herein for Governmental purposes. All other rights are reserved by the copyright owner.
x local streamwise coordinate (from wing Flight Research Facility in April 1989. Aerodynamic leading edge), ft characteristics and performance evaluations, for exam- ple wing pressure, buffet, and lift and drag test re- x/c fraction of local streamwise chord sults, also have been discussed in separate subdisci- Y perpendicular distance above upper wing 8 11 pline reports. - The data indicate that the advanced surface, in.
fighter technology integration ( AFTI) / F-111 MAW air- y spanwise coordinate, ft craft had significantly improved aerodynamic charac- teristics, com pared to the the basic F-111 A and tran- a: indicated angle of attack corrected to wing sonic aircraft technology (TACT) designs.
reference plane, deg This paper provides a correlation of the multidis-
o.r wing reference angle of attack (o. + ~a:), deg
ciplines, showing how the interrelationships from the /3 aircraft angle of sideslip, deg wing pressure, buffet, boundary-layer and FDMS anal- ~a: correction for pitching moment and upwash yses strengthen and support each other. Also included effects are descriptions of the flight maneuvers used to obtain the data. The results are for a wing sweep of 26°, a 8LE/TE leading- and trailing-edge camber deflection, Mach number of 0.85, leading- and trailing-edge cam- deg bers (8LE/TE) of 0/2 and 5/10, and angles of attack TJ orifice row semispan locations, 2y/b from 3.0° to 14.0°. The data presented are for dynamic O"anwr rms of buffet component of normal pressures of 300 and 600 lb/ft with the majority of the acceleration, g data shown being at 300 lb/ft • Background Introduction The last research program conducted on the A wing configuration that would allow smooth cam- AFTI/F-111 research aircraft was the testing of the ber changes throughout the flight envelope can pro- MAW concept. The AFTI/F-111 MAW aircraft was vide additional aerodynamic performance at all flight initially an F-lllA airplane, which was modified for conditions. Variable camber alone has been a proven 12 13 use in the F-111 TACT Program. • The original de- concept for enhancing maneuverability for nearly all sign of the F-111 aircraft used a variable-sweep wing to flight conditions. On airplanes such as the F-16 increase the number of optimum flight conditions. The and F-18 aircraft the variable camber is achieved TACT Program combined a supercritical airfoil with through discrete flap positions. Better performance planform and twist changes to improve transonic cruise can be achieved with smooth variable camber. De- and maneuver performance relative to the conventional 2 3 sign studies • to develop a smooth, variable-camber 15 16 F-111 wing. • The MAW Program used a smooth, supercritical wing resulted in the mission adaptive wing variable-camber supercritical wing to provide high lev- (MAW). The MAW consisted of leading- and trailing- els of aerodynamic efficiency over a range of subsonic, edge variable-camber surfaces that can be deflected in transonic, and supersonic flight conditions. Previous flight to provide a near-ideal wing camber shape for any supercritical wing designs tended toward a fixed ge- flight condition. These variable-camber surfaces fea- ometry shape that was a compromise for specific mis- tured smooth, flexible upper surfaces and fully enclosed sion requirements. The MAW minimized penalties for lower surfaces, distinguishing them from conventional off-design flight conditions through the combination of flaps that have discontinuous surfaces and exposed or smooth-skin variable camber and variable sweep.
semiexposed mechanisms. The camber shape was con- trolled by either a manual or automatic flight control Maintaining an efficient airfoil shape by the use of system. camber settings was one of the basic design goals. The flight program provided adequate instrumentation to The wing and the aircraft were extensively allow evaluation of this complex design.
5 6 instrumented • to evaluate the aerodynamic perfor- mance of the MAW. Instrumentation located on Correlation of the buffet characteristics with the the MAW included orifices for surface pressures, a wing pressure distributions adds to the technical in- boundary-layer rake, a flight deflection measurement terpretation of the data. Boundary-layer data at the system (FDMS), wingtip accelerometers, strain gages, wing trailing edge support the buffet and pressure data and control position transducers. with respect to separation. The wing deflection data provide a wing definition with load factor and angle of Results from the MAW Program were summarized attack. The resulting in-flight deflections supplement at the final symposium held at the NASA Dryden the wing pressure data.
Other accelerometers (such as the cockpit accelerome- Description of Airplane and Wing ter a ) were filtered during the data analysis after 1 nckpt The AFTI/F-111 MAW airplane and the camber the flight. Stability and control parameters and surface shape of the wing are shown in Fig. 1. The airplane was position indicators were also used in the buffet analysis.
initially an F-111 airplane with the wings replaced for Boundary-Layer Rake the TACT/F-111 Program. The TACT wing, except for the wing box, was subsequently replaced with the The "12-probe" boundary-layer rake was installed MAW. Modifications were made to the TACT wing on the upper wing surface at T/ = 0.76 (see Fig. 2(a)).
planform to accommodate installation of the smooth-
The leading edge of the center probes was at x/c ~
skin leading- and trailing-edge variable-camber sys- 0.96, where xis the streamwise coordinate and c is the tems. The wing design coordinates at a 26" wing sweep streamwise local chord. Three impact pressures were
for the 1-g cruise MAW with 8LE/TE = 0/2, and the
measured at each probe height. For 3 flights, 31 upper wing-splash coordinates at orifice row semispan loca-
surface orifice transducers from rows T/ = 0.76 and 0.59
tion (ri) = 0.76 and 8LE/TE = 0/2 are given in Ref. 19.
and 5 spare wing transducers were connected instead to Selected MAW design coordinates and the correspond- the boundary-layer rake shown in Fig. 2(b ). Local flow ing coordinates from the 1/12-scale wind-tunnel model direction was calculated using the pressures from the are compared in Ref. 20. The variable-camber leading- two side probes (cut at 45°) and the calibration tech- and trailing-edge surfaces of the MAW are illustrated nique described in Ref. 22. Because of transducer prob- in Fig. l(b ). Note that the positive direction is down.
lems, the pressures at the perpendicular distance above
the upper wing surface (Y) = 0.03 in. and 4.47 in. (see
Instrumentation Fig. 2(b)) were not measured.
