Wind-Tunnel Investigation of the Flight Chasacteristics of a Canard Genera1:Aviation Airplane Coniiguration

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,Wind-TunnelInvestigation of the Flight Chasacteristics of a Canard Genera1:Aviation Airplane Coniiguration

I Dale R. Satran

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NASA NASA Technical Paper 2623

1986 Wind-Tunnel Investigation of the Flight Characteristics of a Canard General-Aviation Airplane Configuration

Dale R. Satran Langley Research Center Hampton, Virginia

National Aeronautics and Space Administration Scientific and Technical Information Branch Summary figuration. A 0.36-scale model was used in the free- flight investigation and was also used to obtain static A 0.36-scale model of a canard general-aviation and dynamic force data to aid in the interpretation airplane with a single pusher propeller and winglets of the free-flight test results. The free-flight tests was tested in the Langley 30- by 60-Foot Wind Tun- were conducted for angles of attack ranging from 7O nel to determine the static and dynamic stability and to 14O. The investigation included tests of the model control and free-flight behavior of the configuration. with high and low canard positions, three center-of- Model variables made testing of the model possible gravity locations, outboard wing leading-edge droop, with the canard in high and low positions, with in- winglets and a center vertical tail, and several roll- creased winglet area, with outboard wing leading- and yaw-control systems. Dynamic force tests were edge droop, with fuselage-mounted vertical fin and also conducted on the 0.36-scale model using the rudder, with enlarged rudders, with dual deflecting forced-oscillation test technique to study the effects rudders, and with ailerons mounted closer to the of two canard vertical positions and using the instal- wing tips. lation of winglets and outboard leading-edge droop The basic model exhibited generally good longitu- on the roll damping of the model. Wool tufts were dinal and lateral stability and control characteristics. installed to aid in flow visualization of the stall pat- The removal of an outboard leading-edge droop de- tern of the wing and canard during the static force graded roll damping and produced lightly damped tests and free-flight tests. roll (wing rock) oscillations. In general, the model exhibited very stable dihedral effect but weak direc- Symbols tional stability. Rudder and aileron control power were sufficiently adequate for control of most flight All longitudinal forces and moments are refer- conditions, but appeared to be relatively weak for enced to the stability-axis system, and all lateral- maneuvering compared with those of more conven- directional forces and moments are referenced to the tionally configured models. body-axis system. The midpoint of the center-of- gravity range is 0.71C ahead of the leading edge of Introduction the wing mean aerodynamic chord. (See fig. 1.) The wing reference area corresponds to that area obtained As part of the NASA general-aviation stall/spin by extending the outboard leading and trailing edges program, advanced aircraft configurations are be- of the wing without leading-edge droop to the fuse- ing investigated that offer unique safety benefits. lage centerline. All dimensional quantities are ex- One such configuration, the R.utan VariEze, uti- pressed in both the International System of Units lizes a high-aspect-ratio canard, a swept-back wing, (SI) and U.S. Customary Units. Measurements were winglets, and a pusher propeller. Full-scale flight made in U.S. Customary Units, and conversion fac- tests of the homebuilt canard aircraft have demon- tors from reference 5 were used to obtain equivalent strated advantages for such a design from the stand- SI dimensions. point of increased stall departure and spin resistance. (See ref. 1.) Several models of this configuration were wing span, cm (in.) tested in different facilities to document the flight drag coefficient, characteristics of the VariEze. Reference 2 contains 9 static wind-tunnel data for a full-scale mode! of the iift coefficient, configuration tested in the Langley 30- by 60-Foot 3 Wind Tunnel. Rotary-balance tests were conducted rolling-moment coefficient, on a 0.22-scale model in the Langley Spin Tunnel, and the results showed the configuration to have in- pitching-moment coefficient, herently good stall departure and spin resistance. Pitching moment qsc (See ref. 3.) In addition to improved safety fea- tures, the configuration has all-composite construc- yawing-moment coefficient, tion, which makes possible a smooth surface finish and, for this particular configuration, the realization effective thrust coefficient at zero angle of attack, Drag (power off)-Drag (power on) of performance gains through large improvements in 9s natural laminar flow. (See ref. 4.) The purpose of this investigation was to use side-force coefficient, Sid:p the free-flight test technique in the Langley 30- by mean aerodynamic chord, cm (in.) 60-Foot Wind Tunnel to study the dynamic stability and control and general flight behavior of the con- frequency of oscillation, Hz moment of inertia about X axis, kg-m2 Abbreviations: (slug-ft2) BL butt line moment of inertia about Y axis, kg-m2 c.g. center of gravity (slug-ft2) FS fuselage station moment of inertia about 2 axis, kg-m2 L.E. leading edge (slug- ft2 ) max maximum incidence angle of canard, positive WL waterline trailing edge down, deg reduced frequency parameter, wb/2V Model and Apparatus roll rate, rad/sec The basic configuration is depicted in a three-view diagram in figure 1, and photographs of the model free-stream dynamic pressure, Pa (psf) are shown in figure 2. The mass and dimensional characteristics are included in table I. The 0.36-scale wing reference area, m2 (ft2) model is representative of the Rutan VariEze, a two- free-stream velocity, m/sec (ft/sec) place, advanced general-aviation airplane. For all tests, the nose gear was retracted. The mass and in- spanwise coordinate, m (ft) ertial characteristics were scaled to correspond to op- angle of attack, deg eration at 1524 m (5000 ft) altitude (standard atmo- sphere). The wing, winglet, and canard of the model angle of sideslip, deg were constructed of balsa wood and fiberglass. The rate of change of sideslip, rad/sec fuselage was made of fiberglass and foam sandwich construction with an internal aluminum structure. increniental rolling-moment coefficient The control surfaces were actuated for free-flight (control deflected - control neutral) tests by electroprieurrlatic servos. The controls con- iricrernental yawing-moment coefficient sisted of a slotted canard flap used as an elevator, ailerons located inboard on the main wing, and rud- (control deflected - control ricutral) ders mounted on the winglets. The basic rudders increinrnt a1 side-force coefficient deflected independently and outward only; that is, (control deflected - control neutral) for a left turn, the trailing edge of the left rudder only would move to the left. The control deflections aileron deflection, positive for left roll, were limited during free-flight tests to f20" for the deg ailerons, f30" for the rudders, and f5" for the el- elevator deflection, positive for trailing evator. Thrust to fly the model was supplied by a edge dow11, deg propeller driven by a turbine-air motor using compressed air. flap deflection, positive for trailing edge Static force tests were made with several rud- down, deg der modifications using sheet-metal tabs to simulate dual, split, and enlarged winglet rudders. These rud- rudder deflection, positive for left rudder der modifications are shown in figure 3. The dual and trailing edge left, deg split rudders had an area equivalent to that of the angular frequency, 27r f , rad/sec basic rudders. For dual-rudder control, both rudders deflected simultaneously and in the same direction; Stability derivatives: that is, for a left turn, the trailing edges of both rudders moved to the left. For split-rudder control, the inboard arid outboard surfaces of one rudder split and deflected outward from the neutral position, and the rudder oti the opposite winglet remained undeflected. The enlarged rudders operated in the same manner as the basic rudders, but had twice the chord, and extended in height to the tip of the winglets. The hinge line of the rudder was unchanged for all rudder modifications. A center vertical fin and conventional

