1、, ,. . LOW.-SPEED WIND-TiIJXNEL RESULTS FOR A THIN WITH DOUBLE SLOTTED.FLAPS F0.R AERONAUTICS 1, :I . -._ . . . . . . .: : . :. .:,WASHINGTO.N - : :. , ,. ,. July 27, 1956 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L56D03 NATIONAL ADVISO
2、RY COMMITTEE FOR AERONAUTICS LOW-SPEED WIND-TUNNEL RESULTS FOR A THIN ASPECT-RATIO-1.85 POINTED-WING-FUSELAGE MODEL WITH DOUBLE SLOTTED FLAPS By Albert E. Brown SUMMARY Results are presented of a wind-tunnel investigation at low speeds of a thin aspect-ratio-1.85 pointed-wing-fuselage model equipped
3、 with double slotted flaps, including the effects of a straight and a delta hori- zontal tail on the static longitudinal stability and the effect of a delta vertical tail on the static lateral stability. The results indi- cated that flap effectiveness increased with increase of flap deflection up to
4、 52.5. For flap deflections greater than 525O, flap effective- ness decreased with increase of flap deflection. With a flap deflection of 52.5, the lift coefficient at an angle of attack of 0 was 0.66 and the maximum lift coefficient was 1.53. Most of the lift-coefficient increment at an angle of at
5、tack of 0 held throughout the angle-of- attack range to near stall. For longitudinal stability of the model with the double slotted flaps deflected, the satisfactory location for a straight or delta horizontal tail was rearward and below the wing chord line extended. However, the straight horizontal
6、 tail studied would not provide longitudinal trim. The delta vertical tail provided static- directional stability of the model except at high lift coefficients and generally increased the effective dihedral. INTRODUCTION Previous investigations (refs. 1 to 4) have shown that large incre- ments of tr
7、im lift coefficient can be obtained on delta-wing airplanes by use of double slotted flaps and that static longitudinal stability can be maintained up through the stall by the use of a properly located horizontal tail. The large increments of lift coefficient were limited to the low and moderate ang
8、le-of-attack range and only relatively small gains in maximum lift coefficient were obtained because of the reduction in flap effectiveness at angles of attack near the stall. Shifting the hinge line of the double slotted flaps to the delta-wing trailing edge Provided by IHSNot for ResaleNo reproduc
9、tion or networking permitted without license from IHS-,-,-2 - NACA RM 5603 (extended double slotted flaps) resulted in a configuration in which the flap effectiveness held to angles of attack near the stall (ref. 5). The present investigation was made to determine whether the attainment of flap effe
10、ctiveness through the angle-of-attack range such as thatof reference 5 might be obtained on an aspect-ratio-1.85 pointed-wing plan form with double slotted flaps. Less rearward movement of the flap would be necessary for this configuration than for that of reference 5, and thus less mechanical compl
11、ication and less diving moment for a given lift-coefficient result. The hinge line of the sweptforward trailing- edge double slotted flap of the present investigation was along the 83 percent chord line which has a sweep of -3 .bo. The hinge line of the constant-chord extended double slotted flap of
12、 reference 5 was unswept. Results of high-speed investigations made on a pointed wing with flap controls having an unswept hinge line are presented in reference 6. Included in the investigation were the effects of a delta and a straight horizontal tail on the longitudinal stability and control char-
13、 acteristics of the pointed wing with double slotted flaps. Both tails had approximately the same variation of lift with angle of attack. In order to make a preliminary evaluation of the static lateral stability of the model, a few tests were made with and without a delta vertical tail at angles of
14、sideslip of f5 through the lift-coefficient range. COEFFICIENTS AND SYMBOLS The results of the tests are presented as standard coefficients of forces and moments about the stability axes. The positive directions of forces, moments, and angles are shown in figure 1. All moments are referred to the qu
15、arter-chord point of the wing mean aerodynamic chord projected on the plane of symmetry as shown in figure 2(a). The coeffi- cients and symbols are defined as follows: CL CD Cm Cn drag coefficient, 9 pitching-moment coefficient, moment qSE rolling-moment coefficient, Rolling moment qsb yawing-moment
16、 coefficient, Yawing moment SSb Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L56DO3 lateral-force coefficient, Lateral force qs variation of rolling moment with sideslip per degree, aC a , measured between p = t5O 21 p variation of yawing
