Flight Investigation of the Effects of an Outboard Wing-Leading-Edgemodification on Stall/Spin Characteris Tics of a Low-Wing, Single-Engine, T-Tail Light Airplane

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NASA Tech ni ca I Paper 2691

July 1987 Flight Investigation of the Effects of an Outboard Wing-Leading-EdgeModification on Stall/Spin Characteris tics of a Low-Wing, Single-Engine, T-Tail Light Airplane

NASA NASA Technical Paper 2691

1987 Flight Investigation of the Effects of an Outboard Wing-Leading-EdgeModification on Stall/Spin Characteristics of a Low-Wing, Single-Engine, T-Tail Light Airplane

H. Paul Stough 111, Daniel J. DiCarlo, and James M. Patton, Jr. Langley Research Center Hampton, Virginia

National Aeronautics and Space Administration

Scientific and Technical Information Off ice Contents List of Figures ...... v Summary ...... 1 Introduction ...... 1 Symbols and Abbreviations ...... 1 Description of Airplane ...... 3 Instrumentation ...... 4 Evaluation Procedure ...... 4 Results and Discussion ...... 4 Stall Characteristics ...... 4 Idle-power stalls ...... 4 Power-on stalls ...... 5 Full-flaps stalls ...... 5 Spin Characteristics ...... 5 Analysis of Spin Resistance ...... 7 Effect of Wing Modification on Airplane Lift and Drag ...... 8 Summary of Results ...... 8 References ...... 9 I Tables ...... 10 Figures ...... 21 List of Figures

Figure 1 System of body axes. Arrows indicate positive direction of quantities. Figure 2 Test airplane with modified wing leading edge. Figure 3 Three-view drawing of test airplane. Dimensions are in feet. Figure 4 Wing-leading-edge modification. Figure 5 Airplane design center-of-gravity envelope showing loadings at test altitude. Figure G True angle of attack as a function of measured angle of attack at wing-tip boom location. Figure 7 Slow deceleration to idle-power, lg, wings-level stall with flaps and gear retracted. Test weight = 2438 lb; c.g. = 0.2777C. Figure 8 Visualization of flow over wings by means of tufts. Figure 9 Idle-power, lg, wings-level stall with maximum stabilator deflection, flaps and gear retracted. Test weight = 2438 lb; c.g. = 0.2777C. Figure 10 Idle-power stall of modified airplane from 30" banked left turn with flaps and gear retracted. Test weight = 2677 lb; c.g. = 0.2823C. Figure Idle-power stall of modified airplane from 45" banked skidding left turn \cith flaps and gear retracted. Test weight = 2677 Ib; c.g. = 0.2823C. Figure Maximum-power, lg, wings-level stall of modified airplane with flaps and gear retracted. Test weight = 2438 lb; c.g. = F.2777C. Figure Maximuni-power stall of modified airplane from 30" banked skidding left turn with flaps and gear retracted. Test weight = 2677 lb; c.g. = 0.2823C. Figiire 14 Maximuin-power stall with flaps extended 40" and gear retracted. Test weight = 2438 lb; c.g. = 0.2777C. Figure 15 Maximum-power stall of modified airplane from 30" banked skidding left turn with flaps extended 40" arid gear retracted. Test weight = 2438 lb; c.g. = 0.2777C. Figure 16 Coinparison of response of basic and modified airplane to pro-spin control inputs at idle power, ailerons neutral, flaps and gear retracted. Test weight = 2438 Ib; c.g. = 0.2777C. Figure 17 Comparison of response of basic and modified airplane to pro-spin control inputs at idle power, ailerons with the spin, flaps and gear retracted. Test weight = 2438 Ib; c.g. = 0.2777C. Figure 18 Comparison of response of basic and modified airplane to pro-spin control inputs at idle power, ailerons against the spin, flaps and gear retracted. Test weight = 2438 lb; c.g. = 0.27775. Figure 19 Comparison of response of basic and modified airplane to pro-spin control inputs at idle power, ailerons neutral, flaps deflected 40", gear retracted. Test weight = 2438 lb; c.g. = 0.2777C. Figure 20 Comparison of response of basic and modified airplane to pro-spin control inputs at maximum power, ailerons neutral, flaps and gear retracted. Test weight = 2438 lb; c.g. = 0.2777C.

V Figure 21 Right and left spins of modified airplane at maximum power entered from lg, wings-level flight, ailerons neutral, flaps and gear retracted. Test weight = 2690 Ib; c.g. = 0.32072. Figure 22 Right spin of modified airplane at maximum power entered from lg, wings-level flight, ailerons neutral, flaps and gear retracted. Test weight = 2829 lb; c.g. = 0.27872. Figure 23 Right spin of modified airplane at maximum power entered from 30' banked turn, ailerons neutral, flaps and gear retracted. Test weight = 2670 lb; c.g. = 0.28422. Figure 24 Left spin of modified airplane at maximum power entered from zoom maneuver, ailerons against the spin, flaps and gear retracted. Test weight = 2677 lb; c.g. = 0.28232. Figure 25 Left spin of modified airplane at maximum power entered from zoom maneuver, ailerons against the spin, flaps and gear retracted. Test weight = 2678 lb; c.g. = 0.25942. Figure 26 Left spin of modified airplane at maximum power entered from zoom maneuver, ailerons against the spin, flaps and gear retracted. Test weight = 2678 lb; c.g. = 0.23972. Figure 27 Distributions of times for first turn of rotation following pro-spin control input. Figure 28 Maximum attainable angle of attack and stall-margin concept. Figure 29 Variation of maximum trim angle of attack with center-of-gravity position for full trailing-edge-up stabilator deflection. Figure 30 Variation of maximum attainable outer wing angle of attack with center-of-gravity position. Figure 31 Airplane lift characteristics with and without wing-leading-edge modification. Figure 32 Airplane drag characteristics with and without wing-leading-edge modification.

vi Summary yawing motions which the pilot may find difficult to control. High angular rates may result from unsta- Flight tests were performed to investigate the ble damping in roll and may lead to autorotation at change in stall/spin characteristics due to the addi- higher angular rates and at angles of attack at which tion of an outboard wing-leading-edge modification the vehicle may exhibit a developed spin mode. to a four-place, low-wing, single-engine, T-tail, gen- Wing-leading-edge devices such as slots, slats, or eral aviation research airplane. Stalls and attempted flaps can significantly improve the damping-in-roll spins were performed for various weights, center- characteristics of wings near thc stall. Early research of-gravity positions, power settings, flap deflections, of the National Advisory Committee for Aeronautics and landing-gear positions. Both stall behavior and (NACA) by Weick, et al. (refs. 5 through 9) identified spin resistance were improved compared with the the potential aerodynamic benefit of such wing modi- baseline airplane. The baseline airplane would read- fications; however, many of the concepts proposed by ily spin for all combinations of power settings, flap the early research efforts proved impractical because deflections, and aileron inputs, but the modified air- of attendant degradation in aerodynamic perfor- plane did not spin at idle power or with flaps ex- mance, complexity of construction, excessive main- tended. With maximum power and flaps retracted, tenance, and cost. Also, certain types of leading- the modified airplane did enter spins with abused edge, high-lift devices improved stall behavior but loadings or for certain combinations of maneuver and aggravated spin characteristics and degraded spin control input. The modified airplane tended to spin recovery. at a higher angle of attack than the baseline airplane. Model tests at NASA Langley more fully explored the effects of these and several other wing-leading- edge modifications. Promising modifications were Introduction selected for flight testing based on the criteria that, for practical application to light aircraft, such a In response to the need for improving the stall/ modification should have a minimum adverse effect spin characteristics of general aviation airplanes, on airplane performance below the stall, should be the National Aeronautics and Space Administ~ration lightweight, should be passive (not require sensors (NASA) is conducting a comprehensive program to and actuators or pilot activation), should require a develop new stall/spin technology for this class of minimum of maintenance, and should be low cost. airplane (ref. 1). The program incorporates spin- tunnel model tests, static- and rotary-balance wind- A discontinuous, outboard wing-leading-edge tunnel tests, analytic studies, and airplane and model droop modification which appears to meet these re- flight tests for a number of configurations represen- quirements was conceived and flight tested on an air- tative of typical general aviation airplanes weighing plane with an untwisted rectangular wing (ref. 10). under 4000 lb. Subsequently, a similar wing modification was devel- Stalling and spinning are major causal factors in oped (ref. 11) for the twisted, tapered wing of the general aviation accidents (refs. 2 through 4). Exam- low-wing, T-tail airplane of reference 12. ination of the circumstances involved suggests that This report presents results of flight tests of the the majority of these stall/spin accidents occur at low low-wing, T-tail airplane with the wing leading edge altitude and involvc inadvertent loss of longitudinal modified to increase spin resistance. Stall charac- or lateral-directional control, spin entry, and ground teristics and the results of attempted spin entries impact before the spin becomes fully developed. In are presented along with the effects of center-of- recognition of the potential danger associated with gravity position, power, flaps, and control inputs on the inadvertent stalls at low altitude, studies are be- these characteristics. Comparisons are made with ing conducted at the NASA Langley Research Center the stall/spin characteristics of the airplane prior to to define concepts which improve the stall character- modification of the wing. istics and spin resistance of light airplanes. Because the inadvertent spin usually occurs at low altitude, Symbols and Abbreviations emphasis is being placed on spin resistance rather than spin recovery. Measurements are referred to the set of body axes Lateral stability and controllability at the stall of with the origin at the airplane center of gravity, as airplanes with unswept wings are characterized by shown in figure 1. Measurement,s were made and the tendency of such wings to experience unstable quantities are presented in U.S. Customary Units. damping in roll and autorotation near the stall. Un- Symbols in parentheses refer to labels on computer- stable damping in roll can result in rapid rolling and generated figures. P (ROLL RATE) roll rate (positive for (NORM ACC) nornial acceleration at rolling right wing down), center of gravity (positive deg/sec in negative zb direction), I g units (PROP SPEED) propeller and engine speed, rpm I

(ALTITUDE) pressure altitude, ft I 9 (PITCH RATE) pitch rate (positive BL butt line; lateral displace- for pitching nose up), , ment from centerline of deg/sec fuselage, in. R spin radius, ft; R x b wing span, ft 9 , ft (&I2 tan ff CD drag coefficient, zpv2s I RIS rate of sink, ft/sec I CL lift coefficient, Lift r (YAW RATE) yaw rate (positive for ipv2s yawing nose right), - C mean aerodynamic chord, deg/sec ft (RES ACC) resultant linear acceler- c.g. center of gravity ation, (LONG ACC2 + LAT ACC2 NORM FS fuselage station; longi- + ACC~)'/~,g units tudinal coordinate measured along waterline, (RUD FORCE) rudder pedal force positive moving aft, in. (positive for forces tending to push right acceleration due to 9 pedal forward), lb gravity, 32.2 ft/sec2 S wing area, ft2 moment of inertia about xb,Yb, and zb body axis, T period of spin, x F,sec respectively, slug-ft2 (THRUST) propeller thrust deter- L.E. leading edge mined using engine performance chart and as- (LAT ACC) lateral acceleration at sumed propeller efficiency center of gravity (positive of 0.85, lb in positive Yb direction), g units V (SPEED) true airspeed, ft/sec (see fig. 1) (LAT FORCE) lateral wheel force (positive for forces tending to WL waterline; vertical coordi- rotate wheel clockwise), nate in airplane plane of lb symmetry measured per- pendicular to reference (LONG ACC) longitudinal acceleration line (reference line is WL at center of gravity 40.00 and passes through (positive in positive xb propeller shaft centerline direction), g units at back of spinner), in. (LONG FORCE) longitudinal wheel force WS wing station, in. I (positive for forces tending to pull wheel body axes through air- aft), lb plane c.g. (see fig. 1) m mass of airplane, slugs 2 distance rearward from leading edge of mean (MPR) engine manifold pressure, aerodynamic chord to in. Hg center of gravity, ft distance between center R (TURN RATE) total angular velocity of of gravity and fuselage airplane, (p2 + 42 + r2)1j2, centerline (positive when with sign of r, deg/sec center of gravity is right of centerline), ft Description of Airplane distance between center of gravity and fuselage The test airplane was a four-place, low-wing, reference line (positive single-engine, retractable-gear, T-tail design. This when center of gravity is airplane was a one-of-a-kind research airplane, but below line), ft it was considered representative of this class of aircraft. A photograph and three-view drawing of the (ALPHA CG) true angle of attack at airplane are presented as figures 2 and 3, respectively; airplane center of gravity, physical characteristics of the baseline airplane are deg (see fig. 1) presented in table 1. The mass characteristics and angle of attack at left- inertia parameters for representative loadings tested wing tip boom location, with the modified wing are presented in table 2. The baseline wing incorporated a modified NACA deg 652-415 airfoil section. The leading edge of the ta- measured angle of attack pered portion of the wing was drooped and transi- at instrumentation boom, tioned smoothly from no droop at the start of the ta- deg per to maximum droop at the wing tip. The wing had angle of attack at right- slotted trailing-edge flaps and plain ailerons. The wing tip boom location, horizontal tail was mounted on the vertical tail at the top of the rudder and consisted of a stabilator with deg a geared antiservo tab to provide longitudinal trim. true angle of attack at An adjustable spring in the rudder control cables instrumentation boom, served to trim out rudder forces. Aiieron trim was deg not adjustable in flight. The airplane was equipped (BETA CG) angle of sideslip at with a spin-recovery parachute system (ref. 13), a airplane center of gravity, quick-release door on the right side, and a pyrotech- deg (see fig. 1) nic egress panel in the absence of a door on the left side (ref. 14). (AILERON) average aileron deflection The baseline wing was modified by installation of (positive for right aileron a glove over the forward part of the airfoil to pro- trailing edge down), 1/2 vide a 2.9-percent chord extension and droop, which (Right aileron deflection increased the leading-edge camber and radius, as + Left aileron deflection), shown in figure 4. The droop geometry duplicated deg (see fig. 1) the leading-edge airfoil configuration developed in (STARILATOR) stabilator deflection reference 11 as a means of improving lateral stability (positive for trailing edge at the stall. Transition from the original airfoil sec- down), deg (see fig. 1) tion to the drooped leading edge was an abrupt dis- continuity. Coordinates of the baseline wing sections flap deflection (positive are presented in table 3; coordinates of the new air- for trailing edge down), foil sections created by addition of the glove are pre- deg sented in table 4. The modified wing-leading-edge ge- (RUDDER) rudder deflection (posi- ometry varied linearly in the spanwise direction from tive for trailing edge left), the inboard end to the tip end of the modification. deg (see fig. 1) Tests were conducted at center-of-gravity positions from 0.2394C to 0.32072 (12.66 to 4.53 percent airplane relative density, static margin) at test weights of 2438 to 2829 lb (take off weights of 2498 to 2889 lb). Figure 5 shows the 2% test loadings relative to the design c.g. envelope. The air density, slugs/ft3 inertia yawing-moment parameter, calculated from bank angle (positive for measured moments of inertia, varied from -62 x right wing down), deg to -24 x