2 rudder mounted on the fuselage directly ahead of the The static force data were measured at a nominal propeller were also tested. dynamic pressure of 464 Pa (9.7 psf), corresponding Force tests were also conducted with outboard- to a Reynolds number of 0.535 x lo6 based on wing mounted ailerons and with differential elevator de- mean aerodynamic chord or 0.218 x lo6 based on ca- flection for roll control. The outboard ailerons were nard mean aerodynamic chord. The static force tests simulated using sheet-metal tabs mounted on the were conducted over an angle-of-attack range from trailing edge of the wing on the outer 25 percent of -loo to 90' and an angle-of-sideslip range of f15', the span. (See fig. 4.) Landing flaps were simulated although some of the tests were made over reduced in exploratory force tests by deflecting the ailerons angle-of-attack ranges. Static sideslip derivatives symmetrically. were determined from 3x5' sideslip angles. Wind- Tests were conducted with the canard mounted on tunnel flow-angularity corrections were applied to all the top of the fuselage, as the basic location, and with data based on wind-tunnel flow surveys. Because the the canard mounted on the bottom of the fuselage, size of the model was small relative to that of the test as an alternative location. Canard incidence was set section, no jet boundary corrections were included. at 0' and 55' in both positions. As a result of preliminary wind-tunnel testing Forced-Oscillation Tests and guidelines presented in reference 6, no. 60 grit was applied along the midchord of the upper sur- Dynamic force tests were conducted using the face of the canard to avoid laminar flow separation forced-oscillation test equipment diagramed in fig- at the low test Reynolds numbers. Also incorporated ure 6. The model was mounted on a strut that forced into the basic configuration were wing leading-edge the model to oscillate sinusoidally about the roll axis droop modifications on the outer 25 percent of the while an internal strain-gauge balance measured the span. These droop modifications resulted in a wing forces and moments on the model. Reduction of the chord extension of about 6 percent and a camber in- data provided measurements of the oscillatory stabil- crease of 3 percent. The leading-edge droop mod- ity derivatives that contain both damping and linear ifications were similar to those in reference 7. A acceleration components. It has not been possible to diagram of the airfoil-section modification and place- accurately separate the two contributions, but expe- ment of the leading-edge droop is included in fig- rience has shown that reasonably good accuracy in ure 4. Several tests were conducted with vortex gen- dynamic stability calculations can be obtained using The dynamic force test tech- erators located at the wing midchord, between the the combined form. outboard end of the ailerons and the inboard end nique is discussed more fully in reference 8. The of the leading-edge droop, to investigate their effect data reduction scheme is presented in the appendix of on the stall characteristics of the wing. The vortex reference 9. generators were sheet-metal tabs (1.27 cm (0.5 in.) The model was tested in roll on the forced- square) angled f45' to the free stream. (See fig. 4.) oscillation apparatus at a dynamic pressure of 440 Pa During the investigation, the model produced an (9.2 psf). A value of k = 0.12 was selected as rep- asymmetric stall with the right wing stalling at a resentative of full-scale flight, which corresponds to lower angle of attack. Templates of the wing airfoil a model oscillation frequency of 0.40 Hz at this dy- were made at several stations, and these templates namic pressure. The amplitude of the roll oscillations showed that a discrepancy in the leading-edge radius was f5'; however, an amplitude of flO' was also had been built into the right wing panel. The tested with no significant differences ir? the results. contours are shown in figure 5. The effect of this . The model was tested with the canard in the high model construction error is discussed in the section position with 0' incidence and in the low position "Static Lateral-Directional Stability." with 5' incidence to correspond to the free-flight test configurations. Data were also taken with the canard Testing Techniques off. With the canard in the high position, tests were made with the winglets installed and removed and Static Force Tests with the leading-edge droop installed and removed. Static force tests were conducted in the Langley Free-Flight Tests 30- by 6O-Foot Wind Tunnel using a six-component strain-gauge balance mounted internally at the mid- In the free-flight test technique, two pilots fly point of the center-of-gravity range. Also, the canard the model within the open-throat test section of the was isolated from the aircraft by a second internal Langley 30- by 6O-Foot Wind Tunnel. Figure 7 is a balance, so that simultaneous canard loads could be diagram of the setup. A flexible flight cable supplies measured independently. compressed air for power, transmits control signals,

3 and acts as a safety cable for the model. The flight handling qualities of the model. In addition to the cable is kept slack by a safety-cable operator using a qualitative observation of model motions, strip charts high-speed, pneumatic winch. The roll and yaw pi- are used to record control input and model rates and lot is located behind and below the test section; the accelerations. pitch pilot and the throttle and safety-cable opera- tors are located beside the test section. Rate gyros, The free-flight tests were conducted over a range accelerometers, and control-position potentiometers of dynamic pressures from 402 to 263 Pa (8.4 to are mounted in the model, and the output of this 5.5 psf) to investigate the dynamic response and instrumentation is sent to a flight-control computer. flight characteristics for nominal angles of attack The flight-control computer receives the control in- from 7" to 14". These dynamic pressures correspond puts from the pilots and the information from the to a Reynolds number range from 0.498 x lo6 to model instrumentation and combines them according 0.403 x IO6, based on wing mean aerodynamic chord. to preprogrammed control laws that allow incorpo- The model was tested with the canard in the high ration of automatic control mixing and artificial sta- position at 0" incidence and with the wing leading- bilization. A more complete discussion of this tech- edge droop installed and removed. Also, tests were nique is contained in reference 8. For the present conducted with the canard in the low position at 0' general-aviation study, rate gyros were installed in and 5" incidence, but only with the wing leading-edge the model to provide a stable platform for conve- droop installed. Ballast was adjusted so that three nience in exploratory studies, but all flights were re- center-of-gravity locations were tested: the basic peated with gyros turned off. location of 0.71C ahead of the leading edge of the The free-flight test results are in the form of mean aerodynamic chord of the wing, and locations pilot observations and motion-picture records of the 0.10C ahead of and behind the basic location.