Wing Pressure Orifices Flight Deflection Measurement System For this study, the pressure instrumentation was lo- The electro-optical FDMS used in this study was . d . l 23 24 cated on the right wing of the AFTI/F-111 airplane an updated version of a system use prev10us y. ' (see Fig. 2(a)). A detailed discussion of the static The MAW FDMS consisted of a control unit, a re- pressure instrumentation is found in Ref. 21. There ceiver, a target driver, and 13 infrared light-emitting were 152 flush-surface static pressure orifices located diode targets (Fig. 2(c)). The targets were mounted on on the upper and lower surfaces in four chordwise rows points of structural interest on the lower surface of the aligned with the free-stream airflow at a leading-edge left wing. The receiver was mounted behind a window sweep back angle of 26°. The pressure orifices were panel in the left side of the fuselage below the wing.
spaced so that the closest spacing was in the mid- From this receiver location all targets could be viewed section of the upper wing surface. The number of for all camber settings when the wing was swept to 26°.
upper and lower orifices at each semispan station is The control unit and target driver were mounted on the presented in Fig. 2(a). Nine of the 10 pressure trans- right-hand instrumentation pallet located in what had ducer boxes were located inside the flexible leading- been the weapons bay.
and trailing-edge flap· surfaces. This required connect- The FDMS control unit used the end-of-frame pulse ing the leading- and trailing-edge surfaces with a flexi- from the pulse code modulation (PCM) system as a ble fluorosilicon tubing (0.07-in. inside diameter). The synchronization signal. The control unit would com- orifices located on the surface of the wing box were con- mand the target driver to momentarily energize each nected by stainless steel tubing (0.12-in. inside diame- target in order. Prior to the illumination of each tar- ter). In most cases the length of the pressure lines from get the control unit would initiate a sequence within the orifices to the transducers was limited to less than the receiver. This sequence involved clearing its linear 5 ft, thus pressure lag effects were minimized. More de- diode array, scanning the array to sample the back- tails about the pressure orifices can be found in Ref. 19.
ground light signature and then scanning again with Buffet Accelerometers the target on. This process was necessary to accom- plish the automatic background light compensation.
Figure 2(a) shows the right wingtip locations for the The background light signal was used to modify the high-frequency normal accelerometers used in the buf- target light signal to improve system operating range fet analysis. The locations for the cockpit, left wingtip and tolerance of ambient light. Each target data sam- and horizontal tail accelerometers are given in Ref. 9.
ple was transferred to the PCM system as two 10-bit The accelerometers used only for the high-frequency digital words. One word contained target identification analysis were filtered in the airplane instrumentation and error messages, while the other carried the target package to remove the low-frequency maneuver com- position data. References 6 and 25 provide more com- ponent (for example, the wingtip and horizontal tail prehensive information on the FDMS.
normal accelerometers, anwr and anHr, respectively).
Aircraft Measurements necessary, this provided an optimum working environ- ment for the optical measurement system. Most of the Free-stream flight parameters, Mach number (M ), data were obtained for stabilized or quasi-stabilized static pressure (p ), dynamic pressure (q ), and 00 00 times because of concerns that spurious signals from angle of attack ( a, or) and angle of sideslip ((3) the FDMS could possibly affect the high-frequency re- were measured and derived from sensors installed on sponse instrumentation.
the AFTI/F-111 airdata boom. Mach number data Buffet Data from a modified MA-1 type uncompensated pitot-static probe were corrected for position error. Angle of at- The buffet analysis primarily consists of determining tack and angle of sideslip were measured using a flight- the root mean square (rms) value of the buffet accelera- path accelerometer vane system. Angle of attack is tions for increasing angles of attack. The low-frequency referenced to the wing reference plane for consistency maneuver components for the wingtip accelerometer with the wind-tunnel data. Because the noseboom is are filtered in the airplane instrumentation package, canted 2.5°-down relative to the vehicle body axis, and leaving only the high-frequency response. The rms the MAW is set to an angle 1.0°-up relative to the values and power spectral density estimates were then vehicle body axis, a 3.5° correction was added to the computed. The fluctuating accelerations were analyzed indicated vane angle of attack to obtain the indicated for continuous 1-sec time segments during periods of angle of attack (a). This angle of attack was corrected increasing angle of attack. The rms value of the accel- for pitching moment and upwash effects to obtain erations for each continuous time segment are shown as a:r. This corrected angle of attack was used for the buffet loads in the data figures. Power spectral density wing pressure data analysis.
techniques indicate the power and frequency distribu- All the instrumented parameters were recorded dig- tion of the buffet parameters. The natural frequencies of the primary structure for the wing and the frequen- itally on an airborne PCM system. The PCM system cies obtained from the wingtip accelerometer analysis had a sampling rate of 20 to 800 samples/sec. Each wing surface and boundary-layer pressure was sampled showed good agreement. This agreement lends confi- at 20 samples/sec. Each high-frequency accelerometer dence in the instrumentation installation and analysis used for the buffet study was low-pass-filtered on the techniques (see Ref. 28).
airplane at 160 Hz. The system sampling rate for the The buffet intensity rise was defined as the point FDMS data channel was 200 samples/sec. This means where the rms buffet component of normal accelera- that with the 13 FDMS targets installed and 3 spare tion (aan ) begins to increase rapidly with respect WT channels, each target was sampled 12.5 times a second.
to increasing airplane normal-force coefficient (CNA) (knee of C NA as a function of a an curve). This is Analysis Techniques WT discussed in a later section.
Pressure Data Test Points The data used for the surface and boundary-layer Flight data presented in this paper are for a Mach pressures were chosen from stabilized and quasi- number of 0.85 and a 3.0° to 14.0° angle-of-attack stabilized flight conditions to minimize concerns about range. Free-stream dynamic pressures were 300 and pressure and PCM sampling lag. When selecting data 600 lb/ft , with most of the data being for 300 lb/ft .
for analysis, maximum deviations from the desired In most cases, data were selected for analysis for an- flight conditions for M and ay were 0.01 and 0.25°, gles of sideslip near 0°. Flight Reynolds number was respectively. For the boundary-layer data, the Mach 6 1 6 approximately 2.3 x 10 ft- (26 x 10 based on the number and velocity calculations used the assumptions mean aerodynamic chord (cMAC) = 11.2 ft).
(1) that the local static pressure was constant through the boundary layer and (2) that total temperature was Test Maneuvers constant through the boundary layer and equal to the free-stream value. The local static pressure was the The diverse nature of the research objectives in this
surface static pressure at x/c = 0.96 (directly ahead of
flight test program resulted in the use of several types of the rake).
flight maneuvers. Wing pressures and boundary-layer profiles required slow controlled windup turns to mini- FDMS Data mize concerns regarding pressure and PCM system lag.
