78-1476

A Method for Localizing Flow Separation \* at to Alleviate Entry Tendencies T. W. Feistel and S. B. Anderson, NASA Ames Research Center, Moffett Field, Ca.; and R. A. Kroeger, University of Michigan, Ann Arbor, Mich.

AlAA SYSTEMS AND TECHNOLOGY CONFERENCE L Los Angeles, CalifJAugust 21-23,1978

This is a US. Government work and is not copyrightable under U.S.C. 105.

. -- A \IETIIOD FOR 1I)CALIZING WISC fLW SEPARRTION .\T STALI TO ALI.FVIAT? SPIN ENTRY TENDENCIES

T. W. Fe15t~!l' and S. B. AndersonT Am. Research Center, NASA. MJffs>tt Field. Calrforn1.3 md R. &. Krocger' university Of Mlch1g.m. Ann li-bor. Mic:irgan

ABSTRACT Theoretical models of three-dimensional , using a nonlinear-lifting-line approach with a sim- A wing leading-edge modification has been ulated Stalled wing section, had suggested that developed, applicable at present to single-engine strong vorticity would be Shed at the edges of the light aircraft, which produces stabilizing vortices unattached section. A wind-tunnel model was fabri- at stall and beyond. These Vortices have the effect cated with partial span slats added along the of fixing the stall pattern of the wing such that entire leading-edge except for a ma11 length near the various portions of the wing upper surface stall the mid-semispan. These differences in leading- nearly symmetrically. The lift coefficient produced edge configuration were intended to produce a is essentially constant to very high angles of strong streamwise vorticity around the stalled sec- attack above the Stall angle of the unmodified wing. tion and thus, due to a decrease in the local It is hypothesized that these characteristics will induced , keep the other areas help prevent inadvertent spin entry after a stall. attached to high angles of attack. By varying the Results are presented from recent large-scale wind- spanwise position and width of the unslatted see- tunnel tests of a complete light aircraft, both tion, a post-stall lift curve shape Could be pro- with and without the modification. duced, which varied from practically flat on top to double-peaked, depending on the spanwise position of the gap in the leading-edge slats. INTRODUCTION This mid-semispan flow-control technique was Stalls and spins have continued to be a major first developed in experiments in the University of cause of fatal and nonfatal accidents involving Michigan 5- by 7-Foot Wind Tunnel and in the NASA- general aviation aircraft. As discussed in a his- Ames 7- by 10-Foot Wind Tunnel. torical review of stall/spin characteristics,' the aerodynamic factors that affect stall/spin behav- A second series of tests was performed in the ior have been studied for many years and are well NASA-Ames 7- by 10-FOot Wind Tunnel; the results known; however, the incorporation of the proper are reported in Ref. 2. In these Studies, using a L combination of these factors to provide stall/spin half-span model, the slats were replaced by avoidance in current general aviation aircraft has leading-edge gloves which added camber and a larger proved to be a difficult design challenge. radius to the leading edge, similar to a CAW-1 air- foil section. The results showed similar flow con- A key part of providing acceptable stall/spin trol capabilities but the effect was not quite as behavior involves the wing aerodynamics. Lateral dramatic on the post-stall lift curve as the slats. instabilities and the loss of lateral control, com- However, the gloves were capable of producing a mon to most aircraft when in a stall, are due to a flat-top lift curve without showing a perceptible rapid spread of flow separation on the outer portion drag penalty with respect to the clean wing. In of the wing. Many methods to control wing-flow addition, they were simple enough to constitute an separation have been examined, including aerodynamic acceptable type of add-on to a general aviation twist or geometric washout, wing slots or Slats, production aircraft. Subsequent wind-tunnel stud- change in section, variable thickness ratio, ies of a full-span wing in the NASA-Ames 7- by 10- and the USE of leading-edge Stall stripe. Although Foot Wind Tunnel showed that sideslip did not sig- some of these methods have been somewhat successful nificantly alter the effectiveness of this flow- in improving stall/spin behavior, the increased com- control concept. plexity of the wing design and loss of performance have acted as deterrents to widespread acceptance by AS a next Step, the decision was made to test the general aviation industry. the flow-Control method on a typical light airplane in the NASA-Ames 40- by 80-Foot Wind Tunnel, both Recently an improvement in post-stall aerody- with and without engine power and with various con- namic flow control has been made in a research pro- trol surface deflections. This paper presents and gram conducted jointly at Ames Research Center and discusses some results of these recent studies. at the University of Michigan. Basically, the con- The aircraft chosen was a Beechcraft Musketeer, cept involves the shedding of vortices at Stall at Model 23A. A photograph of the aircraft/wind- the mid-semispan leading edge, which Serves to pre- tunnel model mounted in the tunnel is shown in serve the lift, both inboard and outboard, to very Fig. 1. large angles of attack.