3 Instrumentation a right spin), or against the spin (left wheel for a right spin). In some instances, additional control in- The airplane was instrumented to measure and puts were made during the post-stall gyrations in an record flow angles and true airspeed ahead of each effort to drive the airplane into a spin. Spin entry was wing tip, linear accelerations along the body axes, also attempted from the conditions of the stall/spin angular rates about the body axes, control surface scenarios described in reference 4, as representative positions, control wheel and rudder pedal forces, en- of typical inadvertent spin entries. gine power parameters, altitude, and spin-recovery The effects of center-of-gravity position, weight, parachute load. The onboard data system was power level, and flap position were investigated in- supplemented by ground-based telephoto video and dividually and in combination. Stall/spin charac- movie canieras and by cockpit- and wing-tip-mounted teristics were evaluated by NASA pilots over the cameras. Pilot comments were recorded on the weight and c.g. range. Test pilots from three light- ground-based videotape. All data were time corre- airplane manufacturers and from the FAA evaluated lated and provided a continuous time history from the airplane stall/spin characteristics at the aft c.g. maneuver entry through recovery. The data were (0.2804C) and at a gross weight of 2690 lb. telenietered to a ground station and were monitored in real time along with a video display of the airplane. For debriefing purposes, the videotape and telemetry Results and Discussion records were reviewed shortly after each flight. Stall Characteristics Linear accelerations and flow measurements were corrected to indicate conditions at the airplane cen- Airplane stall characteristics with the wing- ter of gravity using the techniques reported in ref- leading-edge modification are presented in table 5 erence 15. Angle-of-attack measurements were cor- and are illustrated via time histories in figures 7 and 9 rected for upwash by applying the flow correction through 15. Comparisons are made with the un- presented in figure 6 (taken from ref. 12). Because modified airplane where data are available. Qual- true angles of attack from both wing tips were av- itatively, the baseline airplane stall characteristics eraged, corrections to angle of attack due to side- had been described as “good” and more docile than wash at the wing tips have a tendency to cancel out. those of most current airplanes of this class (ref. 12). Likewise, because measurenients from both wing tips The pilots described the modified airplane stall were averaged, corrections to sideslip due to up- characteristics at the aft-center-of-gravity (0.2804C) wash at the wing tips have a tendency to cancel out gross-weight loading as “outstanding”, “very docile” , (ref. 16). “benign”, and “controllable” . Good controllability throughout the stall and excellent aileron control at Evaluation Procedure minimum speed conditions with the control wheel full back were also noted. The results of the investigation were based on pi- Stall characteristics of the modified airplane were lot comments, time-history records of airplane mo- evaluated with flaps and gear retracted, with flaps tions and controls, and films and videotapes of the extended 10’ and 40°, and with flaps and gear ex- tests. The majority of maneuvers were flown by three tended. Stalls were entered from level flight, from NASA research pilots. Maneuvers were also flown by left and right turns, and from pullups with and three industry test pilots and two FAA test pilots. without sideslip at idle power, power for level flight Initial tests evaluated airplane stall and depar- (75 percent), and maximum power. Both the slow ture characteristics. Maneuvers included lg and ac- (1 knot/sec) and rapid (3 to 5 knots/sec) decelera- celerated (banked) stalls with various combinations tion rates to the stall, as specified in reference 17, of engine power, bank angle, and sideslip as shown were used. For the modified airplane, no uncontrol- in table 5. lable motions were encountered during the 127 stalls Spin tests were performed at the NASA Wallops performed over the loading range tested (0.2394 Flight Facility. Test altitudes ranged from 12 000 to 0.32074. Airplane stalling characteristics met the to 6000 ft. Most spins were entered at an altitude requirements of FAR Part 23 (ref. 17). of ahoiit 10000 ft, which corresponds to a Reynolds number at the stall of about 3 x lo6. Spin-entry con- Idle-power stalls. For the baseline airplane, slow ditions included conibinations of acceleration, roll, deceleration to a lg, wings-level stall with near- pitch, yaw, and power. Pro-spin controls were ap- zero sideslip resulted in a Dutch-roll-type motion plied at or just before the stall. The control positions with the airplane flying in and out of the stall. at entry were wheel back with pro-spin rudder and The baseline airplane stalled at about 20’ angle ailerons either neutral, with the spin (right wheel for of attack (2’ trailing-edge-up stabilator deflection

4 at a center-of-gravity position of about 28 percent bank could be easily controlled with ailerons and 2). Additional trailing-edge-up stabilator deflection rudder (fig. 12). beyond that required to stall the airplane generally For the modified airplane, stalls from coordinated resulted in a roll-off to the left. This roll-off was 30" and 60" banked left and right turns produced often initially uncontrollable even with full rudder an easily controllable right-roll tendency. When and aileron inputs against the roll. Stalls from stalled with sideslip, the modified airplane exhibited coordinated 30" and 60" banked turns produced little easily controllable roll tendencies away from the slip. roll-off tendency. When stalled with sideslip, the Figure 13 illustrates the controllable roll tendency of airplane rolled away from the slip; that is, right the modified airplane when stalled in a 30" banked sideslip produced a left roll. skidding left turn. For the modified airplane, slow deceleration to a With power for level flight, straight-ahead stalls lg, wings-level stall with near-zero sideslip (fig. 7) of the modified airplane were docile. Stalls from co- resulted in a small pitching oscillation (nose bobble) ordinated 30" banked left and right turns resulted in with roll- and yaw-rate oscillations that the pilot a tendency for the modified airplane to roll opposite perceived to be less throughout the maneuver than the direction of the turn. Stalls from slipping or skid- for the unmodified airplane. The modified airplane ding 30" banked turns resulted in a tendency for the stall was sensed at 20" angle of attack (about 4" modified airplane to roll away from the slip. trailing-edge-up stabilator deflection at a center-of- Full-flaps stalls. With flaps deflected 40°, the gravity position of about 28 percent C) and was baseline airplane tended to roll to the right at the usually accompanied by the onset of small pitch-rate stall for both idle- and maximum-power cases. For oscillations; however, the modified wing outer panels did not stall until the local angle of attack reached maximum-power cases, the rolling tendency could be countered by large rudder and aileron inputs. 36" as indicated by tufts on the wing. Figure 8 shows For the modified airplane, slow deceleration to a tufts on the modified wings during an attempted spin entry as photographed from a camera inside the lg stall at idle power with flaps deflected 40" did not airplane cabin. At the instant of this picture, the provide any stall warning; only a mild departure tendency resulted, and this tendency was easily control- outer panel of the ieft wing was just at stall (36" lable with rudder and aileron inputs. Stalls from co- angle of attack); the outer panel of the right wing ordinated 30' banked left and right turns were docile. was at 31" angle of attack. Additional trailing-edge- With the modified wing leading edge, slow decel- up stabilator deflection beyond that required to stall eration to a stall at maximum power with flaps the modified airplane generally resulted in a slight lg deflected 40" produced a very light stall break with roll tendency that was easily controlled with small a little more roll tendency than at idle power, but aileron and rudder inputs. Figure 9 compares the the airplane was characteristics of the modified airplane with the roll- easily controllable with coordinated rudder and aileron inputs (fig. 14). Stalls from coor- off tendericy of the unmodified airplane (ref. 12) for full trailing-edge-up stabilator deflection. dinated 30° banked left and right turns produced easily controllable right-roll tendencies. When stalled Stalls from coordinated 30" banked turns were in skidding turns, the modified airplane exhibited very docile with the modified wing (fig. 10). When easily controllable roll tendencies away from the slip stalled with sideslip, the modified airplane exhibited (fig. 15). little or no roll-off tendency (fig. 11). Roll tendency, For the modified airplane, slow deceleration to a if any, was away from the slip; that is, left sideslip lg stall at the power required for level flight with produced a right-roll tendency. flaps deflected 40" produced a docile stall with good aileron control. When stalled with sideslip, the Power-on stalls. When stalled with maximum modified airplane tended to roll away from the slip. power and near-zero sideslip, the baseline airplane usually rolled to the right. Stalls from 30" and Spin Characteristics 60" banked left and right turns resulted in roll- offs to the right. When stalled with sideslip, the A total of 244 spins were attempted during 25 baseline airplane rolled away from the slip. These evaluation flights with the modified wing. Of these, rolls often could not be controlled with rudder and 13 (5 percent) resulted in spins and 231 (95 percent) aileron until the airplane was unstalled by reducing resulted in spirals or failed to produce a spin or a spi- the up-stabilator deflection. ral. The modified airplane had a markedly increased When stalled with maximum power and near-zero resistance to spin entry compared with the baseline sideslip, the modified airplane was very docile. Any airplane, which, as reported in reference 12, entered rolling tendency was controllable with ailerons, and a spin in 173 (83 percent) of 209 spin attempts.

5 Figures 16 through 20 illustrate the predominant characteristics, and table 8 summarizes the modified- airplane response to pro-spin control inputs by com- airplane spin tendencies. paring motions with and without the wing modifica- With the modified wing, spin entry from lg tion. Thc baseline airplane would readily spin from flight was achieved only for loadings outside the lg flight, power-on or power-off, with the application design center-of-gravity envelope (beyond aft-center- of normal pro-spin controls. It spun at about 43' of-gravity position or over gross weight) and in- angle of attack. Application of recovery controls did volved full trailing-edge-up stabilator deflection, full- not always stop the spin. With the modified wing rudder deflection, and ailerons neutralized at the stall leading edge, no spins were obtained for idle-power (figs. 21 and 22). For these loadings, no special con- settings or with flaps deflected; spins were obtained trol inputs were required to achieve spin entry. After only with maximum power and flaps retracted. one turn or more of sustained pro-spin controls, the When intentional spins were attempted at ei- airplane would quickly pitch up into a spin. In one in- ther idle power or maximum power by applying aft- stance, a left spin at maximum power with the center stick, full-rudder deflection, and with ailerons neutral of gravity at 0.3207C1recovery control inputs were un- (fig. 16), with (fig. 17), or against (fig. 18) the rudder able to stop the spin, and the spin-recovery parachute input, the modified airplane typically entered a steep, was deployed (fig. 21). Because the baseline airplane controlled spiral with a slow turn rate and a high sink occasionally failed to respond to initial spin-recovery rate. Angle of attack was above that at which the control inputs, and because the baseline airplane was basic wing stalled but below that at which the outer not tested with a center-of-gravity position as far wing panel stalled. Tufts showed that the outboard aft as 0.3207?, it is not known whether the baseline part of the wing remained unstalled. Loss of lift on airplane would have exhibited an unrecoverable spin the inboard wing panels and airframe buffeting from under these conditions. the separated flow on the inboard wing panels pro- For loadings of the modified airplane near the vided the pilot with cues similar to a conventional design aft-center-of-gravity position (0.2804?), spins stall, although the wing outer panels continued to could be entered at maximum power from a 30' operate below stall and maintain roll damping. Ro- banked right turn by applying full trailing-edge-up tation stopped if the controls were not held in the stabilator, full right rudder, and neutral ailerons pro-spin position. (fig. 23). Typically, as the spiral progressed, the stabila- Spins were obtained for loadings of the modified tor rotated off the trailing-edge-up stop, even though airplane within the design loading envelope by us- the pilot exerted increased stick force. The reduced ing a zoom maneuver. With maximum power, the control-surface deflection is attributed to increased modified airplane was put into a dive to build up stabilator hinge moment with increased dynamic speed. It was then pulled up into a steep climb to pressure and to control-system flexibility. With full bleed off speed and achieve an extreme nose-high pro-spin controls, the typical motion of the modified pitch attitude at the stall. Pro-spin controls were airplane is somewhat spiral-like and somewhat spin- applied at the stall with aileron deflection against like. The root section of the wing is stalled, which the spin. Only left spins were obtained from zoom gives buffet and drag as in a spin. The tip panels maneuvers (figs. 24 through 26). are attached and produce roll damping and lateral At the forward center-of-gravity position tested control niuch like in a spiral. The modified airplane (0.2395?), spins were attempted with the modified reaches a steady terminal velocity at a speed much airplane loaded with 30 gal more fuel in the right higher than in a spin. The terminal velocity of the wing tank than in the left wing tank (200 Ib right spiral-like motion is dependent upon the magnitude wing heavy). Airplane spin tendencies were not of the trailing-edge-up stabilator deflection, which degraded by this severe loading asymmetry as shown determines how high an angle of attack is achieved in table 8. in the maneuver. For the few maneuvers that resulted in spin en- For the modified airplane, attempted spins with tries, the modified airplane spun at a higher angle flaps deflected typically produced a bucking spiral as of attack than the baseline airplane (table 7). Spin shown in figure 19, or, with maximum power applied, angle of attack increased as the center of gravity was a spiral as showii in figure 20. With flaps up or flaps moved aft and as weight was increased. Spin-entry down, relaxing the controls returned the modified tendency also increased as the center of gravity was airplane to normal flight. moved aft and as weight was increased (table 8). Table 6 presents the configurations and entry ma- The motions of the modified airplane following neuvers of the modified airplane for the 13 spins that application of pro-spin control inputs were compared were achieved. Table 7 presents representative spin with those of the baseline airplane in terms of the