The data used for the FDMS analysis were from left- The aircraft would be stabilized at the desired Mach hand turns, since this meant the left wing was down number and altitude before entering the windup turn.
and looking at a darkened background ( dark com- During an ideal slow windup turn, the aircraft was sta- pared to the sky). The left-hand turn also avoided bilized for a few seconds or longer at each desired Mach including the Sun in the background. Although not number and angle-of-attack combination. Because of thrust limitions, it was not possible to hold altitude trim or specified start conditions during the setup of and Mach number constant at the higher angles of data runs.
attack (> 8.0°) and/or flap settings. Windup turns Results And Discussion performed for buffet and loads measurements were at a higher turn rate because accelerometers and strain Span Effects on Pressure Distribution gages were not susceptible to pressure lag. The FDMS data were obtained during the pressure maneuvers to In Fig. 5, the chordwise pressure coefficient ( Gp) dis- correlate the FDMS and pressure data.
tributions as a function of x/c are shown at the four semispan stations for M = 0.85, q ~ 300 lb/ft , 00 00 Wing Pressure
8LE/TE = 0/2 (baseline camber configuration), and
Figure 3 illustrates the slow windup-tum maneu- for ar = 8.0° and 10.0°. These two angles of attack ver and the response of one pressure transducer and were selected to show the effects of trailing-edge flow a wingtip accelerometer. During the initial portion of conditions on the pressure profiles. The a:r = 8.0° this particular maneuver, the pilot was adjusting al- data have good presssure recovery at the trailing edge, titude to achieve the desired conditions later in the and the midspan profiles are typical of supercritical maneuver. The initial part of the maneuver was un- airfoils at or near the wing design conditions. All the steady in nature, but developed into a steady turn upper surface pressure profiles have a strong negative culminating at approximately 80 sec at the maximum pressure peak at the leading edge; however, for the aim angle-of-attack and dynamic pressure values, with midspan rows (11 = 0.59 and 0.76) for ar = 8.0° the the Mach number within acceptable limits (see Analy- peak is followed by nearly constant Gp plateaus. For sis Techniques section). The boundary-layer data and ar = 8.0° the Gp profiles "shock down" from approx- FDMS data were gathered using this type of windup imately x/c ~ 0.40 at 11 = 0.93 to x/c ~ 0.70 at 11 = turn. Soon after reaching the desired conditions, the 0.40. Following the aft shocks, the Gp values indicate onset of separation is indicated by the traces for the a region of re compression that continues to the trailing wingtip accelerometer and the trailing-edge pressure edge. At the trailing edge, the pressures show good orifice at x/c = 0.96 and 11 = 0.93; then the maneuver recovery for all the semispan stations. The Gp profiles is terminated.
for ar = 10.0° also show strong leading-edge negative pressure peaks. The midspan Gp plateaus have dis- Buffet appeared with the movement of the aft shock forward Figure 4 illustrates a windup-tum maneuver for the consequent to the separation of the boundary layer at baseline configuration thE/TE = 0/2. This was a typ- the trailing edge. The shading on the trailing edge ical windup-tum maneuver used for the buffet evalu- of the wing represents an approximate region of sepa- ation. This maneuver was started at trim and con- rated flow determined by analysis of the chord wise Gp tinued to maximum allowable angle of attack. Mach profiles for a:r = 10.0°.
number was held nearly constant and altitude and dy- Figure 6 illustrates the Gp profiles for the same namic pressure were sa~rificed where available thrust Mach number and angles of attack as Fig. 5 but for was limited. As angle of attack is increased, there is a
8LE/TE = 5/10. The leading-edge camber of 5° pro-
sudden increase in buffet (time ~ 33 sec, a ~ 10.0°), duces a rounded leading-edge Gp profile. The midspan known as the buffet intensity rise (BIR). This BIR Gp profiles are semiflat, followed by aft shock recom- for the wingtip is followed by initial buffet at the pi- pression near the trailing-edge-flap line (x/c ~ 0.70).
lot's station (time ~ 34 sec). Next are the simultane- The wingtip row (11 = 0.93) is similar except for the ous onset of wing rock and BIR for the horizontal tail large negative pressure area aft of the flap line, fol- (time~ 38 sec). Buffet characteristics for the MAW lowed by a large secondary velocity peak. As in Fig. 5, are discussed in more detail in Ref. 9.
the aft shock has moved toward the leading edge as a:r Other Types of Maneuvers increased to 10.0°. The larger trailing edge deflection angle and angle of attack, both 10.0°, combine to in- Pushover pullup (POPU) maneuvers were used to crease the approximate region of separated flow shown, gather data for many of the tests points in the per- which is indicated by the shaded area.
formance part of the program. This type of maneuver is usually rapid in nature but will keep the aircraft Angle-of-Attack Effects on Aft Shock Location near the initial premaneuver flight conditions. It was Figure 7(a) shows the relationship with o:r of aft generally not possible to pause and hold angle of at- shock location (see Ref. 19 for discussion) along row tack. Another maneuver was the level acceleration used 11 = 0.76 for M = 0.85, q ~ 300 lb/ft , and 00 00 primarily for performance and evaluation of the auto-
8LE/TE = 0/2. As angle of attack increases from
matic control modes. In addition, all the disciplines 6.0° to approximately 8.5°, the location moves from used data from any suitable maneuver, including the approximately 30-percent x/c to approximately 60- has been shown to be associated with the larger flap
r
9 I percent x/c. As angle of attack continues to increase, settings. Also appearing (for time > 110 sec) in the the shock location then begins moving forward again to anwr trace and the trailing-edge pressure traces of rows approximately 30-percent x / c as O.T increases to 11.0°.
