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A Method for Localizing Wing Flow Separation at Stall to Alleviate Spin Entry Tendencies

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A Method for Localizing Wing Flow Separation at Stall to Alleviate Spin Entry Tendencies 78-1476 A Method for Localizing Wing Flow Separation \* at Stall to Alleviate Spin 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 AIRCRAFT 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 wings, 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 angle of attack, 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 airfoil 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 airfoils 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 aileron 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.
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