6 time required for the airplane to rotate 360" in attack was used because it is the closest to the ini- heading or in roll (fig. 27). All maneuvers for which tial conditions of the test maneuver. This would best the modified airplane completed one turn of rotation, enable the influence of such variables as static mar- whether in a spin or spiral, were included. In general, gin, power, control inputs, and control phasing to be the modified airplane provided the pilot more time to readily detected. The extreme angles of attack that make corrective control inputs before completing one can be achieved by means of zoom maneuvers are turn of rotation. also shown in figure 30, but such maneuvers are not cvrisidered typical of inadvertent spin entries. Analysis of Spin Resistance The outer wing stall angle of attack was deter- Initial analysis of the spin resistance of the air- mined from measured angle-of-attack data and from plane with the outboard-wing modification was based corresponding films of tufts on the wings. These on the assumptions that (1) an airplane would be values for the basic and modified-airplane configu- highly spin resistant if roll damping could be main- rations are indicated by the horizontal lines in fig- tained for all maneuvers, and (2) roll damping could ure 30. Flight-test results indicated that the base- be maintained if the outer wing panel did not stall. line wing outer panel stalled at 20" angle of attack. Conceptually, spin resistance was thus related to the The intersections of the outer wing angle-of-attack difference between the maximum attainable angle of curves and the stall angle-of-attack curves suggest attack and the stall angle of attack. The maximum center-of-gravity positions forward of which the air- attainable angle of attack was defined as the maxi- plane should not spin and aft of which it would have mum local angle of attack of the wing outer panel, the potential to spin. For example, the airplane including dynamic effects; the stall angle of attack with the basic wing would be expected to spin fr was defined as the local angle of attack at which the all center-of-gravity positions shown, regardless outer-panel flow separated. For analysis purposes, power, since the maximum angles of attack exceed maximum attainable angle of attack was divided into the stall value of the basic wing. Indeed, the air- a baseline trim angle of attack and increments due plane with the basic wing did readily enter spins at to center-of-gravity changes, power effects, and dy- about 28 percent C (the only center-of-gravity posi- namic effects (fig. 28). If the summatiori of these tion tested for the baseline airplane) for both idle increments did not exceed the outer-panel stall an- and maximum power. gle of attack, the airplane was not expected to spin. An airplane whose maximum attainable angle of at- The modified wing outer panel stalled at 36O angle tack greatly exceeded the stall angle of attack was of attack. As shown in figure 30, the airplane with expected to enter a spin more readily than an air- the modified wing would not be expected to spin at plane whose maximum attainable angle of attack just idle power until the center-of-gravity position was aft slightly exceeded the stall angle of attack. During of about 28 percent C. rriariy rnaneuvers, the stall angle of attack was ex- The modified airplane was tested at center-of- ceeded temporarily without leading to a spin entry. gravity positions of 0.238C to 0.32?, and it did not The length of time that the stall angle of attack is spin with idle power. Figure 30 also indicates exceeded, as well as the magnitude of the increment that the airplane would be expected to spin with in angle of attack beyond that required for stall, may the use of maximum power, but spin tendencies be a factor in spin entry. would be small until the center of gravity was aft The static trim angles of attack with full nose- of about 28 percent 1.. With maximum power, spin up pitch-control input were measured for a range of entries were obtained for all loadings tested, includ- center-of-gravity positions. These were measured at ing 24 percent C; however, it became increasingly dif- maximum gross weight with maximum power, 50 per- ficult to enter a spin as the center of gravity was cent power, and idle power for the modified airplane moved forward, and very deliberate, prolonged in- with flaps retracted and deflected 40". Results are puts of a zoom maneuver were required to utilize presented in figure 29 as plots of static trim angle of the dynamic effects to cause spin entry for center-of- attack versus center-of-gravity position. The maxi- gravity positions forward of about 28 percent E. mum angle of attack presented in figure 30 is the peak The stall-margin concept of figure 28 and its value of the first transient in local angle of attack at quantification, such as shown in figure 30, may help the outer wing panel after pro-spin control inputs. an airplane designer assess trade-offs and their rela- The region between the curves of the maximum an- tive impact on spin resistance. The consequences of gle of attack and the trim angle of attack represents increasing control deflection, extending the center- the transient effects of the dynamics associated with of-gravity envelope, or increasing the installed power the spin-entry maneuver. The first peak in angle of are easy to visualize. For example, the change in stall

7 margin due to an increase in power could be offset single-engine, T-tail, general aviation research air- somewhat by moving the aft-center-of-gravity limit plane. Tests were conducted at weights of 2438 forward. Likewise, changes which reduce the stall to 2829 Ib (89 to 103 percent of maximum gross margin tend to reduce spin resistance. If a point weight) with center-of-gravity positions of 0.2394 is reached where the outer wing-panel stall angle of to 0.3207 mean aerodynamic chord (12.66 to 4.53 per- attack can be exceeded with subsequent loss of roll cent static margin). The following results were damping, the airplane could spin. With the modified indicated: wing, this airplane would be expected to spin at a higher angle of attack and may be more difficult to 1. Addition of the modification increased stall stop than the airplane with the baseline wing. angle of attack of the outer wing to 36", which was approximately twice the stall angle of attack of the Effect of Wing Modification on Airplane Lift basic wing (20'). and Drag 2. The modified airplane had improved stall behavior and increased spin resistance compared with Airplane lift and drag were measured for both the baseline and modified airplane configurations by con- the baseline airplane. ducting level flight runs at constant airspeeds from 3. Spin resistance has been related to the vari- maximum level flight speed to near stall onset. As ation in attainable outer wing-panel angle of at- shown in figure 31, there was no discernible difference tack with center-of-gravity position, power level, and in lift up to about 10' angle of attack. As shown in dynamic effects. figure 32, within the measurement accuracy, no dif- 4. For the few maneuvers that did produce spins ference was found in airplane drag for lift coefficients with the wing modification, the modified airplane typical of cruising flight. For lift coefficients between tended to spin at a higher angle of attack than the about 0.5 to 1.0, the modified wing had increased baseline airplane. drag. 5. Airplane cruise drag was unchanged by addition of the wing modification. Summary of Results

Flight tests were conducted to evaluate the change NASA Langley Research Center in spin resistance caused by the addition of an out- Hampton, Virginia 23665-5225 board wing-leading-edge modification to a low-wing, April 2, 1987

8 References 9. Weick, Fred E.; and Abramson, H. Norman: Investiga- tion of Lateral Control Near the Stall. Flight Tests With 1. Chambers, Joseph R.: Overview of Stall/Spin Technol- High- Wing and Low- Wing Monoplanes of Various Con- ogy. AIAA-80-1580, Aug. 1980. figurations. NACA TN 3676, 1956. 10. Staff of Langley Research Center: Ezploratory Study of I 2. Silver, Brent W.: Statistical Analysis of General Avia- the Effects of Wing-Leading-Edge Modifications on the tion Stall Spin Accidents. [Preprint] 760480, SOC.Auto- Stall/Spin Behavior of a Light General Aviation Airplane. motive Engineers, Apr. 1976. NASA TP-1589, 1979. 3. Ellis, David R.: A Study of Lightplane Stall Avoidance 11. White, E. Richard: Wind-Tunnel Investigation of Efects and Suppression. FAA-RD-77-25, Feb. 1977. of Wing-Leading-Edge Modifications on the High Angle- 4. Hoffman, William C.; and Hollister, Walter M.: General of-Attack Characteristics of a T-Tail Low- Wing General- Aviation Pilot Stall Awareness Raining Study. Rep. No. Aviation Aircraft. NASA CR-3636, 1982. FAA-RD-77-26, Sept. 1976. (Available from DTIC as 12. Stough, H. Paul, 111; DiCarlo, Daniel J.; and Patton, AD A041 310/4.) James M., Jr.: Flight Investigation of Stall, Spin, and 5. Weick, Fred E.; and Wenzinger, Carl J.: Eflect of Length Recovery Characteristics of a Low- Wing, Single-Engine, of Handley Page Tip Slots on the Lateral-Stability Factor, T-Tail Light Airplane. NASA TP-2427, 1985. Damping in Roll. NACA TN 423, 1932. 13. Bradshaw, Charles F.: A Spin-Recovery Parachute System for Light General-Aviation Airplanes. 14th 6. Weick, Fred E.; and Wenzinger, Carl J.: Wind-Tunnel Aerospace Mechanisms Symposium, NASA CP-2127, 1980, Research Comparing Lateral Control Devices, Particu- pp. 195-209. larly at High Angles of Attack. VII-Handley Page Tip 14. Bement, Laurence J.: Emergency In-Flight Egress Open- and Full-Span Slots With Ailerons and Spoilers. NACA ing for General Aviation Aircraft. NASA TM-80235, TN 443, 1933. 1980. 7. Weick, Fred E.; and Wenzinger, Carl J.: Wind-Tunnel 15. Sliwa, Steven M.: Some Flight Data Extraction Tech- Research Comparing Lateral Control Devices, Particularly niques Used on a General Aviation Spin Research Air- at High Angles of Attack. I-Ordinary Ailerons on Rect- craft. AIAA Paper 79-1802, Aug. 1979. angular Wings. NACA Rep. 419, 1932. 16. Mod, Thomas M.: Determination of Corrections to Flow 8. Weick, Fred E.; Sevelson, Maurice S.; McClure, Direction Sensor Measurements Over an Angle-of-Attack James G.; aid F!anagan, Marion E.: Ifzwstlgation Range From 0" to 8s". NASA TM-86402, 1985. of Lateral Control Near the Stall-Flight Investigation 17. Airworthiness Standards: Normal, Utility, and Acro- With a Light High- Wing Monoplane Tested With Vari- batic Category Airplanes. Federal Aviation Regula- ous Amounts of Washout and Various Lengths of Leading- tions, vol. 111, pt. 23, Federal Aviation Administration, Edge Slot. NACA TN 2948, 1953. June 1974.