11 = 0.93 and 0.40 are flucuations that are a function The windup-tum time history in Fig. 7(b) illustrates of the variations in angle of attack.
the effect on the orifice pressures of aft shock movement Wingtip Twist Effects over the upper surface of the wing for the same flight conditions shown in Fig. 7(a). In Fig. 7(b), absolute The incremental change in wingtip twist caused by pressures from six representative locations are shown load (wingtip delta twist) was calculated as the dif- plotted as a function of time. One is from a wing ori- ference between the changes in deflections of the for- fice near the leading-edge area (x/c = 9 percent), four ward and aft wingtip targets. This wingtip delta twist are from the midsection (x/c = 37, 47, 56, and 59 per- is the incremental twist caused by load, not the total cent) and one is near the trailing edge (x/c = 96 per- twist. The wingtip delta twist, in degrees, is shown cent). As angle of attack increases during the windup in Fig. 8(a) as a function of free-stream dynamic pres-- tum, the aft shock moves rearward over the orifices sure for three camber settings at M = 0.85. The for the midchord pressures and reaches approximately trailing-edge-up twist is negative. The close agreement 60-percent x/c, then it retraces its movement forward. between the data for OLE/TE = 0/2 and 10/2 show The traces in Fig. 7(b) show that as the aft shock moves that the effect of leading-edge camber on wingtip twist rearward, lower orifice pressure is measured in a region is insignificant compared to that of trailing-edge cam- between the strong negative pressure peak at the wing ber. The FLEXSTAB predictions reported in Ref. 30 leading edge and the aft shock. The aft shock never show the same trend. Only one data point was avail- reaches the trailing-edge orifice (x/c = 96 percent) but able for OLE/TE= 5/10 at q f'::/ 600 lb/ft . The verti- there are indications of disturbed flow for o.T > 9.0°, cal spread in the groups of measured wingtip twists is which can be noted by a high-frequency content in the caused by the variation in aircraft angle of attack. For a pressure traces between the aft shock and the trailing given dynamic pressure and Mach number, an increase edge. This may result from disturbances at the base of in angle of attack causes a corresponding increase in the aft shock and from the beginning of trailing-edge load factor, which in turn directly affects the wingtip separation. delta twist. Thus, increasing aircraft angle of attack causes more negative wingtip twist, which is also re- Figures 7(c) and (d) present a detailed compari- ferred to as "washout". The resulting local angle of at- son of two windup-turn time histories at M f'::/ 0.85 tack that the wingtip experiences is therefore less than showing pressure traces for all four trailing-edge ori- the aircraft angle of attack for a positive normal accel- fices and their associated traces of angles of attack and eration maneuver. This wingtip washout may explain wingtip accelometers for OLE/TE = 0/2 and 5/10. In why the wingtip pressure traces (11 = 0.93) in Figs.
attached flow at the trailing edge, the flow should re- 7( c) and (d) differ in the separation indicated. In Fig.
cover to the free-stream static pressure, but a decreas- 7( c) a washout of 1.5° or larger could delay separation ing pressure indicates nonrecovery to free-stream static enough to show little, if any, effect. But in Fig. 7(d) pressure and separated flow at the trailing edge. In the 10° trailing-edge flap plus the angle of attack of Fig. 7(c) for OLE/TE= 0/2 the trailing-edge traces are the windup turn would be large enough to overcome smooth until approximately 9.0° angle of attack. Be- any washout indicated in Fig. 8(a) for OLE/TE = 5/10.
yond this angle of attack, three of the four pressures Thus, the wingtip pressure trace indicates trailing-edge
(11 = 0.76, 0.59, and 0.40) break toward lower pressure
separation.
values producing an indication of trailing-edge sepa- ration. The wingtip accelerometer trace, anwr, also Figures 8(b) and (c) compare wing surface pressure closely correlates with the trailing-edge pressures, both profiles at two span locations (TJ = 0.93 and 0.76) at fluctuating (time > 60 sec) with small changes in angle M = 0.85, 0.T = 8.0°, OLE/TE = 0/2, and q ~ 00 00 of attack. For OLE/TE= 5/10 (Fig. 7(d)), the two mid- 300 and 600 lb/ft • In the figures, the wingtip twist wing trailing-edge pressures (11 = 0.76 and 0.59) have effects, if any, are minimal at the inboard row location a pronounced break to a lower pressure near o. = 8.0°. while the outboard row shows only a small difference in As the windup turn continues (time >100 sec), the the wingtip delta twist for the two dynamic pressures.
aircraft begins to lose altitude rapidly because of the The wingtip delta twist is only 0.5° for OLE/TE= 0/2, higher drag from the larger flap settings (the altitude which would suggest that large effects would not be trace is not shown). This results in the higher pres- anticipated for the two dynamic pressures investigated.
sures observed for all four trailing-edge pressures. The Buffet Intensity anwr trace indicates an increased level of activity (see Fig. 7(d)) compared with the OLE/TE= 0/2 case over Figure 9 presents the normal-force and buffet in- the entire windup turn. This residual buffet "buzz" tensity characteristics for the baseline configuration, is attached and well-behaved. For a:r > 8.6°, the buf-
OLE/TE = 0/2, and for the 8LE/TE = 5/10 config-
fet data, as well as the pressure and velocity profiles,
uration. The 8LE/TE = 5/10 configuration is re-
garded as one of the better fixed-flap configurations indicate the flow is separated.
for transonic maneuvering. The normal-force curves In Fig. lO(d), information pertinent to the data in
for the 8LE/TE = 0/2 and 5/10 configurations have
Figs. IO(a), (b), and (c) are presented as a function of breaks that imply the presence of significant areas of a:r. The GPTE and GNA. curves are repeated, and anwT flow separation on the wing. These breaks occur at a is now shown as a function of o:r. The curves for aft
CNA. ~ 0.80 and a:= 9.5° for the 8LE/TE = 0/2 config-
shock position in percent of x/c and the ratio of local uration (Fig. 9(a)), and at a GNA. ~ 1.00 and a:~ 10.3°
velocity to edge velocity (U /Ue) values for Y = 2 in. are
for the 8LE/TE = 5/10 configuration (Fig. 9(b)). The
also shown. All of the pressure data are for rJ = 0. 76.
difference in the GNA. values at the normal-force break The shaded band at approximately o: = 8.5° indicates (approximately 0.20 G NA) indicates the influence of the the region of incipient separation. Good agreement was wing trailing-edge deflection on the coefficient data.
found between all the data sources (trailing-edge pres- The buffet intensity data (GNA. as a function of sure, aft shock location, boundary-layer velocity ratio, aan ) indicate slightly lower BIR values in terms of normal-force coefficient, and therms of the buffet).
cN:than the normal-force-break data. Similar im- In Fig. 11, the pressure and buffet characteristics for provements in the BIR and the intensity characteris- M = 0.85 and q ~ 300 lb/ft are presented for 00 00 tics are shown for the 8LE/TE = 5/10 configuration,
8LE/TE = 5/10. The upper and lower surface pressure
with respect to the normal-force curves. However, for profiles (Fig. ll(a)) are again for o:r = 5.0°, 6.0°, 8.0°, the 8LE/TE = 5/10 configuration and low GNA. values, and 10.0°. Boundary-layer data were not obtained for the intensity data (aanwT) indicate a large offset when M = 0.85. However, the boundary-layer data ob-
compared with the lhE/TE = 0/2 data. This offset in-
tained for Moc, = 0.80 showed separated flow at the
dicates a low-level separation occurring before the BIR trailing edge for all angles of attack studied (o:r = 4.0° with a maximum value of aan ~ 0.25. A similiar WT to 8.0). Because trailing-edge flow separation occurs at comparison for the 8LE/TE = 0/2 configuration indi- lower angles of attack as Mach number increases, the cates a aan ~ 0.06. The offsets are pointed out in
flow at the trailing edge for M = 0.85 would also be
WT 00 Fig. 9.
separated. Figure ll(b) shows GPTE as a function of a:r and Fig. 11 ( c) shows the buffet characteristics. The Summary of Pressure and Buffet pressure profiles show the expected rearward movement Characteristics of aft shock location as o:r increases from 5.0° to 8.0°.