17、moment with sideslip per degree, bCn/ap, measured between p = i5O variation of lateral force with sideslip per degree, dCy/ap, measured between p = +5O free-stream dynamic pressure, TPV , lb/Sq ft 12 wing area, 8.63 sq ft wing mean aerodynamic chord, 2.88 ft, rbl2 c2dy (see fig. 2) uJ 0 wing span, 4
18、.00 ft free-stream velocity, ft/sec mass density of air, slugs/cu ft flap deflection relative to wing-chord plane, measured from flap-chord plane in a plane normal to hinge line (positive when trailing edge is down), deg angle of attack of wing, deg local wing chord, ft local flap chord, measured no
19、rmal to flap leading edge, ft lateral distance from plane of symmetry measured parallel to Y-axis, ft vertical distance from wing-chord plane positive when above chord plane, ft (fig. 2(c) distance of tail quarter-chord position rearward of the wing quarter-chord position, ft (fig. 2(c) incidence of
20、 horizontal tail measured from wing-chord plane, deg 3 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4 Subscripts : max t maximum horizontal tail NACA RM L56DO3 ./“- MODEL AND APPARATUS The model was tested on a single-support strut in the Langley
21、300 MPH 7- by 10-foot wind tunnel. The aerodynamic forces and moments were meas- ured on a six-component mechanical balance system. The pointed wing (fig. 2(a) and table I) was essentially a flat steel plate 5/8 inch thick with beveled leading and trailing edges, having 600 sweep of the leading edge
22、, -23 .lo sweep of the trailing edge, and rounded tips. The thickness varied from 0.012 at the root to a maximum of 0.047 at 0.746b/2. The double slotted flap arrangement tested (fig. 2(b) and tables I1 and 111) consisted of a tapered flap constructed of steel with a wood leading edge and a tapered
23、vane consisting of a steel spar with wood covering. For the flap in the deflected position the leading edge of the vane was along the hinge line which was the 83-percent chord line and the inboard edge of the flap was skewed relative to the fuselage. With a flap deflection of 52.5 the ,inboard tip o
24、f the flap trailing edge was 5.36 inches from the plane of symmetry. For the undeflected position of the flap, relative movement between the vane and flap would be necessary to stow the vane; since stowage space for the vane was not provided in the construction of the model, the vane was removed for
25、 the undeflected- flap tests. The horizontal-tail configurations and locations tested on the model are shown in figure 2( e). The 60 delta tail had an aspect ratio of 2.31 and was constructed of l/ positioning the tails above and below the wing chord line was accomplished by supporting the tail on 1
26、/2-inch steel vertical struts. The 60 delta vertical tail was constructed of 1/8-inch-thick alu- minum and had an area which was 18.4 percent of the wing area (fig. 2( a) ) . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-5 TESTS The tests were made
27、 at a dynamic pressure of approximately 25 pounds per square foot corresponding to an airspeed of about 100 miles per hour. The corresponding Mach number was 0.13. Reynolds number based on the mean aerodynamic chord was approximately 2.7 x 10 . The model was tested at angles of attack from -12O to 3
28、2. 6 Data were obtained for flap deflections of Oo, 40, 45O, 50, 52.5, 55, 60, and 65 for the model without a horizontal tail. Tests of the model with straight and delta horizontal tails were made at several tail incidences with flap deflections of 50 or 52.5 for the tail locations shown in figure 2
29、( e) . The parameters Cnp, Cyp, and C2 were deter- mined from tests at sideslip angles of 25 for the model with tails off, with a delta vertical tail, and with a delta vertical and horizontal tail. P CORRECTIONS Jet-boundary corrections, obtained by methods outlined in reference 7, have been applied
30、 to the angle of attack, the drag coefficient, the pitching-moment coefficient, and rolling-moment coefficient data. Blocking corrections have been applied according to the method of reference 8. A buoyancy correction has been applied to the data to account for a longitudinal static-pressure gradien
31、t in the tunnel. The angles of attack have been corrected to account for airstream inclination. REWLTS AND DISCUSSION The data are presented in the following figures: Longitudinal aerodynamic characteristics: Figure Effect of flap deflection, tail off . 3 Effect of straight tail . 4 Effect of delta
32、tail 5 Lateral stability characteristics: Lateral stability parameters 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 NACA REI 5603 Longitudinal Aerodynamic Characteristics Effect of double slotted flap deflection, tail off.- For the flap- defle
33、ction range (40 to 52.5), substantial increments of lift coeffi- cient, which increased with flap deflection, were obtained throughout the lift-coefficient range to near Cbx with the 2 = 1.56E after- body section (fig. 3). For flap deflections greater than 52.j0, flap effectiveness decreased with in