9 Table 1. Physical Characteristics of Baseline Airplane Overall dimensions: Span, ft ...... 35.43 Length, ft ...... 27.80 Height, ft ...... 8.26 Powerplant: Type ...... Reciprocating, four cylinder horizontally opposed Rated power at sea level, hp ...... 200 Rated continuous speed, rpm ...... 2700 Propeller: Type ...... Two blades, constant speed Diameter,in...... 76 Pitch (variable), deg ...... 14 to 31 Wing: Area, ft2 ...... 173.7 Root chord, in...... 74.0 Chord of constant section, in...... 63.0 Tip chord, in...... 42.2 Mean aerodynamic chord, in...... 62.16 Aspect ratio ...... 7.24 Dihedral, deg ...... , 7.0 Incidence at root, deg ...... 2.0 Incidence at tip, deg ...... -1.0 Airfoil section ...... NACA 652-415- modified Flap: Chord, in...... 12.26 Span, in...... 85.50 Area (each), ft2 ...... 7.3 Hinge line, percent flap chord ...... 20.0 Deflection, deg ...... 40 down Aileron: Mean chord, in...... 10.01 Span, in...... 100.05 Area (each), ft2 ...... 6.95 Deflection, deg ...... 30 up, 16 down Stabilator: Area (including tab), ft2 ...... 25.0 Chord (constant), in...... 30.0 Aspect ratio ...... 4.0 Dihedral, deg ...... 0 Hinge line, percent stabilator chord ...... 26.97 Deflection, deg ...... lOup,lOdown Airfoil section ...... NACA 0012 Table 1. Concluded Tab: Chord (constant). in ...... 6.0 Span. in ...... 106.2 Area. ft2 ...... 4.4 Tab hinge line to stabilator hinge line. in...... 15.91 Vertical tail: Area (including rudder) . ft2 ...... 17.6 Root chord. in ...... : ...... 54.52 Tip chord. in ...... 28.62 Mean aerodynamic chord. in ...... 42.91 Span. in ...... 60.84 Aspect ratio ...... 1.47 Leading-edge sweep back. deg ...... 33.91 Rudder: Area. ft2 ...... 5.8 Average chord aft of hinge line. in ...... 11.85 Span (parallel to hinge line). in ...... 61.6 Deflection. deg ...... 28 left. 28 right Test weight at altitude. lb ...... 2420 Relative density p at 9000 ft altitude ...... 6.7 Representative center-of-gravity position: FS. in ...... 95.85 (0.2782C) BL. in ...... -0.56 WL. in ...... 38.15 Moments of inertia about body axes (based on represenative center-of-gravity position): IL.slug-ft2 ...... 1789 Iy.slug-ft2 ...... 2486 I,. slug-ft2 ...... 3796

Inertia yawing-moment parameter. L-Iy ...... -74 x 10-~

11 600

I I

12 13 Table 4. Coordinates of Modified Leadirig-Edge Contour in Percent Chord of Basic-Wing Airfoil Sections

WS 114.07 WS 205.9 Stat ion Ordinate St ation Ordinate 2.441 2.816 3.424 3.093 2.181 2.726 3.080 2.872 1.953 2.666 2.737 2.666 1.709 2.549 2.392 2.449 1.465 2.432 2.049 2.168 1.221 2.295 1.706 1.888 ,977 2.148 1.362 1.649 ,732 2.021 1.018 1.362 ,488 1.855 ,675 1.115 .244 1.689 ,332 .793 ,000 1.543 -.011 ,346 -.244 1.348 - 353 ,065 -.488 1.172 - .695 -.285 -.732 ,977 - 1.036 -. 746 - ,977 ,771 -1.378 -1.192 -1.221 .498 - 1.719 -1.681 - 1.465 .273 -2.059 -2.252 - 1.709 .ooo -2.397 -3.004 - 1.953 -.285 -2.732 -4.116 -2.197 - ,674 -2.859 -5.225 -2.441 -1.172 -2.71 1 -6.055 -2.686 -1.758 -2.358 -6.743 -2.832 -2.441 -2.007 -7.183 -2.686 -3.076 - 1.657 -7.470 -2.441 -3.525 - 1.309 -7.646 -2.197 -3.760 -.961 -7.809 - 1.953 -3.926 -.614 -7.902 - 1.709 -4.053 -.267 -7.954 - 1.465 -4.131 .080 -7.978 -1.221 -4.199 .426 -8.002 - .977 -4.238 ,773 -8.026 -.732 -4.287 1.119 -8.022 -.488 -4.297 1.465 -7.977 -.244 -4.307 1.811 -7.946 ,000 -4.316 2.157 -7.928 .977 -4.326 2.503 -7.883 1.953 -4.346 2.848 -7.849 2.930 -4.375 3.194 -7.820 4.883 -4.414 3.540 -7.789 6.836 -4.473 4.232 -7.740 8.789 -4.521 5.616 -7.559 10.742 -4.580 6.999 -7.433 12.695 -4.619 9.767 -7.168 14.648 -4.678 12.534 -6.862 16.602 -4.736 15.301 -6.569 18.555 -4.775 18.068 -6.262 20.508 -4.824 20.835 -5.983 22.461 -4.883 23.602 -5.705 23.680 -4.919 26.369 -5.398 29.136 -5.092 Chord = 61.445 in. Chord = 43.320 in

14 Table 5. Modified-Airplane Stall Characteristics

4, P, Description of maneuver Result -Power deg deg Stable; slight nose bobble IC 0 -5 Slow deceleration to lg stall, then controls fixed

0 -1 to4 Slow deceleration to lg stall, then Stable; slight left-roll tendency easily wheel full back and controls fixed controllable with small right aileron and rudder inputs; very docile

0 0 From level flight near stall, wheel Right or left-roll tendency controlled abruptly pulled full back and controls with small aileron and rudder inputs fixed opposite to roll

0 2 Slow deceleration to lg stall, then No roll-off tendency wheel full back and controls fixed, then use ailerons and rudder

0 -18 lg stall with left sideslip (full No roll-off tendency; controllable right rudder)

-30 -20 Stall from slipping left turn (full Slight roll-off to the right; easily right rudder) controllable

-45 10 Stall from skidding left turn (3/4 left Airplane stays in the bank; controllable rudder)

-30 2 Stall from left turn Mild roll to the left 4 sec after stall; countered with rudder and aileron inputs

-30 0 Slow deceleration to stall in left turn, Very docile then wheel full back

30 4 Slow deceleration to stall in right Very docile turn, then wheel full back

0 -3 Pull-up to 1.59 accelerated stall No roll-off tendency; docile

For level 0 f 12 Wheel back to stall, then use ailerons Docile Ai and rudder

0 14 Wheel full back at stall with No large angular rates produced right sideslip

0 - 10 Wheel full back at stall with Very docile; only small roll and yaw left sideslip (full right rudder) oscillations

30 2 Stall from right turn Rolls to left

- 30 -2 Stall from left turn Rolls to right

30 - 10 Stall from skidding right turn Rolls to right

-30 -13 Stall from slipping left turn Rolls to right

15 Table 5. Continued

6, P, Power deg deg Description of maneuver Result Maximum 0 -2 Slow deceleration to lg stall, then Very docile; rolling tendency wheel full back controllable with ailerons; can control bank with ailerons and a little rudder

0 2 Wheel abruptly pulled to full back Easily controlled and held

0 - 10 lg stall with sideslip (full right rudder) Tendency to roll off to right countered with ailerons

30 3 Stall from right turn, then wheel full back Slight right-roll tendency easily countered with normal controls; good control

-30 -2 Stall from left turn, then wheel full back Roll to the right easily countered

30 -12 Stall from skidding right turn, then Roll to the right wheel full back

-30 12 Stall from skidding left turn, then Mild roll-off tendency; controllable wheel full back with ailerons; good control

30 13 Stall from slipping right Rolls level turn, then wheel full back

-30 -7 Stall from slipping left turn, then Roll to the right; controllable with wheel full back ailerons; good control

60 2 Stall from right turn, then wheel full back Slight roll off to right, mild maneuver

-60 -4 Stall from left turn, then wheel full back Slight right-roll tendency; easily controllable

60 - 12 Stall from skidding right turn Slight right roll-off tendency; easily recovered with ailerons

-60 7 Stall from skidding left turn Slight roll-off tendency; easily controllable

0 2 Pull-up to stall, then wheel full back Pronounced break; good control

16 Table 5. Concluded

4, A Power deg deg Description of maneuver R.esult Idle 0 -2 Slow deceleration to lg stall, No stall warning; mild departure then wheel full back tendency; easily controlled with small rudder and aileron inputs; good control

30 3 Stall in right turn, then wheel full back Docile

- 30 -3 Stall in left turn, then wheel full back Docile

0 2 Rapid deceleration to stall, then wheel Docile full back

For level 0 -4 Slow deceleration to lg stall, then Docile; follows aileron input wheel full back with full right rudder and left aileron

0 12 Slow deceleration to lg stall with full Left roll and yaw changed to right roll and yaw left rudder and right aileron, then at full back wheel apply opposite rudder

30 14 Stall from slipping right turn (full left rudder) Rolls to left

-30 -15 Stall from slipping left turn (full right rudder) Rolls to right

Y 0 0 Slow deceleration to lg stall, then Very light stall break; a little more roll wheel full back tendency than with idle power and harder to control, rudder fixed; readily controlled with coordinated rudder and ailerons

30 1 Stall from right turn, then wheel full back Right-roll tendency easily countered

-30 0 Stall from left turn, then wheel full back Left then right-roll tendency easily countered

30 -8 Stall from skidding right turn Gentle roll to the right

-30 9 Stall from skidding left turn Roll to left easily recovered with rudder

Y 0 1 Zoom to stall. then wheel full back Good control

17 Table 6. Conditions for Spin Entry of Modified Airplane

Center-of- Number of Number of gravity seconds pro- turns pro- Number of position, spin input spin input turns for 2 Power - Spin-entry maneuver maintained maintained recovery Direction 0.1 Zoom to stall; left rudder pulse MU 10.5 21/ 2 2 Left prior to stall; pro-spin rudder and full trailing-edge-up stabilator at stall; ailerons against 1.2 sec after pro-spin input

Zoom to stall; left rudder pulse 7.2 13/ 4 3 Left prior to stall; pro-spin rudder and full trailing-edge-up stabilator at stall; ailerons against 1.5 sec after pro-spin input

.2594 Zoom to stall; pro-spin rudder and full 17.0 411 8 111 8 Left trailing-edge-up stabilator at stall; ailerons against 3.2 sec * after pro-spin input

,2778 Zoom to stall; full pro-spin rudder 12.5 21/ 2 11/ 2 Left and trailing-edge-up stabilator at stall; ailerons against 2.8 sec after pro-spin input

,2823 Zoom to stall; pro-spin rudder and full 6.7 31 4 3/ 4 Left trailing-edge-up stabilator at stall; ailerons against 2.0 sec after pro-spin input

Zoom to stall; pro-spin rudder and full 10.7 2 11/ 2 Left trailing-edge-up stabilator at stall; ailerons against 2.9 sec 1 " after pro-spin input

,2787 lg, wings-level stall; pro-spin rudder 27.1 81/ 2 2 Right and full trailing-edge-up stabilator at stall; ailerons neutral

,2790 Stall in 30' banked right turn; full 10.5 2l/ 4 1/ 2 Right pro-spin rudder and trailing-edge-up stabilator at stall; ailerons neutral

,2842 22.2 731 4 15/ 8 Right

,2998 lg, wings-level stall; full pro-spin 25.8 8 1'1 2 Right rudder and trailing-edge-up stabilator at stall; ailerons neutral

,3207 19.8 5'71 8 11/ 2 Right

29.0 9 15/ 8 Right

" 1 23.7 6 Chute Left

18 ......

19 N00-0m3-4m

20 ‘b

Figure 1. System of body axes. Arrows indicate positive direction of quantities.

21 M

.ec 3

22 I I\ I ws 207.40

n >I I

IL. 35.43 I

0.26 WL 38.15 1- 1- 1- 21.80 -=

Figure 3. Three-view drawing of test airplane. Dimensions are in feet.

23 _------r

Airfoil section at 0.53 b/2 Airfoil section at 0.96 b/2

Leading-edge modification (a) Diagram.

(b) Photograph. L-83-5680 Figure 4. Wing-leading-edge modification.

24 2900 2800 00 2700 2600 2500 2400 E 2300

c, LI Ol .r al 3

n" 10 20 30 40 Center-of-gravity position, percent E

Figure 5. Airplane design center-of-gravity envelope showing loadings at test altitude.

0 20 40 60 80 100 ampdeg

Figure 6. True angle of attack as a function of measured angle of attack at wing-tip boom location.

25 125 125 loo 100 75 75 4 n 50 50 - w' o 25 I 25 d E 00 -. LL O2 0 -25 -25 (3 3 -50 -50 $ -1 -75 -75 -100 -100 -1 25 -1 25

Figure 7. Slow deceleration to idle-power, lg, wings-level stall with flaps and gear retracted. Test weight = 2438 lh; c.g. = 0.27772..

26 0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 75 40 30 b 430 0 W Stall -0 10 10 c3 \ .* f\ ,nd- __-_- a- - -y _. 0 0 00 2 -10 -10 2 :-20 -20 : -30 -30 -40 -40

200 200 175 175 150 150 125 125 100 100

0 75 75 0 $ 50 50 $ 0 25 25 0 -_--e-. 9 2 y---7--A.*.\.-+I-_-_-* 02 2 -25 -25 2

2 -50 -50 2 > -75 -75 2- -100 -1 00 -125 -1 25 -150 -150 -175 -1 75 -200 -200

0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 75 TIME, sec Figure 7. Continued.

27 0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 75

3000E -Modified wing 25mP 2000 wd 15OOW rooo~ n 500 0 CY On

30 __-_-----z5 m /-- 20 = / C I 15'- I I 10 ------A 5= 0

I---- I ,/'

30 j30 20 g lou

\-- 0g -1oyCY

-20 a -30 n 4 50t

200 200 150 i 150 i 100 100 > 20) u50 I_ 50 4 eo - .--e. t. *- -- OP 2 -50 -50 2 -1 -100 -100 a -1 -1 g -150 -150 g -200 -200

1.0 O1 I ------__-- .5 0' __.- 02 -.5 p -1.00,

0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 75 TIME, sec

Figure 7. Continued.