In Fig. 10, the pressure and buffet characteristics However, unlike the 8LE/TE = 0/2 profiles, the o:r =
for M = 0.85 and q ~ 300 lb/ft are presented
5.0° and 6.0° profiles have a secondary velocity peak
for {hE/TE = 0/2. The upper and lower surface pres-
at x/c ~ 0.75. For o:r = 10.0°, the forward movement
sure profiles for T/ = 0.76 are given in Fig. l0(a) for
of the aft trailing shock indicates that the flow charac-
a:r = 5.0°, 6.0°, 8.0°, and 12.0°. The pressure coeffi-
teristics over the wing have changed, and there is the
cient on the wing upper surface at x/c = 0.96 (GPTE)
possibility that the wing has separated flow. None of as a function of a:r is also shown. In Fig. lO(b ), the
the pressure profiles have a recovery to Gp = 0 at the
boundary-layer velocity profiles for rJ = 0.76 and x/c =
trailing edge, which supports the trailing-edge separa- 0.96 are shown for a:r from 5.2° to 8. 7°. The airplane tion indicated by the boundary-layer data. The CPTE normal-force coefficient and buffet intensity are given curve for o:r = 4.0° to 10.0° does not have the well- in Fig. lO(c). The pressure profiles show the expected
defined break of the 8LE/TE = 0/2 data, and therefore
rearward movement of the aft shock location over a cannot be easily used to obtain the angle of attack for supercritical airfoil as a:r increases to 8.0°. The Gp wing separation. The less negative values occurring at
for o:r = 8.0° shows a well-developed supercritical dis-
approximately o:r = 7.0° are a result of the secondary tribution. The pressure profiles indicate that separa- shock. From the buffet intensity data in Fig. ll(c), the
tion at the trailing edge occurs between o:r = 8.0° and
BIR occurs for o:r ~ 9.6°, which supports the possible 12.0°. From the break in the curve for GPTE• separation wing separation observed for the o:r ~ 10.0° pressure is seen to occur for o:r ~ 8.6°. The velocity profiles for profile. The offset in the buffet intensity data and the a:r :S 8.0° show larger losses as a:r increases from 5.2° low level of "buzz" seen in the anwr time history in
to 8.0°. For a:r = 8.6° and 8.7°, the velocity profiles
Fig. 7(d) support the trailing-edge flow separation ob- show incipient separation for the flow at the trailing served for the pressure data. It is apparent from these edge. The buffet data, by the break in the C NA and figures that the global (buffet data) and the local data the BIR point, also show that separation occurs for (wing pressure data) do not have the clear-cut interre- a:r ~ 8.6°. For a:r < 8.6°, the buffet data, as well as lationship observed for the IJLE/TE = 0/2 data.
the pressure and velocity profiles, indicate that the flow Figure ll(d) presents information similar to that shock position, is correlated with the initial separation shown in Fig. 10( d), except that the pressure-derived provided by the buffet analysis and the boundary-layer section normal-force coefficient (c.,..) for T/ = 0.76 is velocity profiles. The wingtip twist measurements pro- shown instead of the boundary-layer velocity ratios. vided an insight into how dynamic pressures for pos- All of the pressure data are for T/ = O. 76. All of the itive normal accelerations affect the wingtip pressure curves derived from the pressure data (trailing-edge profiles.
pressure, aft shock location, and section c.,..) indicate For the well-behaved flow of the 8 LE/TE = 0/2 cam-
changes in the flow at approximately o.r = 7.0°. None
ber, which is a typical cruise camber shape, the local of these curves have a definite break that would indi- and global data are in good agreement with respect to cate an extensive region of wing flow separation. The the flow properties of the wing. This good agreement is section c.,.. curve indicates that the wing is still perform- not observed for the 8LE/TE = 5/10 camber, which is ing well as o.r increases from 7.0° to 10.0°, and that
a maneuvering camber shape. For the 8LE/TE = 5/10
an extensive region of wing flow separation may occur camber, the local and global data have similar trends for o:r > 10.0°. This agrees with the buffet data. The and conclusions.
shaded band at approximately a: = 10.0° indicates the region where extensive separation begins.
References
For the well-behaved flow of the 8LE/TE = 0/2 cam-
Sisk, Thomas R., Friend, Edward L., Carr,
I
ber, which is a typical cruise camber shape, the lo- Peter C., and Sakamoto, Glenn M., Use of Maneu-
I
cal and global data are in excellent agreement with ver Flaps to Enhance the Transonic Maneuverability respect to the flow properties of the wing. This excel- of Fighter Aircraft, NASA TM-X-2844, 1973.
lent agreement is not observed for the 8LE/TE = 5/10
camber, which is a maneuvering camber shape. For Gould, D.K., Boeing Pre-Design of AFTI-111 Mis- the 8LE/TE = 5/10 camber, the local and global data sion Adaptive Wing, Volume II-Aerodynamic Trade have similar trends and conclusions but not the clear- Studies, Theoretical Calculations and Wind Tunnel cut agreement for the breakpoint as observed for the Tests, AFFDL-TR- 78- 73, June 1978.
OLE/TE = 0/2 camber. A possible reason that the lo- Nelson, D.W. and Letsinger, Gary R., AFTI- cal and global breakpoints are not aligned is because F-111 Mission Adaptive Wing Wind Tunnel Analy- of the presence of a secondary velocity peak observed sis Report, Arnold Engineering Development Center
for 0:T = 5.0° and 6.0° in Fig. ll(a), and for 0:T = 8.0°
PWT-16T, Test No. TF550, Boeing Doc. No. D365- and 10.0° in Fig. 6.
10O58-1, rev. A, Apr. 1981.