Presentation of Results The test results are presented in figures 8 to 35, which are grouped in order of discussion as follows: Figure Static force tests: Static longitudinal stability and control: Effect of fixed transition on canard ...... 8 Lift and pitching-moment characteristics: Effect of canard ...... 9 Elevator control deflections (high position) ...... 10 Canard position ...... 11 Elevator control deflections (low position) ...... I2 Canard incidence ...... 13 Power effects ...... 14 Effect of the leading-edge droop: Flow visualization with tufts ...... 15 Longitudinal characteristics ...... 16 Effect of landing flaps ...... 17 Static lateral-directional stability: 1,nteral-directiori;tl characteristics: Effect of sideslip ...... 18 Effect of ci~~lardand leading-cdgc droop ...... 19 Effect of vortex geiicrtit ors and extended leading-edge droop ...... 20 I, at cral-di rec t io1 1a1 static stat) i 1 i ty : Effect of wiriglcts ...... 21

4 Effect of leading-edge droop ...... 22 Effect of canard position ...... 23 Effect of center vertical tail and enlarged winglets ...... 24 Static lateral-directional control: Aileron effectiveness ...... 25 Effect of outboard ailerons ...... 26 Effect of differential elevator ...... 27 Rudder effectiveness ...... 28 Effect of center rudder ...... 29 Effect of modified winglet rudders ...... 30 Forced-oscillation tests: Dynamic roll stability: Effect of leading-edge droop ...... 31 Effect of canard ...... 32 Effect of winglets ...... 33 Interpretation of results: Calculated aileron response ...... 34 Calculated rudder response ...... 35

Results of Static Force Tests range should be representative of stability and control values. Results presented in reference 2 are for The static force data were measured to aid in the Reynolds numbers near flight conditions for the full- analysis and interpretation of the free-flight test re- scale airplane and better predict the performance of sults and are therefore discussed prior to the free- the full-scale airplane. flight test results. The basic configuration was con- Canard configurations require that the center of sidered to have the canard in the high position and gravity be located between the canard center of outboard leading-edge droop on the wing. The ca- lift and wing center of lift for positive stability and nard incidence was set at O', and the data were pre- positive control. If the canard stalls before the sented for the mid c.g. location of 0.71C ahead of the wing stalls, longitudinal stability and airplane stall leading edge of the mean aerodynamic chord unless otherwise noted. resistance are increased. The lift characteristics of the canard (fig. 9) indicate that the canard achieved maximum lift at an angle of attack slightly lower than Static Longitudinal Stability and Control that for model maximum lift. However, many factors The results of fluorescent oil-flow visualization must be considered in order to make a configuration tests indicated that substantial laminar boundary- stable and controllable as well as stall resistant. layer separation occurred on the canard. This sepa- The lift and pitching-moment data for the basic ration resulted in the highly nonlinear lift curve and configuration (fig. 10) indicate that maximum ele- high drag characteristics shown by the canard bal- vator deflection for the mid c.g. position provided ance data in figure 8. Grit was applied to the mid- pitch trim for the model to a = 15', which is below chord of the upper surface of the canard to min- wing stall. The canard is ineffective for trimming the imize the flow separation, and all data presented model to an angle of attack beyond wing stall at the herein were collected with grit applied to the canard. design mid c.g. location; therefore, the canard pro- It should be noted that airfoils suffer performance vides inherent angle-of-attack limiting at the mid c.g. degradation at low Reynolds numbers. Because of position. For the forward c.g. position, the maximum this degradation and the variance in Reynolds num- elevator deflection provided pitch trim to cr = ll', ber between the full-scale airplane in flight and the which still provides a CL value of 1.2. The maximum 0.36-scale model in the wind tunnel, the results from elevator deflection for the aft c.g. position provided the model do not reflect the performance of the full- pitch trim to a = 30' with a potential deep-stall scale airplane. However, results in the linear lift pitch-trim condition at a. = 53'. Figure 1O(c) shows

5 that the model has sufficient elevator effectiveness apparently because the thrust line passed near the at cy = 53" to pitch the model out of the deep-stall model center of gravity. At high angles of attack, the trim point. As pointed out previously, the Reynolds data show that thrust increased the lift and provided number for this investigation was significantly lower a stabilizing diving moment. This diving moment re- than the Reynolds number for full-scale flight condi- sulted partly from the induced slipstream effect over tions. The data presented in reference 2 showed no the wing, which increased the wing lift; however, the deep-stall point. diving moment is probably caused mainly by the de- The aerodynamic data for the low canard position velopment of a propeller normal force by the rotating are compared with data for the high canard position propeller disk, which on a pusher configuration pro- in figure 1l(a), and the results indicate generally sim- duces a stabilizing moment. ilar lift and pitching-moment characteristics for the Photographs showing the results of tuft studies two canard positions. The low canard position re- on the model with and without the outboard wing sulted in a small negative shift in angle of attack for leading-edge droop are presented in figure 15. These C,,, = 0.0. The drag characteristics presented in fig- tuft photographs indicate that the installation of urc ll(b) do not change significantly with a change in leading-edge droop significantly reduced the region the canard position, although drag measurements at of separated flow at the wing tips at high angles of this low Reynolds number are not directly applicable attack. Data presented in figure 16(a) indicate un- to full-scale flight. stable pitching-moment characteristics in the angle- The elevator effectiveness with the canard of-attack range from about a = 5" to a = 15" for the mounted in the low position is presented in figure 12 model without the leading-edge droop. Installation and indicates inherent angle-of-attack limiting, sim- of the leading-edge droop eliminated the unstable ilar to that for the canard mounted in the high po- trends in the pitching-moment curve up to cy = 30'. sition. The low canard position resulted in some The drag characteristics presented in figure 16(b) reduced static stability with the elevator deflected. show no significant drag penalty for the installation This reduction was indicated by the reduced slope of of the leading-edge droop at positive lift coefficients tlic pitching-moment curve at approximately a = (3". at the test Reynolds number. The higher Reynolds This loss in stability with elevator deflection for the number data of reference 2 indicated a cruise drag low canard position was not as apparent for the high penalty of 0.0040 for the leading-edge droop. cuiard position. (See fig. 9.) This loss in pitch stabil- Landing flaps were simulated by deflecting both ity may be associated with canard-wing interference, ailerons symmetrically. Figure 17 indicates that which occiirred as a down load on the inboard wing flap deflections of 22' increased CL,~~~by 0.1 and and an upwash outboard on the wing. This adverse decreased the angle at zero lift by 2'. However, flow behavior of the canard on the wing was appar- flap deflection increased the nose-down pitching mo- ently more pronounced for the low canard configura- ment such that 20' elevator deflection trimmed to tion than for the high canard configuration. only CL = 1.0, which is lower than the niaxiniuni The results of tests to show the effects of adjust- trinimablc CL without flaps. ment of canard incidence on the aerodynamic characteristics of the high and low canard configurations are Static Lateral-Directional Stability presented in figure 13. The data show that increasing canard incidence to 5" increased the static stability of The static lateral-directional characteristics are the configuration at higher angles of attack by caus- presented as a function of sideslip angle in figure 18 ing the canard to stall in the region where partial and are generally linear with sideslip angle up to wing stall wonld cause a reduction of static stability. a = 18". The slopes of the data indicate a trend The data of figure 13(b) show effects of incidence toward decreasing directional stability C,, and in- change for the low canard position similar to those creasing the dihedral effect C18 with increasing an- for the high canard position. In choosing the proper gle of attack up to a = 14'. The static lateral- caitard incidence angle, consideration must be given directional coefficients at ,8 = 0" are presented in to cruise drag penalties and stability characteristics. figure 19 and indicate that an asymmetry in roll and Tlici low Reynolds nimibers of the sutiject tests do yaw to the right occurred through an angle-of-attack riot providc data suitable for such trade-off studies. range froin 15' to 25' that was not affected by the Resiilts of tests to drterinirie the effects of pro- canard or leading-edge droop. Tuft studies indicated pcllrr thrust on the aerodynamic characteristics of that the right wing stalled more abruptly than the tlic niodcl arc presented in figure 14. Figure 14 also left, and further investigation deterinined that a dis- shows that there were no significant thrust effects on crepancy in the leading-edge radius of the right wing the lift and pitching moment at low angles of attack, of the model caused the asymmetric stall to occur.