34、crease of flap deflection. Although a flap deflection of 52.5 produced about the maximum flap effectiveness, the largest value of Chx (1.58) occurred for a flap deflection of 400. (The value of Chx for the model with flaps undeflected was approx- imately 1.20. ) For flap deflections up to 55O the in
35、crement of lift coefficient for a given flap deflection varied only slightly throughout the lift-coefficient range to near Chx (similar to the delta wing with extended double slotted flaps of ref. 5). With a flap deflection of 52.5 the lift coefficient at an angle of attack of Oo for the pointed win
36、g was 0.66 and was 1.53. For the delta wing of ref- erence 4 with a flap deflection of 5z0, the value of the lift coeffi- cient at an angle of attack of 0 was 0.89 and Chx was 1.53. It should be noted that the flap-wing area ratio (0 .llgl) of the present investigation was smaller than that of refer
37、ence 4 (0.1483) and accounts largely for the lower lift coefficient at an angle of attack of 0. An angle of attack of about 23 was required for the plain wing to obtain a lift coefficient of 1.0, whereas an angle of attack of about 7O was required to obtain the same lift coefficient with a flap defl
38、ection of 52.5. cL.max At lift coefficients above 0.65 the lift-drag ratio for the wing with double slotted flaps deflected 52.5 was higher than that for the plain wing. With the exception of some neutral stability over the high lift- coefficient range (fig. 3), the tail-off pitching-moment curves o
39、f the present investigation for flap-deflection angles up to 52.5 generally remain stable up to the stall. Pitching-moment characteristics exhibited undesirable nonlinearity when flap deflections were increased above 52.30, correaponding to flap deflections for which flap effectiveness decreased (fi
40、g. 3(b),. In the delta-wing investigation of reference 4, instabili$y occurred with the tail-off configuration for all flap deflections tested. For lift coefficients up to the maximum, the tail-off pitching-moment curves of the present investigation had less diving moments than for the corresponding
41、 lift-coefficient values of reference 5. Effect of straight tail on the longitudinal stability and control.- A longitudinally stable configuration with flaps deflected occurred with location of the straight tail rearward and below the chord line extended Provided by IHSNot for ResaleNo reproduction
42、or networking permitted without license from IHS-,-,-NACA RM L56D03 7 (1 = 1.56, z = -0.20E) and longitudinal instability occurred for the model with rearward position of the tail above the chord line extended (fig. z = 0.60) longi- tudinal trim would be unattainable throughout the lift-coefficient
43、range. From consideration of the similarity of the pitching-moment curves of the present investigation to those of reference 4 it is believed that longi- tudinal trim could be obtained throughout the lift-coefficient range with the delta tail located rearward and below the chord line extended (2 = 1
44、.56F, z = -0.20). Lateral Stability Characteristics Yawing moment and side force due to sideslip.- The delta vertical tail provided static directional stability of the model with horizontal tail off and flaps deflected 52.5 except at high lift coefficients, where large reversals in both CnP and CyP
45、occurred (fig. 6) . Addi- tion of the delta horizontal tail further increased the static direc- tional stability up to a lift coefficient of 1.06. Rolling moment due to sideslip.- Positive effective dihedral -CzB (fig. 6) increased with increase of lift coefficient throughout most of the lift-coeffi
46、cient range for the tails-off configuration with flaps deflected 52.5; however, a reduction in the effective dihedral occurred at approximately a lift coefficient of 1.1. Ehploying the delta vertical tail generally resulted in additional positive dihedral, the increment of which decreased with incre
47、ase of lift coefficient. The effect of the delta horizontal tail on the rolling moment of the tails- on configuration was small. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 CONCLUSIONS A low-speed investigation to determine the longitudinal cha
48、racter- istics of a thin aspect-ratio-1.83, pointed-wing-fuselage confipation with double slotted flaps including the effects of a straight and a delta horizontal tail on. the longitudinal -stability and the effect of a delta vertical tail on the static lateral stability indicates the following conc
49、lusions: 1. The angle of attack required to obtain a given lift coefficient was considerably reduced with deflection of the double slotted flap. An angle of attack of about 23O was required for the plain wing to obtain a lift coefficient of 1.0, whereas an angle of attack of about 7 was required to obtain the sa