FULL-SCALE WIND-TUNNEL TESTS

The aircraft was modified by attaching a *Research Engineer. removable fiberglass leading-edge glove which was +Research Assistant for Interagency Programs. installed in Segments. The design of the glove was Associate Fellow AIAA. similar to that used in the earlier 7- by 10-foot Wead, Aircraft Research Laboratory. Member wind tunnel tests, Le., by matching the nose of a AIM. GAW-1 airfoil to the leading edge of the wing such

1 that the upper surfaces of the two approx- Starting with the bottom pair of photos, the angle " imately coincide over 20-30% of the chord (a sketch of attack is 12O. As expected for this pre-stall is Shown in the lower part of Fig. 2). This angle, the flow is about the same on both wings, results in a larger leading-edge radius as well as with a Small amount 0.f Separation occurring at the greater camber at the nose; the lower surface is trailing edge in the wing root region. The tuft faired flat so that it blends with the bottom of patterns at a = 16' and a = 20' (not shown) reveal 4 the wing at about 30% chord. This Simple modifica- little to distinguish between the two configura- tion is by no means optimum, but it has been shown tions. At a = 24'. in the next pair of photos to delay leading-edge separatjon to significantly shown, the favorable effect of the leading-edge higher angles of attack. modification is especially well illustrated, with the flow ahead of the breaking down on the The leading-edge glove segments were designed unmodified wing while it is still well attached on so they could be removed and rearranged to produce the modified version; it remains so through a = an unprotected gap, varying from 1/16 to 1/4 of the 28O. In the final set of pictures, at a = 36'. e. semispan in width, at various spanwise positions on the flow separation at the tip of the modified each of the wings. A Sketch of the layout and wing, which was partial at 32- (not shown), is com- nomenclature is shown in Fig. 2. Figure 3 is a plete (it is interesting to note that the tuft pat- close-up photograph of a typical modification. The tern here is similar to that for the unmodified location and width of the unprotected gap were var- wing at n = 24'). ied systematically during the exploratory part of the tests. These tests were run with the horizontal Lateral Characteristics tail removed, in order to focus on the wing charac- teristics, at an airspeed of about 77 mph (124 kph). The rolling-momnt data for these two config- urations with neutral Controls, and with Some max- imum control limits for full aileron deflec- Results with Modified Leading-Edge roll tion (represented by the open points at selected and for Basic Aircraft angles of attack), are shown in Fig. 6. AS would be expected from the tuft photos, the rolling The most desirable leading-edge modification moments for the modified wing (in Fig. 6.3) are configuration tested in this phase. based on the fairly well-behaved to an angle of attack of about shape of the lift curve and the rolling moments pro- 32'. above which they Start to depart. The excur- duced at stall, was approximately the same as that sion at 20°-219 (Point "A") is thought to be due to wed in the earlier 7- by 10-foot wind tunnel teats the leading-edge stall in the unprotected gap with a semispan wing,' i.e., a 1/8 semispan gap occurring on one wing first. Subsequent tests with located just inboard of the mid-semispan (position a sharper leading-edge radius in the gap have in Fig. 2). 4 reduced this excursion. For the unmodified wing (in Fig. 6b) the divergence in roll is much more Longitudinal Characteristics extreme, with large uncontrollable excursions v occurring at 22'-2G0; these excursions are due to The tail-off lift curve for this configuration asymmetric wing flow separation which was observed is shown in Fig. 4.3, along with the basic tail-off in the tufts. AS be the aileron effec- aircraft characteristics for comparison in Fig. 4b. can seen, tiveness was significantly higher for the modified (Note that these data are for the configuration wing at the higher angles of attack, with that of with tail off, power off, and. flaps up so that wing- the unmodified wing dropping to very low values at body effects only are being shown.) It can be Seen a = 28' to 36O. The yawing moments for both ver- that both the modified and unmodified configurations sions (not shown) were relatively small. A cursory have approximately the same CL,~ The shape Of . look at the contribution Of the modification to the top of the lift curve, however, is quite dif- dihedral effect Cgg and, to a lesser extent, ferent. The lift Of the modified configuration, directional stability Cn8 (for the configurations instead of steadily decreasing, remains essentially constant to an angle of attack of about 32'. It is with tail on), indicates that they are enhanced hypothesized that this characteristic implies somewhat at the higher angles of attack. The sensi- improved roll damping past stall: i.e., from the tivity of the post-stall lateral characteristics of tuft photos, the flow on the outer portion of the, the modified aircraft to small yaw angles is still wings stays attached, with separation occurring in being investigated. the vicinity of the mid-semispan and inboard. The tips then, Which the largest contributors to are Results with Two Other Types of roll, presumably have positive resulting in d CLa Leading-Edge Modifications an improved roll damping for the wing. The lift of the basic configuration, on the other hand, falls In order to investigate other means of obtain- off steadily after the maximum. .This negative ing the same results, two other leading-edge mod- slope and the observed tip-flow separation imply ification schemes were tested, both with a discon- negative roll damping as is known to occur in the tinuity at position 4 (Fig. 2)-that position found classic post-stall case. optimum for the gap in the leading-edge glove.