28 0 5 10 15 20 25 30 55 40 45 50 55 60 65 70 75

14000 1- 13000 13ooo 12000 12000 11000 11000 10000 1M)o

z -100 -100 0: -150 I3 ;-150 I- -200 -200 -250 -250

-1 F 4 -1

0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 75 TIME. sec

Figure 7. Concluded.

29 PRECEDING PAGE BLANK !COT FiiMED 31 0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 75

0 =4 *OCj 30 4: 20 I 103 0

15 0 10 4 5cT' 0 _.-- Ot 4: -5 1 m -10 2 -15

125 125 100 100 75 75 e n 50 50 - J W' o 25 25 o E -. IY LL00 - .-- T -25 -25 a 4 -50 -50 '1 -75 -75 -1 00 -100 -1 25 -125

100 100 80 80 p 60 60

-2400) 40k u- 20 20 + W Po Ot 4 4: E -20 -20 [L 5 -40 -40 5 I- E -60 -60 -80 -80 -100 -100

0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 75 TiME. sec

Figure 9. Idle-power, Ig, wings-level stall with maximum stabilator deflection, flaps and gear retracted. Test weight = 2438 lb; c.g. = 0.2777C. 0 5 10 15 20 25 30 35 4.9 45 50 55 60 65 70 75

40 40 30 30 p 20 20 !g -0 -0 10 10 cj a- 00 OU 2 -10 -10 3 g -20 -20 : -30 ..30 -40 -40

30 30 rn 20 20 !g -0 10 10-0 go 0g 0 0 0 -10 -10 0 3 3 rY -20 -20 E -30 -30

200 200 150 150 0 100 loop 50 50g

KO 08 2 LL -50 -50 n n 2 -100 -1002 -i50 -150 -200 -200

200 1 75 150 12s loo 75 g 50 $ 25 9) 0; -25 2 -50 3 -75 >

"' .5 .5 "' Yo ! - - 0s:

I-4 -.5 -.5 5 J -1.0 -1.0 -I

0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 75 TIME. sec Figure 9. Continued.

33 0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 75

3000 E 2500 E 2000 wd 1500 1000 v, a 500 0 0:

25 IT 20 = C 15'- /-- I --- lo E 5= 0

1000 750 9 .500t-- v) ___---- 250 2 I __.-- OF -250

50 30 g 20 20 gl 10 10 go u- 0g IL IL w -10 -10 w =I =I 4: -20 -20 4: -30 -30 9 n 50 50 -- 8 25 25 8 IL 00 08 LL -25 -25 LL I- 5 -50 -50 2-1 200 200 150 150 0) 2 100 1002 9) 9) 050 500

I I_ fL Llio - ~, w- - ~ 0 w- --/------t- 2I- -50 -50 2 -1 -100 -100 -1 -1 -I $ -150 -150 $ -200 -200

IT 1.0 1.0 * uv- .5 .5 u' I - 0 4:o ----- 04: (3 g -.5 -.5 z 9 -1.0 -1.03

0 5 10 IS 20 25 50 35 40 45 50 55 60 65 70 75 TiME. sec Figure 9. Continued.

34 14OOO 14OOO 13000 13000 12000 12000 11000 11000 loo00 10000 9m I 9000 c c --- 7 w- 8000 \*- - - - - I------8000 w'

7000 7000; 2 6000 F 5 6000 1 -l 5000 5000 -l 4000 Woo0 3000 3000 2000 2000 1000 1000 0 0

250 250 200 200 f 150 150 f 20 100 1002 XJ 50 50 XJ w- 0 0 p- l- 2 -50 -50 2 z -100 -100 z w w -3 -150 -150 I- 2 -200 -200 -250 -250

Figure 9. Concluded.

35 0 5 10 15 20 25 30 35 40 45 60 DP 50 d4 Stall 40 "- 30 30 20 20 2 a 10 10 -I 0 OQ

Ol- Q m -5 m=' 2 -10 -10

75 75 &? n 50 50- w' 0 25 25 K K 00 LL " -25 -25 5 -50 -50 5 1 1 -75 -75 -1 00 -1 00 -1 25 -1 25

100 i100 80 80 60 60 f 240 40 2 20 I 20 4- I-u. W Q V Ok 20 Q -20 -20 111. 6 -40 -40 6 l- -60 -60 E -80 -80 -1 00 -100

0 5 10 15 20 15 30 35 40 45 TIME, sec

Figure 10. Idle-power stall of modified airplane from 30O banked left turn with flaps and gear retracted. Test weight = 2677 lb; c.g. = 0.28232.

36 0 5 10 15 20 25 30 35 40 45

Stall v 10 10 cj 0 00 3 -10 -10 2 :-20 -20 : -30 -30 -40 -40

50 50 F 20 20 g - 10 10- $0 O$ n n 0 -10 -loa 3 3 IY -20 -20 IY -30 -30

200 200 150 150 n n - 100 100 - w'0 50 5og 80LL o! -50 -50 n n 2 -100 -1002 -1 50 -1 50 -200 -200

200 200 175 175 150 150 125 125 100 100 75 75 0 50 50 25 25 P9) 0 OF -25 -25 2 -50 -50 3 Q: -75 -75 *

; l:!i

':!!I !- ':!!I -.5 I- -.5 Q Q 1 -1.0 -1 .o J

0 5 10 15 20 25 30 35 40 45 TIME, sec

Figure 10. Continued.

37 0 5 10 15 20 25 30 35 40 45 ::::El ::::El $ 2000 d dW 2000 Stall 15M) I v, 1000

OF 40

-250 4 -250

200 200 p 150 150 f 2 100 l00k Q) 0) TI 50 500 w- 0 0 w- I- I- 2 -50 -50 2 A -100 -100 -1 -I J -150 -150 $ -200 -200

1.ot 3 1.0 o, .5 .5 u' 0'0 I 0 4:o OQ Z" -.51 5 -1.0 I-.."-1.00,

0 5 10 15 20 25 50 35 40 45 TIME, sec Figure 10. Continued. 14000 14OOO 13MM 15ooo 12000 12000 11000 11000 loo00 I 10000 ;r QOOO - 9000 ;= J 8000 -Li 2 7000 7000 clz 6000 F 5 6000 45oOo 5000 2 1ooo 4ooo 3MM 3000 2000 2000 loo0 loo0 OF 0

-1 4 -1

0 5 10 15 20 25 30 35 40 45 TIME, sec

Figure 10. Concluded.

39 0 5 10 15 20 25 Jo 35 40 45 50

6o m 50 4 406 30 20 a 10J 4: 0

100 100 80 80 $ 60 60 f $40 40 > 4. 20 20 + Id A/ LI I-0 Oh 4: Q ly -20 -20 ly 6 -40 -40 5 k I- a -60 -60 E -80 -80 -100 -100

0 5 io 15 20 25 30 35 40 45 50 TIME, sec

Figure 11. Idle-power stall of modified airplane from 45O banked skidding left turn with flaps and gear retracted. Test weight = 2677 lb; c.g. = 0.2823C.

40 30 50 Stall 20 20 z U -0 10 10 u- cj 00 00 2 -10 -10 2 wm -20 -20 -50 -30

4 100 100 -n w' 0 50 50 [II 00 0; L r- -50 -50 n Q 2 -100 -1002

200 175 150 125 100 0 al 75 $ 50 $ sal 25 e 9 t- 4: 02 IY -25 2 3 -50 -50 3 4: 4: * -75 -75 * -100 -1 00 -1 25 -1 25 -1 50 -1 50 -175 -1 75 -200 -200

1.0 m

z -.5k +5l- 4: -1 -1.0 -1.0-1

0 5 10 15 20 25 30 35 40 45 50 TIME, sec

Figure 11. Continued.

41 3000 E 2500 ? Stall 2000 w wa‘ 2 1500 1500 1000 1000

t--v) 500 500 2 250 1 250 2 Fo OF -250 -250

it.10 OQ CY lY w -10 -1ow =! -J 4 -20 -20 a

n - 50 n 50, 8 25 25 4:

IL80 08 -25 -25 I- LL -50 -50 -J J2 200 200 p 150 150 > 100 100 > 50 50 -0” d I Lio fi Ow’ t- I- 2 -50 -50 2 -100 -I 2 , !-I50 d Lx -200 -200

O1 1.0 0’ .5 I 0 Qo $ -.s 9 -1.0

020 TIME, sec

Figure 11. Continued. 0 5 10 15 20 25 50 35 40 45 50 300 300 250 250 $ 200 200 3 150 d 150 d w 100 100 W y 50 ma 0 0

13000 12000 11000 10000 10000 e 9000 9000 ;r

2000 1000 "EOF 40

250 200 0 150 E look 50 'T3 0 p- -50 2 -100 z -150 -200 -250

-1 LA-1

0 5 10 15 20 25 30 35 40 45 50 TIME, sec

Figure 11. Concluded.

43 0 5 10 15 20 25 30 35 40 45 50 55 60

60 U 50 Stall 40 40 30 30

20 20 10 10 -1 4: 0 0

,- Ob Q -5 -I m -103 -15 -15 v,

lL -25 -25 a 3 -9 -50 5 -1 -75 -75 -1 00 -lW -1 25 -1 25

100 100 80 80 $ 60 60 m 40 40 > al 4- 20 20 u- w -.--e---.* - Ok PoQ < (L: -20 -20 CK

-1 i= -1

0 5 10 15 20 25 30 35 40 45 50 55 60 TIME, sec

Figure 12. Maxiiiium-power, lg, wings-level stall of modified airplane with flaps and gear retracted. Test wciglit = 2438 11): c.g. = 0.2777C.

44 0 5 10 15 20 25 30 35 40 45 50 55 80

40 40 30 30 p 20 20 p -0 Stall -0 10 10 0- cj 00 00 3 -10 -10 2 ;-20 -20 ; -30 -3 -30

-150E. 4-150

200 200 175 175 150 150 125 125 100 100

8 75 75 8 $ 50 50s 0) 25 25 0) .. -- -- I I.- A- / 0; 2 0 2 -25 -25 2

2 -50 -50 2 t -75 -75 > -100 -100 -1 25 -1 25 -1 50 1 1;:; -175 -200 -200

0 1.Ok j 1.0 07

0 5 10 15 20 25 30 35 40 45 50 55 60 TIME, sec Figure 12. Continued.

45 0 5 10 15 20 25 Jo 35 40 45 50 55 60

OF 30

1 000 I 7 1000

-250 p 3 -250

F 20 20 F -0 10 10- go -/I 0s K CY w -10 -1ou 2 r! Q -20 -20 4 -30 L 3 -30 9 n 50 J 50- 8 25 25 e 80 08 -25 -25 LL I- I- Q: -50 -50 Q 1 -1

@o _A A __ n. v'-v " O@ 2 -50 -50 2 1 -100 -100 -1 -I 1 -150 -150 g -200 -200

Ln 1.0 1.0 d .5 I .5 0' 0 *o 02

0 5 10 15 20 25 30 35 40 45 50 55 60 TIME, sec

Figure 12. Continued. 1- 14000 130oo 13OW 12000 1 2000 11000 1 1000 loo00 loo00

9000 9000 'f: w'-a -w' 3 7- 70005 I I- 5 6000 6000 F -1

1000 7 loo0 OF 0

250 250 200 200

-250 F -250

5 5 04 40 63 30' 22 22 m1 &-- W lz KO OK -1 -I

0 5 10 15 20 25 30 35 40 45 50 55 60 TIME, sec

Figure 12. Concluded.

47 0 5 10 15 20 25 30 35 40 45

Stall

a 0

15 15 rn -0 10 10 4 cf- 5 5 d 0 eo - Ob U -5 -5 =I m4 m 3 -10 -10 2 -15 -15

LL -25 -25 c3 3 -50 -50 5 -1 -75 -75 -100 -100 -1 25 -1 25

100 100 80 80

f 60 60 f

> 40 @\, 4- 20 20 4. w W I-0 !-- - Ob- Q -- Q cf -20 -20 lx 5 -40 -40 5 E -60 -60 I-E -80 -80 -100 -1 00

-1 4 -1

0 5 10 ;5 20 25 30 35 40 45 TIME, sec Figure 13. Maximum-power stall of modified airplane from 30' banked skidding left turn with flaps and gear retracted. Test weight = 2677 lb; c.g. = 0.2823~

48 0 5 10 15 20 25 30 35 40 45 40 ed40

m Stall al D 10 10 u- cj 0 OU 2 -10 -10 2 ;-20 -20 ; -30 -40 1 1:: 30 p 20 20 0al

7J 10 10'6 E. A wo YV 0g n a -10 -100 3 3 E -20 -20 rY -30 -30

200 200 150 150 9 100 100 g d 50 50 e 80 OF LL -50 -50 n n 2 -100 -100 2 -1 50 -1 50 -200 -200

200 200 175 175 150 150 125 125 100 100 75 75 0 so 50 z 25 25 I 9 0 .- -^---- OF -25 -25 9 -50 -50 3 6 -75 -75 >

m 1.ot 4 1.0 m "- .5 .5 o- 0 Q:o I- 0: t-< -.5 -.5 2 -1 -1.0 -1.o-I

0 5 10 15 20 25 30 35 40 45 TIME, sec Figure 13. Continued.