6 The asymmetric stall was not a characteristic of the the angle of attack at which Cna became unstable. configuration, but was a result of inaccurate model Neither the center tail nor enlarged winglets affected construction on the right wing panel. As shown in Clo significantly up to the stall angle of attack. figure 20, the asymmetric stall could be nearly eliminated by placing vortex generators at the midchord Static Lateral-Directional Control of the wing or by extending the leading-edge droop inboard to the region of the discrepancy. The results of tests to determine the lateral con- The effects of the winglets on the static lateral- trol power are presented in terms of ACy, ACn, and directional stability derivatives are presented in fig- AC, produced by a right-roll or right-yaw control ure 21. The basic configuration exhibited stable dihe- input. Data showing aileron effectiveness are pre- dral effect, due to wing sweep, that increased linearly sented in figure 25 and indicate that the available with angle of attack up to Q = 14'. The winglets pro- roll control decreased with increasing angle of attack, vided a constant increment to Clo. Above Q = 14', probably because of the progression of separated flow Cia was reduced sharply due to the asymmetric stall into the region of the ailerons. Aileron deflection ex- noted previously. Normally, Clo would be expected hibited favorable yawing moments from cy = 0' to Q = 14', above which the yawing moment became to increase linearly with angle of attack in the re- adverse. The canard vertical position produced only gion where is a linear function of angle of attack lift minor effects on the aileron control data. (Q < 16'). The directional stability of the basic con- The roll-control characteristics of simulated aile- figuration decreased linearly from moderate stability rons mounted outboard, behind the leading-edge at Q = 0' to neutral stability at Q = 14'. Above droop, are compared with data for the basic ailerons Q = 14', Cnp increased sharply to cy = 20' and in figure 26. As expected from the increased lateral- then decreased and became unstable at high angles moment arm, the outboard ailerons greatly increased of attack. The increment to Cap from the winglets the available roll control up to Q = 25'. However, decreased with increasing angle of attack. the outboard ailerons caused moderate values of ad- The effects of wing leading-edge droop on the verse yaw below Q = 8' but eliminated the adverse lateral-directional characteristics of the model are yaw characteristics of the basic ailerons in the stall shown in figure 22. The leading-edge droop made region. no significant contribution to the static lateral- Differential elevator deflection to provide roll con- directional stability at low-to-moderate angles of at- trol is compared with the basic aileron effectiveness tack. In the post-stall region, the droop provided flow in figure 27(a); canard balance data are included for attachment at the wing tips which improved Clo, al- comparison with model balance data. Figure 27(a) though Cna was generally degraded. shows that differential elevator control was less effec- The effect of the canard on lateral-directional sta- tive than the basic ailerons. Below cy = 16', the ca- bility is presented in figure 23(a), and the data indi- nard downwash altered the local angle of attack of the cate that the canard had no significant effect on Clo wing such that the wing produced an opposing rolling but degraded Cno below stall in either the high or moment. Above Q = 16', the wing was stalled, and the rolling moment measured by the canard balance low position. In the post-stall angle-of-attack region, was equal to that of the model balance. Differential the high canard was very destabilizing directionally elevator deflection for the low canard (fig. 27(b)) in- but provided a stabilizing increment to lateral stabil- dicates similar results to those of the high canard. ity. The data of figure 23(b) show that deflecting the The yawing moments produced by differential eleva- canard elevator produced little effect on the lateral- tor deflection were more adverse for the low canard directional stability characteristics. The increase in position than. for the high canard position. Ci, produced by the canard is probably due to the ca- Data showing rudder effectiveness are presented nard downwash being asymmetric on the wing when in figure 28 and indicate that the yawing moment the configuration is sideslipped. produced by rudder deflection was approximately lin- Increased winglet area and a center vertical tail ear with rudder deflection. The rudder effectiveness were tested to provide increased directional stability; remained almost constant with angle of attack up the results are presented in figure 24. The data of to Q = 10' and then decreased linearly through the figure 24 show that the center tail provided a con- stall. Comparison of the data of figures 18 and 28 stant increment of Cn, of about 0.005 that was not indicates sufficient rudder control up to fairly large affected by thrust changes, despite the close proxim- sideslip angles. Comparison of the winglet rud- ity of the propeller to the tail. The enlarged winglets der data to data measured for a rudder mounted provided a substantial increase in C,, and increased on a center vertical tail (fig. 29) indicates that the