Concluding Remarks Hall, Joseph M., AFTI/F-111 Flight Control Sys- tem, AFWAL-TR-87-3O12,1987.
Selected results from the wing surface and boundary- layer pressures, flight deflection measurement system Steers, Louis L. and Bussing, Paul R., "Flight (FDMS) and buffet studies for the advanced fighter Demonstration and Research of the Smooth Variable- technology integration (AFTl)/F-111 mission adaptive Camber Wing," Advanced Fighter Technology Integra- wing (MAW) Program were presented and discussed tion F-111 Mission Adaptive Wing, NASA CP-3055, with respect to each other. The discussions mainly 1990, pp. 71-97.
concerned data for a Mach number of 0.85, and leading- Bonnema, Kenneth L. and Lokos, William A.,
and trailing-edge camber deflections of (6LE/TE) = 0/2
"AFTI/F-111 Mission Adaptive Wing Flight Test In- and 5/10.
strumentation Overview," Paper No. 89-0084, Instru- From a flight test perspective, providing the techni- ment Society of America, May 1989.
cal tools to describe the advantages of a supercritical Advanced Fighter Technology Integration F-111 wing for different cambers is very challenging. This Mission Adaptive Wing, NASA CP-3O55, 1990.
paper describes the different aerodynamic technologies studied on the airplane, and their relationship with Webb, Lannie D., McCain, William E., and Rose, each other.
Lucinda A., Measured and Predicted Pressure Distri- butions on the AFTI/F-111 Mission Adaptive Wing, The pressure profiles had the distribution typical of NASA TM-10O443, 1988. Also published as AIAA-88- a supercritical airfoil for the 8LE/TE = 0/2 and 5/10 2555, June 1988.
cambers investigated in this paper. The midspan pres- sure profiles for both cambers illustrated the nearly Friend, Edward L. and Thompson, Jeffrey M., "Buf- constant upper surface pressure coefficient plateaus ex- fet Characteristics of the Advanced Fighter Technology pected for supercritical wings. The analysis, in terms Integration (AFTI)/F-111 Airplane with the Mission of pressure profiles with respect to angle of attack and Adaptive Wing," Advanced Fighter Technology Integra- F-111 Mi,ssion Adaptive Wing, NASA CP-3055, 1990, tion F-111 Mi,ssion Adaptive Wing, NASA CP-3055, pp. 157-195.
1990, pp. 197-222.
Bussing, P.R., AFT! F-111 Final Instrumentation Wong, Kent J., AFTI/F-111 Mi,ssion Adaptive Report, Boeing Doc. No. D365-10016-2, Dec. 1984.
Wing Lift and Drag Test Results Volume I, AFFTC- Dudzinski, Thomas J. and Krause, Lloyd N ., Flow- TR-87-02, Apr. 1987.
Direction Measurement with Fixed-Position Probes, Phillips Paul W., AFTI/F-111 Mi,ssion Adap- NASA TM-X-1904, 1969.
tive Wing (MAW} Automatic Flight Control System DeAngelis, V.M., "In-Fight Deflection Measure- Modes Lift and Drag Characteristics, AFFTC-TR-89- ment of the HiMAT Aeroelastically Tailored Wing," 03, Feb. 1989.
AIAA-81-2450, Nov. 1981.
Painter, Weneth D. and Caw, Lawrence J., Design Lokos, William A., Predicted and Measured In- and Physical Characteristics of the Transonic Aircraft Flight Wing Deflections of a Forward-Swept- Wing Air- Technology (TACT) Research Aircraft, NASA TM- craft, NASA TM-4245, 1990.
56048, 1979.
DeAngelis, V. Michael and Fadale, Robert, Symposium on Transonic Aircraft Technology "Electro-Optical Flight Deflection Measurement Sys- (TACT}, AFFDL-TR-78-100, Aug. 1978.
tem," SFTE 18th Annual Symposium Proceed- Supercritical Wing Technology: A Progress Report ings, SFTE Technical Paper 22, Sept.-Oct. 1987, on Flight Evaluations, NASA SP-301, 1972. pp. 22-1-22-14.
15 26 Ayers, Theordore G. and Hallissy, James B., His- Webb, Lannie D. and Washington, Harold P., torical Background and Design Evolution of the Tran- Flight Calibration of Compensated and Uncompen- sonic Aircraft Technology Supercritical Wing, NASA sated Pitot-Static Airspeed Probes and Application TM-81365, 1981. of the Probes to Supersonic Cruise Vehicles, NASA TN D-6827, 1972.
Baldwin, A. Wayne, Kinsey, Don W., and Lash, Stanley F., Transonic Aircraft Technology Summary, Sakamoto, Glenn M., Aerodynamic Characteris- AFFDL-TM-78-7-FXS, Jan. 1978.
tics of a Vane Flow Angularity Sensor System Ca- pable of Measuring Flight Path Accelerations For Fehl, John E., AFTI/F-111 Mission Adaptive The Mach Number Range From 0.40 to 2.54, NASA Wing 1/12 Scale Wind Tunnel Model Inspection, TN D-8242, 1976.
AFWAL-TM-80-114-FIMS, Dec. 1980.
Friend, Edward L. and Matheny, Neil W., Prelimi- Gould, Douglas K., "AFTI/F-111 Mission Adap- nary Flight Measurements of the Buffet Characteristics tive Wing," Advanced Fighter Technology Integration of Prototype Lightweight Fighter Aircraft, NASA TM- F-111 Mission Adaptive Wing, NASA CP-3055, 1990, X-3549, 1977.
pp. 29-69.
Stanewsky, E. and Basler, D., "Experimental In- Webb, Lannie D., Powers, Sheryll Goecke, and vestigation of Buffet Onset and Penetration on a Super- Rose, Lucinda A., "Selected Local Flow-Field Measure- critical Airfoil at Transonic Speeds," Aircraft Dynamic ments on the Advanced Fighter Technology Integration Loads due to Flow Separation, AGARD-CP-483, 1990, (AFTI)/F-111 Aircraft Mission Adaptive Wing," Ad- pp. 4-1-4-11.
vanced Fighter Technology Integration F-111 Mission Adaptive Wing, NASA CP-3055, 1990, pp. 115--156. Nelson, D.W., AFTI/F-111 Aerodynamics Final Report, Boeing Doc. No. D365-10110-1, Feb. 1987.
McCain, William E., "Comparison of Two Pre- diction Methods with Flight-Measured Wing Sur- Friend, Edward L. and Sakamoto, Glenn M., Flight face Pressure Distributions from the Mission Adap- Comparison of the Transonic Agility of the F-11 lA tive Wing," Advanced Fighter Technology Integration Airplane and the F-111 Supercritical Wing Airplane, NASA TP-1368, 1978.