49 0 5 10 15 20 25 30 35 40 45

E 3000 3000 E 2500 2500 f! d 2000 2000 d W W 1500 1500 2 v, 1000 1000 Q 8 500 500 0 IY no OE

30 30 m 25 25 m = 20 20 = c c '- 15 15'- g 10 10 g =5 5= 0 0

30 j 30

P n 50 50 -. 25 25

80 08 -25 -25 LL I- < -50 -50 1 2-J 200 200 f 150 150 f -& 100 100 -& Q) U 50 50 4 li 0 0LI.i I- I- 2 -50 -50 2 -I -100 -100 -1 -J -I -150 -150 -200 -200

1.og 3 1.0 *

005 TIME, sec Figure 13. Continued. 0 5 10 15 20 25 30 35 40 45 300

i4000 1 4000 13000 13000 1 zoo0 12000 11000 1 1000 10000 10000 g goo0 9000 ;= w' 8000 I 8000 w- 2 7000 7000 2 - I- E 6000 6000 F 1 4 5000 5000 4 4000 4000 3000 3000 Zoo0 Zoo0 loo0 1000 0 -0

250 250 200 200 0 150 150 $ 100 100 -&

50 50 Q 0 0 E* -50 -50 2 z -100 -100 z n: n: 3 -150 -150 3 t- t- -200 -200 -250 -250

w 0 1 ow -3 3 -1

0 5 10 ;5 20 25 50 35 40 45 TIME, sec

Figure 13. Concluded.

51 I

0 5 10 15 20 25 30 35 40 45 50 55 60 65 70

-Modified wing 6o rn ---Basic wing 40cj -. 30 '\ ,*\ 20 2 ',I a '. 10 2 0

125 125 100 100 75 75 a e 50 50 - W- o 25 25 lY of 00 Eo LL -25 -25 5 -50 -50 5 -J -1 -75 -75 -1 00 -1 00 -1 25 -1 25

100 100 80 80 0 0) 60 60 $ s40 40s Q) -u- 20 20 s W two Ot- a 4 of -20 -20 Iy. 5 -40 -40 = E -60 -60 -80 -80 -1 00 -1 00

0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 TIME, sec

Fig-iirc 14. Maximum-power stall with flaps extended 40' and gear retracted. Test weight = 2438 lb; c.g. = 0.2777T.

52 0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 40 30 20 U 10 ci 00 -103 -20 # -50 -40

50 20 p low .- O$ n -100 3 -20 E -30

200 200 150 1SO -D 100 100 g 50 5og E 00 0; L -50 -50 n n 2 -100 -1002 -150 -lso

200 200 175 175 150 150 12s 12s 100 100 0 75 75 8 1 50 M$

a, 25 25 0) 9 9 EO OP 2 -25 -25 3 -50 -50 3 4: 4: 2- -75 -75 > -1 00 -100 -1 2s -1 25 -1 SO -lso -1 75 -175 -200 -200

1SI 1.0t j 1.0 m 2 .5 .5 0- Yo I 02 I- -.5 -.5 I- 4: 4: -1 -1.0 -1.0 -1

0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 TIME, sec

Figure 14. Continued.

53 3000 E 2500 2000 d W - Modified wing 1500 ---Basic wing 1000 v, a 500 0 OK

30 I -- 30 I m 25 25 0 = 20 20 = C c ’- 15 15’: g 10 10 =5 5= 0 0

lo00 -___-I ~ .\.--.* 750 9 500 250 2 OF -250

30 20 lov 0g -lowK -1 -20 a -50 P n 50 50 -- 8 25 25 8

LL80 08 -25 -25 LL I- I- < -50 -50 < -I 1 200 200 f 150 150 f 2 100 look 9) 9) V 50 50 U

I-w- 0 Ow’I- 2 -50 -50 $ -I -100 -100 -1 -J 1 -150 -150 $ -200 -200

1.0

0‘ .5 I -_-____-_&------%---‘ A -. 0 Qo g -.5 9 -1.0

020 TIME, sec

Figure 14. Continued. 14000 14OOO 13000 13000 12000 12000 1 1000 1 1000 10000 1 0000 g 9000 9000 W' 8000 8000 w- n 7000 70005 I- z 6000 6000 F I -I 2 5000 5000 Q 4000 4000 3000 ------__ 3000 2000 ------______-----__ 2000 1000 1000 0 0

Figure 14. Concluded.

55 0 5 10 15 20 25 30 35 40 45 50

- 80 t I 3 80 24- IFlaps retractedjso:

125 125 100 I 100

I 6 -40 -40 6 I- -60 -80 - ; 4 -80 -80 -100 i -100 5 04 93 ip42 20 -1

0 5 10 15 20 25 30 55 40 45 50 TIME, sec

Figure 1.5. Maximiini-power stall of modified airplane from 30' banked skidding left turn with flaps extended 40' and gear retracted. Test weight = 2438 lb; c.g. = 0.2777C. 0 5 10 15 20 25 30 35 40 45 50

I Flaps retracted 50 20 g U U 10 10 a- cj 00 00 2 -10 -102 :-20 -20 : -30 -30 -40 I -40

-30 I

200 t I 150 100g 50 3 -n o! -50 n -1002 -150

200 175 150 125 100 g 75 cn 50 '.m 0) 25 I I-eo - 2 -25 2 -50 > -75 -100 -125

-175-'..I -200 L

! j 1.0 m

-.5 -.5 !- 5 -l -I -1.0 -1.o-I

0 5 10 15 20 25 30 35 40 45 50 TIME, sec

Figure 15. Continued. 0 5 10 15 20 25 30 35 40 45 50 I I 3000 E ~ A2500 ?' Stall I Flaps retracted iooo("

SOt I 07 25 I I 25 07 = 20 20 = S C '- 15 15'- g 10 10 $ =5 5= I

go~~~i'' 0g K K w -10 -10 w -1 -1 Q -20 -20 7

-25 -25 c I I- Q -50 -50 Q

1.0 1.0 07 d .5 .5 0 !- 0' Qo I 02 $ -.5 -.5 za 9 -1.0 i -1.09

0 5 10 15 20 25 50 55 40 45 50 TIME, sec

Figure 15. Continued. 0 5 10 15 20 25 30 35 40 45 50

500 300 0 250 c 200 Stall 150 150 d d y 100 100 & 50 0 0

1 4000 kp 14000 13000 13000 12000 12000 1 lo00 1 1000 1 DO00 10000 9000 g go00 I c w- J 8000 - 8000 7000 7000 5 I-5 I- F: 6000 6000 F -1 1 4 5000 5000 4 4000 4000 3000 2000 1000 0

250 200 0 150 100 > 0) 50 - --- I OF -50 2 -100 z K -150 3 I- -200 -250

TIME. sec

Figure 15. Concluded.

59 75 75 .D n 50 50 - w' Li v 25 25 o tt LT 00 00 LL LL -25 -25 c3 5 -50 -50 5 -I -75 -75

100 100 80 80 60 60 c40 40 >a :* :* 20 20 '0- - w ako Ol-Q -20 -20 [1: 6 -40 -40 6 at -60 -60 I- -80 -80 -100 -1 00

5t _IC

0 5 10 15 20 25 30 35 40 45 50 55 TIME, sec Figure 16. Comparison of response of basic and modified airplane to pro-spin control inputs at idle power, ailerons neutral, flaps and gear retracted. Test weight = 2438 lb; c.g. = 0.27771..

60 0 5 10 15 20 25 30 35 40 45 50 55 :::j :::j :: g 20 20 g D D 10 10 ci a- 00 ov ;rl -10 -103 2 -20 -20 2 -30 1 3 -30 -40 Spiral turns I 1 I -40 Spin turns I I I I I 1 1

200 3 200 150 n 100 - 50 8 oz -50 n -100 2 -150

100 75 8 50 < cn 25 Q) 9 OF -25 2 -50 3 Q -75 >

cn 1.0- 3 1.0 cn u- .5 .5 o- Yo 02 -.5 -.5 I-4 2 -J -1.0 -1.0 -J

0 5 10 15 20 25 30 35 40 45 50 55 TNE, sec Figure 16. Continued.

61 0 5 10 15 20 25 30 35 40 45 50 55

3000 E

1500 1500 2 -.. 2 In 1000 1000 In R --____--* a 0 500 Spiral turns 500 0

Spin turns .

-250 3 -250

g 20 20 g TI 10 10 go 0g lY lY w -10 -10 LJ A 1 u -20 -20 4:

200

I- 150 f 100 -& 50 % 0 w- t- -50 2 -100 1 -.J -150 -200

cII 1.Ok -1 io"

c3 ? -.5k +-$ 5 -1.0

0 5 10 15 20 25 30 35 40 45 50 55 TIME, sec Figure 16. Continued. 0 5 10 15 20 25 30 35 40 45 50 55

300

3 50 tr, M 5c Spiral turns 1.. 0 0

Spin turns ' 1 I . ! I 14000 i pd 1 4000 13000 12000 12000 13000111000 1 1000 ------_-- - - .-- -. loo00 10000 9000 9000 g ,; ,; 8000 8000 w- 7000 7000 CI I- 2 F 6000 6000 F= -J 4: 5000 5000 4:

2000 1000 3 0 250 250200 E----- 200 -4 -4 0 150 150 100 100 al 50 50 Q 0 0 p- -50 -50 2 -100

-200 -250 -250

5c 35

-1 F q -1

0 5 10 15 20 25 30 35 40 45 50 55 TIME, sec

Figure 16. Concluded.

63 0 5 10 15 20 25 30 35 40 45 50 55 60 85

6o 0 -4 40c5 30 20 2 n \-----10 2

CY- 5 5 CY' 0 20 Ol- Q: 4E -5 -5 =! m 2 -10 -10 2 v, -15F q-15 in

125 100 75 0 50 w- 25 o CY

-25 c3 -50 6 -J -75 -100

100 100 80 80 60 60 f

$400) 402 u- 20 20 5 w I-0 Q: O: CY -20 -20 6 -40 -40 3 I- -60 -60 -80 -80 -100 -1 00

5 5 "4 40 00'3 3g 42 2Q: r I CY1 ICY 0 0 20 02 -1 -1

0 5 10 15 20 25 30 35 40 45 50 55 60 65 TIME, sec Figure 17. Comparison of response of basic and modified airplane to pro-spin control inputs at idle power, ailerons with the spin, flaps and gear retracted. Test weight = 2438 lb; c.g. = 0.2777C.

I 64 200 200 175 175 150 150 125 125 100 100 $ 75 75 z0 2 50 "\, s 25 25 9a) O--- .-.?

I \ ,I -50 3 2 -50 I Q > -75 \ -75 2- \\ ,,\",--q.--/ -+/-v -1 00 -1 00 -1 25 -1 25 -1 50 -1 50 -1 75 -1 75 -200 -200

1.0 0 .5 "' .-\e- .7 ,\- '.' - 0: -.5 I-< -1.0 -I

0 5 10 15 20 25 30 35 40 45 50 55 60 65 TIME, sec

Figure 17. Continued.

65 0 5 10 15 20 25 30 35 40 45 50 55 60 65

1500 1500 / I n \------a 0 500 500 0 IY SDin turns 10 30

OF 30

lo00 1 000 0 750 750 4

0’ .5 .5 0’ 0 ---- 0 Qo OQ a g -.5 -.5 z 9 -1.0 -1.00,

005 TIME, sec

Figure 17. Continued. 0 5 10 15 20 25 30 35 40 45 50 55 60 65

13000 12000 11000 10000 QOOO g 8000 7WoF 6000 F 5000 2 4000 5000 2000

~ -j 1000 0 0

250 250 200 200 0 0 9) 150 150

P) 100 1002 s P) eV 50 50 V 4 CK

5b $5 0 4 40 0' 3 30' 0 4 2 22 cn W 1 'Y CK 0 Om -1 I= 4 -1

0 5 10 15 20 25 30 35 40 45 50 55 60 65 TIME. sec

Figure 17. Concluded.

67 0 5 10 15 20 25 30 35 40 45 50 55

--Wodified wing/spi ral 6o 01 =$ 'OCj

125 125 100 100 75 75 P 50 50 Li wi o 25 25 o [2: fK f0 O2 -25 -25 5 -50 -50 5 -75 -75 -100 -100

60 10 20 0 -20

-1ooE

5 5 04 401 G30 38 (2 24 I 1lY $1 0 20 02 -1 -1

0 5 10 15 20 25 30 35 40 45 50 55 TIME, sec

I Figure 18. Comparison of response of basic and modified airplane to pro-spin control inputs at idle power, ailerons against the spin, flaps and gear retracted. Test weight = 2438 lb; c.g. = 0.27772.