7 center rudder produced approximately half as much Results of Free-Flight Tests yaw control as the winglet rudders. There were slight The majority of the free-flight tests were con- power effects on the center rudder for climb power ducted with augmented roll-, yaw-, and pitch-rate conditions (C> = 0.39). damping to provide a stable platform for ease of fly- Three modifications to the winglet rudders were ing while conducting exploratory studies. However, tested (as noted in the section "Model and Appara- flights performed without stability augmentation in- tus"), and the results are presented in figure 30. The dicated that there were no significant changes in the enlarged rudders and the dual-rudder deflections in- trends or overall flight characteristics other than a creased the available yaw control. The split rudders, reduction in the workload of the pilot that was re- however, provided a reduction in yaw control. Fur- quired to control the model. Where applicable, the ther testing indicated that the reason for the adverse results of the free-flight tests are compared with those control effect from the split rudder was that the drag of the static and dynamic force tests previously dis- from the split rudder, which was the mechanism ex- cussed. Since the free-flight tests were conducted at pected to yaw the model in the direction of the split lower Reynolds numbers than the static force tests, riidder, was offset by the reduction in inward side some characteristics, such as local flow separation, force iiornially produced by that winglet. The inward may occur at slightly different angles of attack. side force from the opposite winglet was greater than Late in the flight-test program, when high-angle- that produced from the split-rudder winglet. The net of-attack stall flight tests were attempted, an asym- effect was that the yaw control was opposite to that metry between the left and right wing panels was de- desired. tected. The asymmetry caused the model to depart in yaw to the right at an angle of attack of 14" for all Results of Forced-Oscillation Tests test conditions. No free-flight tests were conducted after fixes for the asymmetry were found. The roll forced-oscillation tests were conducted to nieasure the oscillatory stability derivatives Cy, + Longitudinal Flight Characteristics Cl. sin a, C,,, + C,, sin a, and Cl, + Cl sin a, and 1' B D the results of the tests are presented in figures 31 The basic configuration with the high canard was to 33. Roll damping Cl, + Cl sin a was of primary stable and well damped in pitch but was sensitive to 0 elevat,or control inputs, particularly for nose-down int crest 1)ccauscof wing rock oscillations encoiintered control. The model dropped and pitched rapidly in the free-flight testing. from trimmed flight with slight nose-down elevator The leading-edge droop was installed to prevent deflection. Elevator deflection required for trimmed preniature wing-tip stall that was found in static tuft flight varied linearly from 6, = 6" at a = 9" to studies. The data of figure 31 show a loss of roll 6, = 14' at 0 = 14' for the mid c.g. position. The duiiping without tlic Icatlirig-cdge droop and unsta- model had good pitch stability at the aft center- ble roll daniping at a = 16'. The addition of the of-gravity location, but the workload of the pilot leading-edge droop helped to nnaintain attached flow increased significantly because of increased elevator at the wing tips arid provided stable roll damping sensitivity. Any attempt to fly at higher angles of throughout the angle-of-attack range. attack, particularly at the aft c.g. position, was ex- The data of figure 32 show that the canard posi- pected to give less desirable pitch stability charac- tion had little effect on roll stability. The roll danip- teristics. The model was not flown above o = 14' ing was progressively reduced as angle of attack in- because of an asymmetry between the left and right creased, and the canard position did not affect this wing panels (fig. 5), which caused the model to de- trwd. The roll damping reniained stable through- part in yaw to the right at o = 14'. This departure out the angle-of-attack range with the canard in ei- was not characteristic of the airplane and was strictly ther position, although the darnping decreased in the a characteristic of the model. anglc-of-attack range from 15' to 20'. The removal of the leading-edge droop did not The effect of winglets on the roll oscillation data is significantly affect the longitudinal flight character- sliowti iii figure 33. data show that the winglcts Thc istics for the limited angle of attack investigated. The Iiatl littlc tffect 011 roll danipiiig at thc lower anglcs of static force data presented in figure 16 indicate that, ;ittil(.k. Thc niiiiii cffect of the winglcts wits to providc with the droop removed, a destabilizing break oc- i~ iiqyLtivc incrciiicnt to C,,, + C,, sill (k ovcr thc tcst M curred in the pitching-moment curve at angles of at- iiiigl(wf-;Ltt iick r:~ngc. For the low-~Lnglc-of-att;tck tack near stall. However, the model exhibited no rarigc., this iiegiit ive incrcnicnt to C,,, C,, sin (1 + P pitch instability with the leading-edge droop removed results in aiadverse yawing niornent due to rolling. at the lower angles of attack. Longitudinal power ef-

8 fects were not significant. As demonstrated by the angles of attack greater than ll', the rudder effec- static data presented in figure 14, the pitching mo- tiveness was not sufficient to recover the model from ment was not affected by thrust changes in the free- large disturbances without the addition of aileron de- flight test angle-of-attack range. flection. Recovery from the departure at cy = 14' due Flights made with the low canard position indi- to wing asymmetric stall could not be made with the cated that the longitudinal flight characteristics were available aileron and rudder control. The aft center- similar to those with the high canard. The model of-gravity position resulted in reduced control effec- displayed similar pitch stability and elevator sensi- tiveness and more erratic flight behavior. tivity. The static data presented in figure 11 indi- Removal of the leading-edge droop caused definite cate a slight nose-down shift in the pitching moment deterioration of flight characteristics in both roll and because of the low canard. This shift was demon- yaw. At cy = O', the model behaved much the strated by a 2' to 3' increase in the elevator deflec- same as with the droop installed, but with reduced tion required to trim the model with the low canard damping in roll. At cy = ll', wing rock could be throughout the angle-of-attack range tested. To re- easily induced by small aileron or rudder deflections. duce the trim elevator deflection with the low canard, The oscillation was stable but only slightly damped, flight tests were also conducted with 5' canard inci- which correlates with the forced-oscillation results of dence. The trim elevator deflection was reduced by figure 31 that indicate unstable roll damping with 3' to 4' with the increased canard incidence. Static the droop removed. data presented in figure 13 indicate an increase in The model flew steadier in pitch with the low ca- static stability and a slight reduction in elevator sen- nard at 5' incidence, which resulted in a reduced sitivity with 5' incidence. These results were verified workload for the lateral-directional pilot. No sig- in the flight tests. The flights with 5' canard inci- nificant difference was seen in the lateral-directional dence were much steadier, and the model did not flight characteristics for the low or high canard con- display the elevator sensitivity found with the high figurations. This result was expected, since the static canard or the low canard with 0' incidence. and dynamic force tests indicated that the low canard position did not alter the static or dynamic lateral- Lateral-Directional Flight Characteristics directional force characteristics. The basic configuration with the high canard The enlarged winglet rudders and center tail were and leading-edge droop installed exhibited adequate tested to increase the directional stability and rudder directional stability and stable dihedral effect, as effectiveness. The static data of figures 24, 29, shown in the static data of figure 21. The model and 30 show the improvements in directional stability motions were stable and well damped in roll and yaw and rudder control, and these improvements were under cy = 14'. At cy = 14', the model departed supported by the free-flight tests. Although the in roll and yaw to the right as the result of an model had improved flight characteristics below the asymmetric stall brought on by an airfoil asymmetry stall, the increased directional stability and rudder between the right and left wing panels. effectiveness were not sufficient to prevent departure The basic lateral-directional control power was at cy = 14' because of the wing contour asymmetry sufficient for adequate control of the model, although problem. the pilot noted that the rudder effectiveness was relatively weak for maneuvering and recovering from Interpretation of Results disturbances. The flight tests showed that the basic The results of the free-flight tests for this config- ailerons produced noticeable adverse yaw, particu- uration showed generally good agreement with the larly above a = 9'. The static force data in figure 25 static force test data. The asymmetry in the model indicate adverse yaw at angles of attack between 15' limited the angle-of-attack range in which testing and 20'. Adverse yaw is usually related to local flow could be performed but did not affect the results un- separation on the wing panel with the aileron de- less otherwise noted. Of course, the low Reynolds flected down. Since the free-flight tests were con- numbers associated with the present tests could cause ducted at lower Reynolds numbers than the static some characteristics, such as local flow separation, force tests, the adverse yaw was probably occurring to occur at slightly different angles of attack under at lower angles of attack. In all cases, the adverse different test conditions. Also, the confined space yaw could be compensated for by rudder deflection. available within the wind tunnel, the rapidity of the The rudder effectiveness was sufficient to control the motions of the model, and the lack of piloting cues model without aileron input at lower angles of at- cause the evaluation of the longitudinal and lateral- tack because of the large bank angles that could be directional control techniques to be qualitative at induced by the strong favorable dihedral effect. At best.