EC85 33205-.017 (a) Airplane in flight. Chordwise dark areas on the right wing indicate the four semispan locations of pressure orifices.
Fig. 1 AFTI/F-111 MAW airplane and wing shape.
Wing box reference line Flex panel Rear spar 920412 (b) The MAW smooth variable-camber flap shape.
Fig. 1 Concluded.
♦ Accelerometer locations rAircraft £. Boundary-layer rake location / centerline □ Transducer box locations Static orifice locations
-
- ----
26°
----
-
-
Leading-edge flap Reference Tl= 0.40 Tl= 0.59 Tl= o.76 Tl= o.93
I I
Number of 31 15 Upper surface 16 40 pressure 12 14 10 Lower surface orifices 920413 (a) Semispan locations of surface pressure orifices, boundary-layer rake, pressure instrumentation, and wingtip accelerometers for the right wing.
Fig. 2 Experiment locations and description.
---
,,,,,..,,, - ......
Silver / ' / solder I Tip ~1in.~--""T'"'"- .....
, detail y ' / , ..... _____ .,,,,.,,, 4.96 Stainless steel tubing, / 4.47* 0.03 in. outside diameter 1 3.99
T-=====t==
/ 3.49 1 2.99 5 in.
-===~~
. h <' 2.33
k b h Ra e pro e e1g ts , _ 1 96 \ 1.47 , 0.97 \ 0.47 , 0.24 \Q.03* *Probes not functional (b) Boundary-layer rake.
Receiver Targets (13) 920414 ( c) FDMS target locations on the lower surface of the left wing.
Fig. 2 Concluded.
a, deg 31.9 X 103 Attl::,de, :: ::k=== ._ __ _,_ ___ ...... : __ -=======--L------L---=--=.-=-=-=~--'-----'-----"=---' ___ __.
qoo, lb/ft2 Moo .85
.86~-~· ... ·······~
.84 .___ __ _.__ __ --'-------''-----....__ __ _.._ __ ___._ ___ ..__ __ _.__ __ __.
p, lb/f1 at 11 = 0.76 x/c = 0.96 Time, sec 920415 Fig. 3 Time history of typical windup-turn maneuver used to obtain wing pressure FDMS and boundary-layer data; M ~ 0.85, Qoo ~ 300 lb/ft , and OLE/TE= 0/2.
a, deg p, deg Roll
rate, 4: ~
deg/sec -40 BIR a "ckpt'
i• •• •·•• I•.•· ••••ut1Jat 1'
II - - g
2-:b-
BIR an WT' .... ,. .. ~,,.,. ,,1,•J•*DJ*f:Jna,_,
1: l
g -10 an HT' 1 t
10 r
o••• •
g -10 0 8 16 24 32 40 48 56 64 Time, sec 920416 Fig. 4 Time history of typical windup-turn maneuver used to obtain buffet and FDMS data; M ~ 0.85, q ~ 00 00 300 lb/ft , and 8LE/TE = 0/2.
•
■ -1.4 -1.2 separation -1.0
for ar = 10°
-.8 .2 .4 .6 .8 x/c 920417 Fig. 5 Steady chordwise pressure distributions at four semispan locations for M = 0.85, q ~ 300 lb/ft , and 00 00 6LE/TE = 0/2. No separation at ar = 8.0°.
•
■ -1.4 -1.2 Area of separation
for aT = 8° and 10°
-.6 -.4 .2 .4 .6 .8 x/c 920418 Fig. 6 Steady chordwise pressure distributions at four semispan locations for M = 0.85, q :::::: 300 lb/ft , and OLE/TE= 5/10.
Flight 1 Flight 2 □ Aft shock location, X/c, percent 0 ..._ _______ ......_ ______ ___.
4 10 12 920419 (a) Relationship of aft shock location with angle of attack; M = 0.85, 8LE/TE = 0/2, and T/ = 0.76.
Fig. 7 Aft shock position as a function of angle of attack and windup-turn time histories of selected chordwise pressures, trailing-edge pressures, and wingtip accelerations, q ~ 300 lb /ft .
Percent p, of chord lb/ft 2 x/c 6x 10
A trailln~-edge p'."'sure •==
: t :
I ' -·~-. : · -=
:t==
,,.;:;::;:( t::\, Forward movement
of aft shock 56 4 47 4
: ~-,!•i-//-/'_'=· -~ M_i_d_c_ho_r_d_p __ r_e_s_s-u-re~s--~~
····•:•:-:,:.· [ A leading-edge pressure
:~-~------------~
0 10 20 30 40 50 60 70 80 90 Time, sec 50 -- 50 -- 70 ----- 70 -- ao ----- go -----120 Range of a.T 920420 (b) Windup-tum time history of selected chordwise pressures showing movement of shock location; M ~ 0.85, 8LE/TE = 0/2, and 77 = 0.76.
Fig. 7 Continued.
a, deg
:: r ___ - -----------a~
5~· : ~
I I
p, 620 f--· I
lb/ft2
520r i -~ ~
at T\ 1:1 0.93 420 .__ __ _._ __ _.__ ___ ..__ __ _.__ __ _.__ _ _.! _ _,_ __ ____,_ ____ ~--~ I I p, lb/ft 2
at T\ = 0.76
p, lb/ft 2 at Tl= 0.59 P, lb/ft 2
at T\ = 0.40
10 20 30 40 50 60 70 80 90 Time, sec 920421 (c) Time history of trailing-edge wing pressures (x/c = 0.96) and wingtip accelerometers as angle of attack increases; M 00 ~ 0.85 and 8LE/TE = 0/2.
Fig. 7 Continued.
a: 8°
a, deg
'~
tI;:;;It=:····:;:i:\:;:i:····;ifffrt{!
p, lb/ft 2
at Tl= 0.93 ::~ i-f----·-: ---~--
::if:){:fffj:j:r:i:i:j:!:\:j:j:fI{il
at ~b~~o\6 ::~ ...... ~-·~:_-·_·_· :~·~~-~-~-~-•-i_i_i_li_ill_ili_lli~-~~-~-~~
p, lb/ft 2 at Tl= 0.59 p, lb/ft 2 at Tl= 0.40
: t O H -:-~-- ·M:•-•--:•·-•:•
0 10 20 30 40 50 60 70 80 90 100 110 120 130 140 Time, sec 920422 (d) Time history of trailing-edge wing pressures (x/c = 0.96) and wingtip accelerometers as angle of attack increases, Moo ~ 0.85 and OLE/TE = 5/10.