I 68 0 5 10 15 20 25 x) 35 40 45 50 55 40 40 50 30 g 20 20 g -0 U 10 10 6 go ov 2 -10 -102 g -20 -20 g -50 -30 -40 -40 Spin turns I I I I I I 1

I I J

200 150 100s 506 OF -50 n -1002 -1 50 -200

200 175 150 125 100 75 $ 50s

25 Q) 0; 4: -25 [I: -50 2 -75 >

rn 1.Ok 4 1.0 m .5 "' - 0: -.5 5 -1.0 -1

0 5 10 15 20 2530 35 40 45 50 55 TIME, sec Figure 18. Continued.

69 0 5 10 15 20 25 30 35 40 45 50 55

3mm E 3mm 3000 E -Modified wing/spiral \ p- j25CQ P j2000k 't, \ ---Basic wing/spin I 4 2000 d

OF 30

lo00 lo00 0 750 750 0

OE -250 F 4 -250

30t 330 g 20 20 7J 10 10 7J go 0g [L [L kJ -10 -low A a -20 -20 4

200 150 loo\, 0) 50-

O@ -50 2

1.0t 4 1.0 .5 u' 02 -.5 $ -1 .o 9

0 5 10 15 20 25 50 35 40 45 50 55 TIME, sec Figure 18. Continued.

70 0 5 10 15 20 25 30 35 40 45 50 55

OF 40

250k , j 250

0 9)

-s9) D kj 6 OI -100 2 -150 -200 -250 -250

0 5 10 15 20 2530 35 )o 45 50 55 TIME, sec Figure 18. Concluded.

71 125 125 100 100 75 75 -n n 50 50 - d LLi u 25 25 u [I: LL: Bo O2 -25 -25 a 5 -50 -50 5 -75 -75 -100 -1 00 -1 25 -1 25

100 100 80 80 60 60 p 340 40-2 v. 20 20 < w lLl I-0 OI- Q 4: [I: -20 -20 [I:

050 TIME, sec

Figure 19. Comparison of response of basic and modified airplane to pro-spin control inputs at idle power, ailerons neutral, flaps deflected 40°, gear retracted. Test weight = 2438 lb; c.g. = 0.2777C.

72 0 5 10 15 20 25 30 35 40 45 50 55 60

40 30 m al -0 d V

-30 4 -30

200 150 n 100 - 50 0: 0: -50 n -100 2 -!50 -200

200 175 150 125 100 g 75 75 g $ 50 50 $ 9Q 25 25 4. woI- OF 2 -25 -25 2 9 -50 -50 3 -l < > -75 -75 + -1::::I 00 1i,-1 00 -1 75 -175 -200

0, 1.0 1.0 0, "' .5 -- .5 o' 20 0: 2 -.5 -.5 2 -1.0 -J -I -1.0 ~

TIME, sec

Figure 19. Continued.

73 -Kodified wing/spiral

2000 d

1000

= 20 C 15-

OF 40

-250 F 3 -250

0 n 50 50 25 25 8 80 OW L -25 -25 2 I- $ -50 -50 Q: 1 -J ,. 200

* 1.0k j 1.0 *

090 TIME, sec Figure 19. Continued. 0 5 10 15 20 25 30 35 40 45 50 55 60

$! 300 $! 300

12000 l3I11OOo ---- -.- .

ZOO0 1000 "I 40 OF

250 250 200 200 0 150 150 100 looh 50 504 b 0 oe -50 -50 2

3 -150 I- -200

5 4ol 30' 22 'Y O@ 4 -1

0 5 10 15 20 25 Jo 35 40 45 50 55 60 TIME, sec

Figure 19. Concluded.

75 125 125 100 100 75 75 2 -0 50 4 11 4 50 - u- I1 I \ LIJ- I1 ' \ 25 o ;: ; \ K 1 /--\- -. 11It ' L.-4 O2 " -25 !I\ -25 111 -50 I 111 -50 5 I 111 1 I 111 s -75 1 II -75 \A- --. II -1 00 \,,- 1' -1 00 -125 -125

100 100 80 80 0 a 60 60 m v) 40 40 >a -3- 20 20 4- W '-, ., W co U. Ol- Q \' -- Q K -20 -20 K 6 -40 -40 6 I- ak -80 -60 -80 -80 -1 00 -1 00

-1 F

TIME, sec

Figure 20. Comparison of response of basic and modified airplane to pro-spin control inputs at maximum power, ailerons neutral, flaps and gear retracted. Test weight = 2438 lb; c.g. = 0.2777C.

76 11 n I 11 -10 I 11 3 I 11 I II -20 fY

0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 TIME, sec

Figure 20. Continued.

77 3000 E 2500 2000 d - Kodified wi ng/spi ral w 1 500 E 1000 v, Q 500 0 Spin turns I IiII' 1 .----- C 15'- \ 10 L,----- L'

OF 30

1000 750 i, ____----500 $* .- 250 2 I OF -250

30 30 20 20 g U 10 IO -0 go - 0g \--.a 1%. LZI 'd E w -10 -1ow -1 -1 a -20 -20 a

9 n 50 50 -- 25 25 8 80 - .-\ . \a - 08 LL -25 LL I- -25 Q -50 -50 -1 -12 200 200 0 0 150 150 ; v) ,A/* \/-.../-,;'1, 2 100 \; ; 100 2 al ': I Q) U 50 I 50 U \ fi - w-I- 0 b' 0 w- I- 2 -50 -50 2 -1 -100 -100 -1 -l -1 -150 -150 $ -200 -200

1.0 ,- .5 0- -\* J '-_---u 04: -.5 g -1.09

000 TIME. sec

Figure 20. Continued. 14000 13000 13000 12000 12000 11000 10000 10000 < 9000 9000 < \ w* 8000 \ 8000 w' *-. 2 7000 I------7000 2 t I- F 6000 6000 F -1 -J 5000 5000 Q 4000 4000 3000 3000 2000 2000 lo00 1000 OF I10

250 k 3 250 200 150 8 100 2 0) 50 0 F- -50 2 z -100 -100 z ac K 3 -150 -150 3 t t -200 -200 -250 -250

m 0' 0 Q m W ac -1 F

0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 TIME, sec

Figure 20. Concluded.

79 a:5 f rJ A. 0 20 T OC 4 -5 -5 4 2 m ;I -10 -10 2 * -15 -15

100 100 80 a0 $ 60 ln 63 2 40 40 'A 4. 20 20 4. Lrl W I- 0 Ob a 4 E .20 -20 lx I o -40 .40 5 at -60 -60 a'1 -80 -80 -1 00 -100

I 80 0 5 IO 15 20 25 33 35 40 45 50 55 60 65 70 75 do 85 90

40 30 g 20 v 10 uo0- ;r -10 # -20

200 200 I50 150 n - 100 ____------_ 1009 W- 50 111 0 50 0 80LL ' ,,.--' 05 -50 : -50 a J (.e, c3 / 3 2 -100 -100 E -150 -150 -200 -200

k- -.5 -.5 F4 - 5 -:,a3 -1.0 J

0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 T!ME. scc Figure 21. Continued.

81 0 5 10 15 20 25 30 35 40 45 50 55 60 65 -0 15 BO 85 90 ------

ilII 11IlI Right turns I 1 1 I ! I I I 30 --1 10 30

Is, 25 ---___ 25 cn = 20 20 -r c C .- 15 15’- IO 10 >5 5= 0 0

1000 1000 5? 750 750 9 &- 500 - - 500 k- 2 250 250 2 Eo I OF -250 -250

30 30 g 20 20 g U .o 10 -0 - go - -- 0g K u: -rw -10 -1Oy 4 -20 -20 3 -30 -30 n 0 50 50 -. e 25 25 8 LLgo 08 c -25 -25 4 -50 -50 5- 200 200 f 150 150 :: I\ ,A\. r ’, 100 - r\ I I ,00g ,’‘L#, 3 50 ?, so$ i.., 1 w’o -- I- OE 2 -50 -50 $ -1 -100 -100 _1 -1 -I g -150 -150 -200 -200 - ul 1.0 1.0 ul .5 .5 0‘0 0‘ 40 9% 4 -.5 -.5 g 3 -1.0 -1 .o 9

0~0 TIME, sec Figure 21. Continued.

I 82 0 5 10 15 20 25 30 35 40 45 50 55 80 65 -0 75 80 85 90

300 250 rn ‘3 200 1- 150 d lAJ 100 Id & 50 0 Left turns 11 Ill!

1 woo t 13000 E 4! 3000 12000 11000 10005 go00 -w 7000; 6000 F 5000+ 1ooo 3000

OF 40

250 250 200 /-, 200 150 150 f -‘ \-.,/\”/\, f & 100 --,f I loo > 0) / I 0) c 50 ,f-.’ 50 0 go 0 p- -50 -50 z -,00 -100 z E IY ’3 -150 c -150 2 -200 -200 -250 -250

5 4w

-1

0 5 10 15 LO 25 30 35 40 15 bO 55 63 65 -0 75 80 a5 90 TIIIE, sec

Figure 21. Concluded.

83 0 5 10 15 20 25 30 35 40 45 50 55 60 65 70

60 6o 0 -0F 50 504 dW *Od 30 30 p 20 20 s n 4 10 10-1 4 a 0 I IIIIIII 0 Turns 1 8 15 15 -0 10 10 4 c25 5 cr- 0 Po Ol-a -5 -5 =' mI m 3 -10 -10 2 -15 -15(n

125 125 100 too 75 75 -n n 50 50 - m- w' v 25 25 u LT LT 00 N- LL v-- - O2 c3 -25 -25 $ -50 -50 5 _I -75 -75 -i00 -1 00 -1 25 -125

100 100 80 80 60 60 $ v) -240 40 > Q) '9 20 20 + Fo - - n Ol-w *: vv a LT -20 w- -20 LT 6 -40 -40 6 l- -60 -60 -80 -80 -1 00 -1 00

5k 35

-1 F

0 5 10 15 20 25 50 35 40 45 50 55 60 65 70 TIME, sec Figiire 22. Riglit spin of modified airplane at maximum power entered from lg, wings-level flight, ailerons neiitral. flaps and gear retracted. Test weight = 2829 lb; c.g. = 0.2787C.

84 0 5 10 15 20 25 30 35 40 45 50 55 60 65 70

40 30 20 g '0 10 cj 00 -102 -20 -50 I IIIIIIIII -40 Turns 1 8

n :: -10 -100 3 3 II: -20 -20 E -30 -30

200 200 150 150 n 0 100 100 - g50 50 8 2 o! -50 -50 n n 2 -100 -100 2 -1 50 -1 50 -200 -200

200 175 150 125 100 75 $ so$ 25 0 0; -25 2 -50 9 Q: -75 t -? 00 -1 25 -1 50 -1 50 -1 75 -1 75 -200 -200

01 1.0 1.0 rn o- .5 .5 0- 20 02 -.5 t- ;-.s 4 -I -1.0 -l.o-I

0 5 10 15 20 25 50 35 40 45 50 55 60 65 70 TIME. sec Figure 22. Continued. 200 200 f 150 150 f > 100 100 0) 4 50 50 D LLio - 0 12 t- V t- 2 -50 -50 2 -J -100 -100 -I -I J $ -150 -150 -200 -200 Turns 1 8 14000 14000 13000 15000 12000 12000 d 1000 1 1000 0000 10000 9000 9000 g 8000 8000 w' 7000 7000 5 I- 6000 F 6000 -I 5000 5000 4 4000 4000 3000 3000 2000 2000 1000 1000 0 0

5 5 074 407 30' 63 0 22 24:

*1 W 'E =o 0= -1 -1

TlldE, sec

Figure 22. Concluded.

87 look -I 1 00 80 60 40 20 4- OF Q -20 LT -40 6 I- -60 C -80 -100 5 04 g3 62 2 E’ 0 20 -1

OJ5 OJ5 TIME, sec Figure 23. Right spin of modified airplane at maximum power entered from 30° banked turn, ailerons neutral, flaps and gear retracted. Test weight = 2670 lb; c.g. = 0.2842~

88 0 5 10 15 20 25 30 35 40 45 50 55 60 65 40 40 30 30 g 20 20 F D 73 10 10 ci d 00 00 3 -10 -10 :-20 -20 : -30 c -30 -40 I IIIll I It -40 Turns 1 8 30 n

-30 E 3 -30

200 150 n 100 - 50 E A A- Lr" ww OF -50 CJ -1 00 2 -1 50 -200

200 175 150 125 100 75 g 50 $ 25 0 0; -25 2 -50 5 Q -75 2-

-1-look 25 -1 50 L 3-150 -1 75 4 -1 75 m 1.Ok j 1.0 m

2 -.5 -.5 I- Q _J -;.o -1 .o -I

0 5 10 15 20 25 50 35 40 45 50 55 60 65 TIME. sec Figure 23. Continued. 0 5 10 15 20 25 30 35 40 45 50 55 80 65

3000 E

loo0 1000 0 750 750 e

JOt 7 30

rY ;:i -25 -25 LL C I- Q -50 -50 Q J _J

ol 1.oc 4 1.0 ol

0 5 10 15 20 25 30 35 40 45 50 55 60 65 TIME, sec Figure 23. Continued.