9 To better evaluate the free-flight test data, calcu- directional characteristics were not significantly af- lations were made using linear analysis techniques of fected by canard vertical position. three-degree-of-freedom lateral-directional equations 4. The addition of an outboard wing leading-edge of motion. Included in the computations are response droop prevented premature wing-tip stall and, as a estimates for a representative conventional low-wing result, increased the roll damping and eliminated a configuration (ref. 10) for comparison with the re- wing rock tendency at high angles of attack. sponse estimate of the advanced-canard configura- 5. Roll control from the basic inboard ailerons tion. The results of calculated response due to ai- was sufficient to control the model in free-flight tests. leron and rudder inputs are presented in figures 34 Outboard-mounted ailerons improved aileron effec- and 35. The data of figure 34 show that the aileron tiveness but increased adverse yaw. roll response for the basic canard configuration was 6. Sufficient rudder control was available for lower than that of the conventional configuration at steady flight conditions. The rudder effectiveness was low lift coefficients and about equal for the two con- increased by dual deflection of the basic rudders or figurations at high lift coefficients. Modifying the ca- by increased rudder area. nard design with ailerons located outboard provided 7. A three-degree-of-freedom lateral-directional more aileron response for the canard design at low equations-of-motion study indicated weak lateral- and high lift coefficients. The rudder response data directional control power, especially when the rud- of figure 35 show that the yaw response from rudder der effectiveness was compared with a conventional deflection was much lower for the canard design than low-wing general-aviation configuration. for the representative conventional design. Enlarging the winglets provided some improvement in rudder response of the canard design, but the response was NASA Langley Research Center still much lower than that of the conventional design, Hampton. VA 23665-5225 particularly at high lift coefficient corresponding to July 14, 1986 clinib condition. The greater response of the conventional design in this analysis was due to the fact that the conventional design had a much greater 1110- References niciit arm to the rudder and that the propeller slip- st rmni effects were very favorable with the tractor arrangciiieiit in thc) conventional design at climh lift 1. R.ut an, Burt : VariViggc,II/VariE55(. Cariard Desigiis for coefficients. Although the convcmtional design has Flying Qua1itit.s and Pcrforriiancr. Homebuilt Experi- larger dircctional control power, the canard configu- mental Aircrajt Theory and Practice, First, cd., William E. ration is also controllable. This is partly a result of Vaseri, ed., Western Periodicals Co., c.1975, pp. 119 126. the lower levels of lateral-directional stability of the 2. Yip, Long P.: Wind- Tunnel Investigation of a Full-Scale canard configiiration. Canard- Configured General Aviation Airplane. NASA TP-2382, 1985. Conclusions 3. Bihrle. William. Jr.; arid Bowman. James S., Jr.: Inflii- mcc of Wing, Fuselagr. and Tail Dcsigri on Rotational An investigation of the static and dynamic longi- Flow Aerodynamics Beyond Maxirriuni Lift. J. Aircr., tudinal and lateral-directional characteristics of a ca- vol. 18, no. 11, Nov. 1981, pp. 920 925. nard general-aviation aircraft configuration was con- 4. Holmes, Bruce J.; Obara, Clifford J.; and Yip, Long P.: ducted. The following conclusions were reached as a Natural Laminar Flow ExperimentR on Modern Airplane result of this investigation. Surjaces. NASA TP-2256, 1984. 1. During free-flight tests, adequate pitch stabil- 5. Standard jor Metric Practice. E 380-79, Arricricaii Soc. ity was denionstrated throughout the angle-of-attack Testing & Mater., c.1980. rmigc from 7' to 14' for the rearniost center-of- 6. Braslow, Albert L.; and KIIOX,Eugene C.: Simplified gravity position, but the model was more sensitive Method jor Determination of Critical Height oj Distrib- to elevator deflection at moderate angles of attack. uted Rouyhness Particles $or Boundary-Layer IPranuition 2. A low canard configuration offered flight char- at Mach Numbers From 0 to 5. NACA TN 4363, 1958. 7. Staff of thr Langley Rcwarch (kritcr: Exploratory Study actcristics siniilar to those of the high canard config- oj the EJects of Winy-Leadiny-Edge Modifications on the iiratioii biit rcquircd adjustrnent iii caiiard inci- an .Stdl/Spin Behavior oj IL Light General Avicition Airplane. tlciicc aiiglc3 of 5' for iniprovcd flight behavior. NASA TP-1589, 1979. 3. The niodcl exliibitcd a strong fiivora1)lc dilic- 8. Parlrtt, Lyslv P.; and Kirby, Robert H.: Test Tech- (lrikl effect hiit low clirc~tioiialstihility. The addition riiqiivs IJscd by NASA for Irivrstigat ing Dyiiaiiiic Stabil- of a fiiscl;~gr-iiioniitcdvertical fin increased the direc- ity Characteristics of V/STOL Models. J. Aircr., vol. 1, tioiial stability, as did enlarged winglets. The lateral- 110. 5, Sept,. Oct. 1964, pp. 260 266.

10 9. Chambers. Joseph R.: and Grafton, Sue B.: Static and Airplane-- Wind-Tunnel Investigation of High-Angle-of- Dynamic Longitudinal Stability Derivatives of a Pow- Attack Aerodynamic Characteristics. NASA TP-2011, ered 1/9-Scale Model of a Tilt- Wing V/STOL Transport. 1982. NASA TN D-3591, 1966. 10. Newsom. William A.. Jr.; Satran, Dale R.; and John- 11. Kelling, F. H.: Experimental Investigation of a High- son. Joseph L.. Jr.: Eflects of Wing-Leading-Edge Mod- Lift Low-Drag Aerofoil. C.P. No. 1187, British A.R.C., ifications on a Full-Scale, Low-Wing General Aviation 1971.

