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

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Wind-Tunnel Investigation of the Flight Chasacteristics of a Canard Genera1:Aviation Airplane Coniiguration https://ntrs.nasa.gov/search.jsp?R=19870000606 2020-03-20T13:58:42+00:00Z ., , -. / ~ ,a ~ i t 1 1 / I -1 I' I , ,Wind-TunnelInvestigation of the Flight Chasacteristics of a Canard Genera1:Aviation Airplane Coniiguration - I Dale R. Satran j . iNA3A-22-2623 ) UIND- TUNKEL INVESTIGkTION OE 138 7 - 1 d 0 3 j IHZ ELIGHT Cii Ait ACiEiiI STILS C F A CANAtiD Fig32 AVI - G AL- A T iUN AIRP LANE CC NEPGUEATION I 60 [NASA) p csci 01A Ullcla s I H1/02 44247 , 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 com- pressed 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 rud- ders 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.
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