90 0 5 10 15 20 25 30 35 40 45 50 55 60 65

14000 14000 13000 13Ooo 12000 12000 11000 11000 I0000 1 0000 9000 9000 < 8000 -w' 7000 7m I-2 6000 6000 F J 5000 5000 4: 4000 4000 3Ooo 3Ooo 2000 2000 ?ooo 1000 0 0

250 4 250 200 200 0 0 # 150 150 $ 2 100 100 > 0) 0) 050 50 -u p- 0 0e 2 -50 -50 2

2 -150 -200

4E- -3

-1 €

0 5 10 15 20 25 30 35 40 45 50 55 60 65 TIME, sec

Figure 23. Concluded.

91 0 5 10 15 20 25 30 35 40 45 50

6o m m4 40cj 300 20 $ a 10J 4 0 Turns 1 3 15 m 104

5 CK- 0 Ol- 4 -5 2 rn -102 -15 v,

125 100 75 D 50 - W‘ 25 u lx O2 -25 -50 Fj -75 -100

60 $ 40.2u) 0) 20 u- W Po Ol- 4 4 o! -20 -20

-1

OF TIME. sec

Figure 24. Left spin of inodified airplane at maximum power entered from zoom maneuver, ailerons against the spin, flaps and gear retracted. Test weight = 2677 lb; c.g. = 0.2823~

92 0 5 101520253035404550 ::j ::j E m Q 20 20 -0 U 10 10 cj - d 0 0 00 3 -10 -103 & -20 -20 : -50 -30 -40 I II -40 Turns 1 3

9 100 - 50g oE -50 n -1002 -150k 4-150

200 200 175 175 150 150 125 125 100 100 75 $ 75 550 50 5 25 0 25 9 + -- h-- - F0 or 2 -25 -25 2 -50 -50 3 z 4: >- -75 -75 >. -100 -100

m 1.Ob 4 1.0 m 0- .5 .5 0-

/5 h__ 20 Y 02 -.5 I- -.5 4: 2-1 -1.0 -1.o-J

0 5 10 15 20 25 30 35 40 45 50 TIME. sec Figure 24. Continued.

93 0 5 10 15 20 25 30 35 40 45 50

? 2500 L

Turns 1 3 sot

n n - 50k 3 50-

-25 I- I- -25, QIL -50-25, -50 Q -I -I 200 200 Q 150 150 f 100 100 > 0 050 50%

--z Ow' I-do I- 2 -50 -50 2 -I -100 -100 j -I g -150 -150 g -200 -200

1.0 0' .5 0 40 g -.5 4 -1.0

0 5 10 15 20 25 30 35 40 45 50 TIME, sec

Figure 24. Continued.

94 c-3- 0 5 10 15 20 25 30 35 40 45 50

300 :250 200 %- 150 d w 100 n v, 50 0 I II Turns 1 3 11ooo 14000 13000 13000 12000 1 2000 11OOo 11000 loo00 loo00 g 9000 go00 w'- -w' 7000 n3 7000 5 I- I- 6000 F 5 6000 1 Q: 5000 ma 1ooo 4OOo 3ooo 3Ooo 2000 2000 loo0 1000 0 0

al -050 50 -0

-25oF 4 -250

5 5

01 4 401 cj 3 36 0 a 2 22 v, w 1 K 0 OW -1 -1

0 5 10 15 20 25 30 35 40 45 50 TIME. sec

Figure 24. Concluded.

95 0 5 10 15 20 25 30 35 40 45 50 55 60 65 80 $50 cia 030 $ 20 -I 10 4 0 Ill I1 Turns 1 4 15 15 cn 4 10 10 3 i5 5i 0 h_ r- Ol- ?o Q: -5 2 2 -5 m 2 -10 8 -10 2 -15 -15

125 100 75 50 e w- 25 o CY O? -25 " -50 zj -75 -100 -125

100 100 80 80 $ 60 60 $ $40 40: 4. 20 20 u-9) W W I-0 7 Oh 4 Q: CY -20 -20 CY

-60 -80 -1 00

5 5 4 4m

I;$ I;$ 01 -1 ,--2-1 0 5 10 15 20 25 30 35 40 45 50 55 60 65 TIME. sec

Figlire 25. Left spiii of inodified airplane at niaxirnum power entered from zooin maneuver, ailerons against tlic spiii, flaps and gear retracted. Test weight = 2678 Ib; c.g. = 0.2594C. 0 5 10 15 20 25 30 35 40 45 50 55 60 65 40 50 20 p U 10 cj 00 -10 5 -20 3 -50 It I11 -40 Turns 1 4 50 50 p 20 20 p * 10 1073 50 0% n n -10 -100 3 3 IY -20 -20 ctf -30 -30

200 200 150 150 n - 100 100 f! w-050 503 80LL 0; -50 -50 n n 2 -100 -1002 -1 50 -1 50 -200 -200

0 1.0 1.0 0 o- .5 .5 "' Yo - 02 t- -.5 -.5 I- Q Q -I -1.0 -1.0 -..I

0 5 10 15 20 25 50 35 40 45 50 55 60 65 TIME. sec Figure 25. Continued.

97 moo€ 2500 f! 2000 dw 1500 E 1000 so08 11 I I) 0: 0: Turns 1 4

t$ 20 20 t$ U 10 lou go -_ * - 0g tY (y: lj-10 -low i a -20 I- - -20 4 -30 -30

9- 50 50, 8 25 25 8 80 08 -25 I- -25 Q -50 -50 -I -I2

0 li I- -50 2 J -100 -100 1 -I -150 -150 -200 -200

035 TIME, sec

Figure 25. Continued.

98 14000 14000 15ooo 13000 12000 12000 11000 11000 loo00 ---- 1 0000 < 9000 9000 < w* 8000 8000 w- 7000 7000 2 2F 6000 8000 E J 2 5000 5000 4 4000 4000 3000 3000 2000 2000 1000 lo00 OF 40

0 150 t 100 50 50 w

z -100 3 k- -150 -150 2 -200 -200 -250 -250

5 :: :: 4m

-1 F -I

0 5 10 15 20 25 50 35 40 45 50 55 60 65 TIME, sec

Figure 25. Concluded.

99 125 100 75 9 50 J 25 o IY O2 -25 " -50 3 -75 -1 00 -1 25

cc -2oE 4 -20 l2

-60 E -80 1I," 5t j5 4E 4 40

-1 F 4 -1

0 5 10 15 20 25 30 35 40 45 50 55 60 65 TiME. sec

Figiirc 26. Left spin of niodified airplane at maximum power entered from zoom maneuver, ailerons against the spiii, flaps and gear retracted. Test weight = 2678 lb; c.g. = 0.2397C. 40 50 20 g -0 10 d 00 -102 -20 $j -50 Turns 1 3 50 N 50 010) 20 -- 20 g -0 10 lorn go 0g n n 0 -10 -100 3 3 fY -20 -20 fY -50 -33

200 200 175 175 150 150 125. 125 100 100

0 I 75 75 I0 2 50 => 25 0) 9 25 Po __ -- 0; 2 -25 -25 2

$ -50 -50 2 > -75 -75 2- -100 -1 00 -1 25 -1 25 -1 50 -1 50 -1 75 -1 75 -200 -200

m 1.0 1.0 0 "' .5 .5 "' $0 02 % -.5 -.s 4I- -1 -1.0 -1.o-l

TIME, sec Figure 26. Continued.

101 0 5 10 15 20 25 30 35 40 45 50 55 60 65

3000 E 2500 2

1500 "'"I1m 500 0 LT LK no,0°,,1 Turns 1 3 On 30 3 30

OF 0

750 750 0 t' 500 500 5- 2 250 250 2 OE

30 A30

200 150 : 100 -$ 0) 50-0 0 p- -50 2 -100 j g -150 -150 $ -200 -200

1.Ob 3 1.0

-.5 g 3 -1.0-.5~ -1.00,

0 5 10 15 20 25 30 35 40 45 50 55 60 65 TIME, sec

Figure 26. Continued. 1 4000 ~~ 14000 13000 13000 12000 12000 1 1000 1 1000 10000 10000 9000 9000 w- 8000 8000 w- 5 7000 7000 5 I- I- i= 8000 6000 F _J _J 4: 5000 5000 4: 4000 4000 3000 3000 2000 2000 1000 1000 OF 0

250 250 200 200 f 150 150 p > 100 100 2 a, '0 50 50 rln- n go - --- c UU' u- O@ 2 -50 -50 2. z -100 -100 z E IY 2 -150 -150 2 i- I- -200 -200 -250

-1 k 4 -1

0 5 10 15 20 25 30 35 40 45 50 55 60 65 TIME, sec

Figure 26. Concluded.

103 Ai rpl ane Time, sec , 1 configuration Baseline wing 5c 0 Modified wing 4.4 8.8 29.0

VI c, a 40 W c, m S .r a VI 4- 0 30 L nW 5

-m c, 20 u0 + 0 c, C WU

aa, 10

2 4 6 a 10 12 14 Time for 1st turn, sec Figure 27. Distributions of times for first turn of rotation following pro-spin control input.

104

- Stall Margin

Outer Wing Maxi mum Attainable Angle Of Attack Dynamic Effects

Center of Gravity

Elevator Trim

Figure 28. Maximum attainable angle of attack and stall-margin concept.

105 0 Idle power 0 50% power 0 Maximum power

60 -

50 -

40 -

(5, 0 -0 30 - " a

20 -

10 -

0. I I I I I 20 24 28 32 36 40

Center-of-gravi ty position, percent C (a) Flaps retracted. Figure 29. Variation of maximum trim angle of attack with center-of-gravity position for full trailing-edge-up stabilator deflection.

106 0 Idle power 0 50% power 0 Maximum power

Center-of-gravity position, percent C

(b) Flaps deflected 40'.

Figure 29. Concluded.

107 Idle power -n-Maximum power

Open symbols exclude zooms and vertical maneuvers

Darkened symbols are for zooms and vertical maneuvers 60

50 a Maximum attainable 0 m-= anqle of attack ,FA with dynamic effects

40 Outboard-wing-panel stall angle of attack

0, a, Maximum trim 30 angle of attack + tl

Basic wing stall 20 angle of attack

0 20 24 28 32 36 40

Center-of-gravity position, percent C (a) Flaps retracted. Figure 30. .Variation of maximum attainable outer-wing angle of attack with center-of-gravity position.

108 Idle power -a - Maximum power

50 Maximum attainable angle of attack with dynamic effects

40 Outboard-wing-panel stall angle of attack m Q, -0 Maximum trim 30 1 angle of attack I- b

Basic wing stall angle of attack

10 -

0 I I I I I

Center-of-gravity position, percent C (b) Flaps deflected 40'.

Figure 30. Concluded.

109 1.2 0 Basic wing

0 Modified wing 1.a

CL .6

0 I I 1 I I I I

a, deg

Figiirc 31. Airplaiic lift cliaracteristics with and without wing-leading-edge modification.

110 1.2 0 Basic wing

0 Modified wing 1.0

c~ .6

.4 cruise range

I I I I I I .02 .04 .06 .08 .10 .12

Figure 32. Airplane drag characteristics with and without wing-leading-edge modification.

111 nv\snNaltonai ~eronaui~cs ana Report Documentation Page Space AdminiSlrat~On

I. Report No. 2. Government Accession No. 3. Recipient’s Catalog No. NASA TP-2691 1. Title and Subtitle 5. Report Date Flight Investigation of the Effects of an Outboard July 1987 Wing-Leading-Edge Modification on Stall/Spin Characteristics 6. Performing Organization Code of a Low- Wing, Single-Engine, T-Tail Light Airplane 7. Author(s) 8. Performing Organization Report No. H. Paul Stough 111, Daniel J. DiCarlo, and James M. Patton, Jr. L-16243 10. Work Unit No. 9. Performing Organization Name and Address NASA Langley Research Center 505-61-41-01 Hanipton, VA 23665-5225 11. Contract or Grant No.

13. Type of Report and Period Covered 12. Sponsoring Agency Name and Address National Aeronautics and Space Administration Technical Paper Washington, DC 20546-0001 14. Sponsoring Agency Code

15. Supplementary Notes

16. Abstract Flight tests were performed to investigate the change in stall/spin characteristics due to the addition of an outboard wing-leading-edge modification to a four-place, low-wing, single-engine, T-tail, general aviation research airplane. Stalls and attempted spins were performed for various weights, center- of-gravity positions, power settings, flap deflections, and landing-gear positions. Both stall behavior and spin resistance were improved compared with the baseline airplane. The baseline airplane would readily spin for all combinations of power settings, flap deflections, and aileron inputs, but the modified airplane did not spin at idle power or with flaps extended. With maximum power and flaps retracted, the modified airplane did enter spins with abused loadings or for certain combinations of maneuver and control input. The modified airplane tended to spin at a higher angle of attack than the baseline airplane.

17. Key Words (Suggested by Authors(s)) 18. Distribution Statement Stall Unclassified -Unlimited Spin Wing design Light airplane General aviation

19. Security Classif.(of this report) 120. Security Classif.loj his page) 21. No. of Pages 22. Price Unclassified I Unclassified 116 A06 NASA FORM 1626 OCT 86 NASA-Langley, 1987 For sale by the National Technical Information Service, Springfield, Virginia 22161-2171