11 TABLE I . MASS AND GEOMETRIC CHARACTERISTICS OF MODEL

Weight. N (Ib) ...... 230.2 (51.75) Monients of inertia: I,. kg-m2 (slug-ft2) ...... 1.818 (1.341) I,. kg-m2 (slug-ft2) ...... 4.233 (3.122) I-. kg-m2 (slug-ft2) ...... 5.713 (4.214) Fiiselage: Ovcrall length. cm (in.) ...... 167.34 (65.88) Mid center-of-gravity location. nose to c.g., cni (in.) ...... 87.10 (34.29) Wing: Span. cni (in.) ...... 243.84 (96.00) A~cR.cin2 (in2) ...... 6449.7 (999.7) Centerline chord. cm (in.) ...... 38.40 (15.12) Root chord. cin (in.) ...... 32.92 (12.96) Tip chord. cik (in.) ...... 14.63 (5.76) Mean aerodynamic chord. 2. cm (in.) ...... 28.32 (11.15) Spanwise location of E. cm (in.) ...... 51.71 (20.36) Loiigitudinal location of E. cm (in.) ...... FS 110.7 (43.6) Aspect ratio ...... 9.18

Tii1)t’r riit io ...... 0.381 Swccpback angle of quarter-chord. deg ...... 25.7 Dihedral angle. deg ...... -4 Iiicitfciice angk at root. deg ...... 1.2 Incidence angle at tip. dcg ...... -1.8 Airfoil section ...... GA(W)-1 (modified) Aileron (per side): Area. cm2 (in2) ...... 241.9 (37.49) Span. cni (in.) ...... 36.63 (14.42)

Iiil, oard end chord. cni (in.) ...... 6.50 (2.56) Oiitboard end chord. cin (in.) ...... 5.54 (2.18) C;in;ird: 2 . ’> Ar(’ii. Clil (111-) ...... 1580.6 (245.0)

Spmi. cni (ill.) ...... 137.16 (54.00)

(‘110rtl (colistiLlit). CIII (in.) ...... 11.53 (4.54)

SwcYyl)ack ;Lllglt>of q11;Lrt Cr-cl1ol.d. ckg ...... 0 Aspect rat io ...... 11.90 Airfoil section (ref. 11) ...... GU25-5(11)8

. TABLE I . Concluded Elevator: Area. cm2 (in2) ...... 453.1 (70.23) Span. cm (in.) .....: ...... 137.67 (54.20) Chord. cm (in.) ...... 3.30 (1.30) Upper winglet (per side): Area, cm2 (in2) ...... 407.0 (63.08) Span, cm (in.) ...... 32.92 (12.96) Root chord, cm (in.) ...... 18.29 (7.20) Tip chord, cm (in.) ...... 6.40 (2.52) Aspect ratio ...... 2.66 Sweepback angle of quarter-chord, deg ...... 26.3 Dihedral angle (from horizontal), deg ...... 86.0 Incidence angle at root, deg ...... 0 Incidence angle at tip, deg ...... -1.0 Rudder: Area, cm 2.2(in ) ...... 89.6 (13.89) Span, cm (in.) ...... 17.88 (7.04) Lower-end chord, cm (in.) ...... 5.49 (2.16) Upper-end chord, cm (in.) ...... 4.52 (1.78) Lower winglet (per side): Area, cm2 (in2) ...... 47.2 (7.31) Span, cm (in.) ...... 6.38 (2.51) Root chord, cm (in.) ...... 4.62 (1.82) Tip chord, cm (in.) ...... 2.74 (1.08) Sweepback angle of quarter-chord, deg ...... 12 Dihedral angle (from horizontal), deg ...... -60

13 FS 147.6 (58.1)

BL 121.7 (47.9) - Leading-edge droop FS 17.0 (6.7)

fi17.8

Low canard position

Figure 1. Three-view diagram of model with mid c.g. position shown. Linear dimensions are in cm (in.).

14 ORIGINAL PAGE IS .= POOR QUALITY

15 a 0 aE 9) M a YM 3c +9) M El .e 3

d 0 .e .e+ a

16 ORjGINAL PAGE IS OF POOR QUALIm

! I V

C c 0 0 v *I- *? mu- aaJ

aJ aJ -0 U

18 7

h d v.- E V .-c

0 .- I1

19 I W C .-0 .I- n I n C .-0 .I- O Q) CT)

a n m I I a 0 m C wLzzz C 0 .- .I- ciI O

20 21

Grit On Off 1.2

- .2 -20 -10 0 10 20 30 90 50 60 70 80 90 100 a, deg (a) Lift and drag characteristics of canard alone.

Figure 8. Transition grit effects on canard aerodynamics. Referenced to canard area; high canard position; be = oo; i, = oo.

23 1 .q Grit 1.2 0 On 0 Off 1 .o

CL .6

.If

- .2 0 -02 .oq -06 .OB .lo .12 .1If -16 -18 -20 -22

(b) Drag characteristics of canard alone.

Figure 8. Concluded. Figure 9. Canard contribution to model lift and pitching-moment characteristics. High canard position; 6, = oo; 2, = oo.

25 .Lt

0.1

cm - .2

-.II

-.6

0 10 20 - 10

(a) Mid c.g. position.

Figure 10. Elevator effectiveness of the basic configuration. High canard position; i, = Oo. -20 -10 0 10 20 30 '40 50 60 70 80 90 100 01, deg (b) Forward c.g. position.

.'4 be t dq .2 00 0 10 0 20 0.0 n -10 cm

-.2

- .Ll

- -6 -20 -10 0 10 20 30 '40 50 60 70 80 90 100 a,deg

Figure 10. Concluded.

27 -.8 Tlll1111~11~~ I I IIIILHlu

(a) Lift aiid pit,cliiiig-nionie~itcharacteristics.

Figure 1 1. Chiart1 position effects on longitudinal aerodyiianiics. 6, = Oo; i, = 00.

20 0 -02 -0'4 -06 -08 -10 -12 -111 -16 -18 a20 -22 CD

(b) Drag characteristics.

Figure 11. Concluded.

29 cm

-20 -10 0 10 20 30 'A0 50 60 70 80 90 100 a. deg Figure 12. Elcvat,or cffwtivcncss with low canard posit,ion. i,. = 0'. .9

0.C I

Cm -.2

-.9

-.6 1111111111111111111IITI - -8

2 .o

1.8

1.6

1.9

1.2

1 .o

0 .o

-.2

- .9

-.6

-.8 -20 -10 0 10 20 30 90 50 60 70 80 90 100 a, aeg (a) High canard position; 6, = 0'.

Figure 13. Canard incidence effects on lift and pitching-moment characteristics.

31 (I)) Low canard position. .II

cm OmC - .2

- .q

- -6 i 0 Prop off 0 1.6 .12

1 .q

1.2

1 .o

CL .6

.II

CD .2

0 .(

- .2

- .q

- -6

-.8 -20 -10 0 10 20 30 q0 50 a, deg Figure 14. Thrust effects on lift and pitching-moment characteristics. High canard position; 6, = 0'; i, = 0'.

I 33 ORWNAL PAGE IS OF POOR QUALITY

(a) Leading-edge droop removed.

- L-86-347 (b) Leading-edge droop installed.

Figure 15. Tuft studies of stall characteristics of wing. o = 16'.

34 0.c

cm -.2

- .6

- .8

droop

(a) Lift and pitching-moment characteristics.

Figure 16. Effect of leading-edge droop on longitudinal characteristics. High canard position; 6, = 0'; i, = Oo.