Fig. 7 Concluded.
22l OLE/TE, deg
0 0/2
0 10/2
~ 5/10 -1 -2 4.3--~ Wingtip 3.5- .....
delta 4- ..........
twist, 6-~ ...............
.....
.....
deg .....
.....
.....
.....
.....
-3 .....
.....
.....
.....
.....
.....
.....
.....
.....
.....
.....
.....
.....
.....
.....
.....
.....
-4 ----~-3.1 Trailing edge up -5 ._ _____ ...._ _____ __._ _____ _, _______ ..__ ____ __, 200 300 400 500 600 q , lb/ft 2 920423 (a) FDMS wingtip delta twist as a function of dynamic pressure for three cambers.
Fig. 8 Comparison of FDMS measured wingtip delta twist with two pressure profiles for M = 0.85 and q ~ 00 00 300 and 600 lb/ft • -1.2 -0- q = 300 lb/tt2 -0- Q = 600 lb/ft -1.0 -.8 -.6 -.4 Cp -.2
I
•
~
.2 .4 .6 .8 0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 x/c 920424 (b) bLE/TE = 0/2, T/ = 0.93, CJ.r = 8.0°.
-1.2 -0- Q = 300 lb/ft -0- q = 600 lb/ft 2 -1.0 -.8 -.6 -.4 Cp -.2 .2 .4 .6 .8 0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 x/c 920425 (c) bLE/TE = 0/2, T/ = 0.76, CI.T = 8.0°.
Fig. 8 Concluded.
J 1.4 1.2 ......
1.0
• ••
• • • ,I' ■ ••
•
••
•• • Break
••
•
.8 • ----1 ►► -·· CN in curve ■ • ,..-BIR A
••
•
•
• 0.06
.6 •
• • ◄ ..
(Offset) • ••
i
••
.4 • 2 4 12 6 8 10 14 .2 aT ,deg 0 4 8 12 16 0 2 a,deg 920426 (a) lhE/TE = 0/2.
1.4 1.2 I ......
;. ..
•••
• •• • ,r, ■ " Break .. ., 1.0
,11•·
in curve • .. ◄
BIR • .rl .8 0.25
•
-~~
CN ◄
•
'
A • (Offset) •
✓-
.6 •
•
■ •
I
.4 •
Ill 2 4 6 8 10 12 14 .2 aT, deg 0 4 8 12 16 0 1 2
a,deg a an ,g
WT 920427 (b) OLE/TE= 5/10.
Fig. 9 Variation of airplane and normal-force coefficient with angle of attack and buffet intensity for M = 0.85 and OLE/TE= 0/2 and 5/10.
-.30 uT deg -.20 0 5.20 CpTE -.10 D 6.02 ◊ 7.00 l::i..
8.01 .10 V 8.58 5 6 7 8 9 10 11 Iii 8.67 o:T, deg UT= 50 -2 Y, UT= 50 . , □ In .
_,,,:1··,·, "\...
UT= 90 ◊ 2 <?
',-;v.s-~..,.._., -c~ %...;.
-1 UT= 12° l::i..
"' " . "
, ... ,y;., ✓ :.;~,,..
\ ">.. '°I~,-"~~ Cp 1\-
/~-~7.: ~ "-"ilf~~~
. ... -
0 '-----'--...1...-1...--.....L..-.,._--J_...L-____,i .3 .4 .5 .6 .7 .8 .9 1.0 1.1 .2 .4 .6 .8 1.0 U/Ue x/c 920428 920429 (b) Boundary-layer profiles for several angles of attack
(a) Pressure profiles for several angles of attack, 'r/ =
0.76 (see inset for trailing-edge pressures). Solid sym- at x/c = 0.96 and 'r/ = 0.76.
bols are lower surface Gp.
1.2 1.0 Break .8 In curve CN .6 A .4 4 14 2 6 8 10 12 .2 uT, deg 0 4 8 12 16 0 1
u,deg cr an ,g
WT 920430 (c) Variation of airplane normal-force coefficient characteristics with angle of attack and buffet intensity.
Fig. 10 The angle-of-attack relationship between pressure coefficients, boundary-layer profiles, and buffet charac- teristics for M = 0.85, q ~ 300 lb/ft , and 8LE/TE = 0/2.
00 00
r Region of separation
o L-==t:=:===:I==::::::iC4._....1..-_-...1..._..-1L.....-_-i- _ __J
1.0 1.1 U/Ue .8 at V = 2 in.
.5 Aft shock location x/c, percent for Tl= 0.76 C PTE for11 = 0.76 -.1 and x/c = 0.96 -.2 5 6 7 11 12 13 920431 (d) Summary of breakpoints for pressure- and buffet-derived quantities.
Fig. 10 Concluded.
-0- aT = 50
-0-
aT = 50
-.4
~ aT = so
---6- aT = 10° -1.5 -.3 0 0
-1.0 0
-.2 CpTE 0 -.5 Cp -.1 .5 0 ..._ __ __. ___ ....... ___ ...._ __ __.
1.0 3 5 9 0 .2 .4 .6 .8 1.0 x/c 920432 920433 (a) Pressure profiles for several angles of attack, T/ = (b) Variation of upper surface pressure coefficients 0. 76. Solid symbols are lower surface Gp. with angle of attack at x / c = 0. 96 and T/ = 0. 76.
1.2 Break 1.0 In curve .8 CNA .6 .4 2 4 6 10 12 14 .2 0 4 8 12 16 a,deg ( c) Variation of airplane normal-force coefficient characteristics with angle of attack and buffet intensity.
Fig. 11 The angle-of-attack relationship between pressure coefficients and buffet characteristics for M = 0.85, q R:: 300 lb/ft , and lhE/TE = 5/10. Boundary-layer profiles not shown because flow was separated.
1.4 Region of separation ~ <ran , WT .7 g 1.2 CNA .6 Aft shock location x/c, percent
for11 = 0.76
- .10 C PrE
for 11 = 0.76
-.25
and x/c = 0.96
-.40 .9 Cn .6
for11 = 0.76
.3 .__ _ ___, __ __,_ __ _._ __ ...._ __ ...__ __ ..___ 3 5 6 7 8 9 10 11 12 13 aT, deg 920435 ( d) Summary of breakpoints for pressure- and buffet-derived quantities.
Fig. 11 Concluded.