Flow Visualization The first of these consisted of the same leading-edge glove full-span in combination with a The tuft photos in Fig. 5 correspond to the large,1.5 in. (3.8 cm) wide, horizontally disposed lift curves shown in Fig. 4. They illustrate the leading-edge spoiler l/8 semispan long. The data flow structure over the wing for a range of angles for this variation, along with a sketch, are shown of attack from immediately pre-stall to deep post- in Figs. 7a and 7b. A somewhat larger drop in CL L stall. The photos of the unmodified version are on resulted after Chax, but recovery was good, with the left and those of the modified are on the right. a reasar.ably Elat top on the lift curve. The coll- COKCLUDING REMAlXS ins moments, as shwm, look bet.t.er than for the modif; cation prasrntfd rarlisr-prohably because of A wing loading-edge modification has been stronger vortices bcinq zhsd by the large leading- Eeveloped that chrngrs the stall pattern su that edge spoiler. The yzwing moment (not shown) was tke onset of separation is lucalized at. the semi- small. Further inveatigat.ions are being made to span leading edge. Vortices shed at t.his point are determi ne whether similar results can he obtai,ned thought ta help relieve the f-lcw on the inboard and with a snnller spoiler. oiltboavd portions of the wing. so that the fiow separatbn pattern is Stabilized and st.ayG fixed to The ncxt. variation in the leading-edge modi fi- large angles of attack. The resulting aerodyniuric cation .scbcmes investigctd resembled the conven- characteristics of the airplane are improved iI: tional st.r\ll pattern control treatment used on most of the important aerx?rt.r affecting spin deppar- current liqht: aircraft. It employed t.he basic ture. For example, while is about the sme %"ax wifig, with no leading-edge glove, wi.th but a 3/8 as for the ilrmodjfied wing, the shci;e of the tup of ir.. (0.95 cml squ-rr "stail strip" at the %?ne the lift curve for the modified wing is i;nprovcd so position 4 (Fig. 2;. This resulted i3 the lift t.hat it is essentially flat. 1.0 npprorirn+tely 3ZE shown Fig. curve in Sa; '4n.x is lower and occurs angle of ottnck. In adSi tion, flow visualization at a lower anrrle of &tack. However, the level of studies showed that the flow over the outLiOard por- CL, i.s maintained for a few dcgrees farther than tion of the wing stays attached to much hiher for the basic winq. but beyond u = 22' it angles, indicating that favorable effects on pst-

:a

1.8 I~

:'2 I

8' CL , t-

4

32mi 7

0

cy

Pig, 7 AirzmZt witl: full leading-edye glovc, tai.l-off - L;?miler at positiuii 4.

c 4

1.2

.8 CL

.4

316 (0.95 cml SQUARE STALL STRIP AT POSITION 4 ISPANWISE LENGTH, 118

0

-8 0 8 16 24 32 40 I? .deg

(a) Lift characteristics. (b) Rolling moments.

Fig. 8 Basic aircraft with stall strip at position 4, tail-off.

40 2.0 - T - 32

1.6 -

- 24

1.2 - .de9

CL - 16

.8 -

8

.4 -

0

-.05 -.M ;03 42 41 0 .01 n .des CP

la) Lift characteristics. (b) Rolling and pitching moments

Fig. 9 Complete aircraft with modified leading edge - landing approach condition.

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