35 1.6

1 .'4

1.2

1.o

c, .6

.'4

0 .C

- .2

- .'4

- .6

- -8 0 -02 -0'4 .06 a08 -10 .12 -1'4 -16 -18 -20 -22 CD (b) Drag characteristics.

Figure 16. Concluded.

36 1.6

1.4

deg 1.2 0 22 1 .o

-8

cr, .6 .Y

o .a

- .2

- WY

-.6

-.8 -20 -10 0 10 20 30 40 50 a, deg Figure 17. Landing flap effects on lift and pitching-moment characteristics. High canard position; 6, = 20°; 2, = oo.

37 .03

.02

cn .O1 0

-.01

-.M

.05

.09

- 03

.02

.a1

c1 0

- .01

- .M

- .OY

- -05

Figure 18. Lateral-directional characteristics of configuration in sideslip. High canard position; 6, = oo; i, = oo.

30 .02 cy 0 - .02 - .w LE. droop 0 High On Low On .010 On Off

% .0°5 0

-.005

.02

.01

- .01

Figure 19. Canard and leading-edge droop effects on lateral-directional characteristics. p = 0'; 6, = 0'; i, = Oo.

39 .02

0 CY - .02

-.0q 0 Basic 0 Extended L.E. droop .010

.005

. .005

.02

.01

- .01 -20 -10 0 10 20 30 q0 50 60 70 80 90 100 a, deg

Figure 20. Effect of installation of vortex generators and extended leading-edge droop on lateral-directional characteristics. p = 0'; high canard position; 6, = 0'; i, = Oo.

40 a-0 00 SZ6 90 ??? I I I $

41 0 ggo-"00 sls 99 ?? 1 I 8 6"

42 In m m0

Lo m

lnIn

0(D

In In 0 0

0In II .o".- *In F: .-0 U I!? I!? .I 0- % "d a mIn 2C 9 0m cr 0 U V cuIn & W h cu0 d v

0 M

0 C ys -.02

- .0tf 0 Off On Prop off 0 On On 0 .ON

,003

c .002 "B .001

- .001

.001

- .001 B

- .002

- -003

- .ooq --15 -10 -5 0 5 10 15 20 25 30 35 tf0 Li5 a, deg

Figure 24. Effect of center vertical tail and enlarged winglets on lateral-directional static stability. Low canard position; 6, = 0"; i, = 0".

45 NO ? e

d s U a a

46 mLn

0 m

Ln m

0 m

lnP

InLn

la0

0Ln

9m F 0 a 'ti mLn m0

NLn

Ln I

47 TN 99

48 TNON go~m 00 0 00 0 09 I1 UJ sa a

49 Lom

0)Lo

003

0 0 LoP II 'C1u 0P ... 0 0 II Lo(0 2... c 0 0(D .- .-U 8 3 a 2c: 0u) d 0 c hc =rLo E 8 0- "ti

0ln

Nln

0 N

Lo ? I $ "8584s + I I U a

50 Ln m

0 m

In03

Ln P

c-. c-. 0

m(D

0(D

Lnv)

0Ln

sIn 8 0- 'd

Ln0

m(u

0(u

Lo (nnggos0 TO 0 gg O??? 9 ? 09 I I I U" U Q 4

51 .^ 0 0 01 I

Y -a, bo .-C 3

rr 0

52 L.E. drooD

,Q Stabi I izi ng

-5 0 5 10 15 20 25 30 35 110 115 a, deg Figure 31. Effect of leading-edge droop on dynamic roll stability. High canard position; 6, = 0'; z, = Oo.

53 2 C t Cy. sina yP 0 - I I I I I I I rt I I I I I I I I I -2 Canard position

.2 Cn t C sina P nb 0 .o

- .2

- .q

Sta bi I izi ng clP + Czb sins

-5 0 5 10 15 20 25 30 35 40 95 a?deg

Figurc 32. Effects of canard on damping in roll characteristics. Leading-edge droop on; 6, = 0'; i, = Oo.

54 .9

Cn + Cn. sin a a2 PP 0

-.2..- ...... I.# a,,,,,,

0 Stabil izing

-5 0 5 10 15 20 25 30 35 90 95 a, deg Figure 33. Effect of winglets on damping in roll characteristics. Leading-edge droop on; high canard position; 6, = oo;i, = oo.

55 c =0.2 = 1.0 L c L 35 3c 25 2c Roll rate, deglsec 15 10 5 0 -5 I I I I I I I I I I I I 0 .05 . X .l5 .2G .25 0 .05 .10 .15 .20 .25 Time, sec Time, sec Figure 34. Comparison of calculated aileron response for conventional and canard designs.

cL= 0.2 cL= 1.0 2o r

Yaw rate, deglsec

0- .05 .10 .15 .20 .25 0- .05 . 10 .15 .20 .25 Time, sec Time, sec

Figure 35 Coniparison of calculated rudder response for conventional and canard designs. Standard Bibliographic Page

. Report No. 12. Government Accession No. 3. Recipient's Catalog No. NASA TP-2623 . Title and Subtitle 5. Report Date Wind-Tunnel Investigation of the Flight Characteristics October 1986 of a Canard General-Aviation Airplane Configuration 6. Performing Organization Code

'. Author(s) 505-45-43-01 Dale R. Sat,ran 8. Performing Organization Report No. L-15929 1. Performing Organization Name and Address 10. Work Unit No. NASA Langley Research Center Hampton, VA 23665-5225 11. Contract or Grant No.

13. Type of Report and Period Covered 2. Sponsoring Agency Name and Address National Aeronautics and Space Administration Technical Paper Washington, DC 20546-0001 14. Sponsoring Agency Code

5. Supplementary Notes

6. Abstract A 0.36-scale model of a canard general-aviation airplane with a single pusher propeller and winglets was tested in the Langley 30- by 60-Foot Wind Tunnel to determine the static and dynamic stability and control and free-flight behavior of the configuration. Model variables made testing of the model possible with the canard in high and low positions, with increased winglet area, with outboard wing leading-edge droop, with fuselage-mounted vertical fin and rudder, with enlarged rudders, with dual deflecting rudders, and with ailerons mounted closer to the wing tips. The basic model exhibited generally good longitudinal and lateral stability and control characteristics. The removal of an outboard leading-edge droop degraded roll damping and produced lightly damped roll (wing rock) oscillations. In general, the model exhibited very stable dihedral effect but weak directional stability. Rudder and aileron control power were sufficiently adequate for control of most flight conditions, but appeared to be relatively weak for maneuvering compared with those of more conventionally configured models.

~ ~~ ~ 7. Key Words (Suggested by Authors(s)) 18. Distribution Statement Canard Unclassified-Unlimited General aviation Stall/spin Stability and control Wing leading-edge modifications

9. Security Classif.(of this report) 20. Security Ciassif.(of this page) 21. No. of Pages 22. Price Unclassified Unclassified 57 A04

For sale by the National Technical Information Service, Springfield, Virginia 22161 NASA-Langley, 1986