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Aviation History
1936
1936 - 1591.PDF
SUPPLEMENT TO FLIGHT 6526 38 THE AIRCRAFT ENGINEER JUNE 18, 1936 mining the lift over the body which is compatible with downwash and lift considerations over the entire system. The method developed was then applied to the Burnelli UB14-A Monoplane, which has been successfully tested during the last year and also to the larger Burnelli Bomber which proved of unusually high efficiency based on thorough wind tunnel research. The results of the calculation by this method show : (1) The aerofoil body acts as a completely effective lifting surface, in spite of its large thickness, and in spite of the discontinuity produced. (2) Throughout the normal flying range, the body has the effect of producing a continuous lift distribution over the span, except for the break in this curve occurring at the body and wing intersection. (3) The method developed provides an approved basis for stress analysis giving full and official recognition to the lifting characteristics of the fuselage, a subject very important for "structural design in which full advantage is taken of the lift of the fuselage which indicates that the wings are correspondingly relieved of load with valuable saving in weight. APPENDIX. DERIVATION OF EQUATIONS FOR DETERMINATION OF LIFT DISTRIBUTION INVOLVING DISCONTINUITY. Basic Assumptions : (1) The plan form of the Burnelli type wing combination is assumed to conform to the general shape as given in Fig. 1, with the dimensions represented by the notation as shown. (2) The circulation is assumed to be represented by a continuous function of (v) which applies over the entire span, plus a constant increment of circulation which is added over the body portion. (3) The loss in circulation occurring at the points of discontinuity y = ± 6/2) is assumed to be lost between 120 9 0 O J fi-0 30 LIFT CI RVE a • 5' SPAN FEET r~ Fir; ? SMNWISC LIFT DISTRIBUTION \ n 1-0 •8 •6 •* 2 THEOR-MODEL UB-I4A WIN Fl( / u - t / WIND TUNNEL TEST LIFT CURVES- G and BODY COMBIN WON 20 25 30 OC* FROM ZERO LIFT The following is a summary of important physical dimensions and characteristics of this particular aerofoil body combination. (See Fig 1.) Total semi-span = S = 35.615 ft. Body semi-span = b/a = 6.25 ft. Distance between chord planes at Trailing edge = d — 2.00 ft. Chord of body = 280 in. Root chord of wing = 168 in. Tip chord of wing = 47.8 in. Aerofoil section of body N.A.C.A., 4323. Root chord aerofoil section N.A.C.A. 2412. Tip chord aerofoil section N.A.C.A. 2409. In Fig. 2 are shown lift distribution curves for the UB14-A, at angles of attack of 5 deg. and 10 deg. from zero lift. These curves show that the discontinuity at the wing and body intersection, while causing a break in the span distribution curves, does not adversely affect the lift at other points. If these curves of Fig. 2 are integrated, the lift curve as shown in Fig. 3 is obtained. Also, on this graph is plotted the lift curve obtained from a wind tunnel test of a similar model. the points " A " and " B " (see Fig. 1), and to come off as a vortex sheet. With these assumptions : md = \Jbj2) - \~ b',2) where : d = distance between A and B m »B circulation per ft. being lost between point A and B |t = circulation between ± bJ2 |„ = circulation between " bl2 " and " S " and " - 6/2 " and " - S " If we let: (1) •nb A 0V -f l„ = continuous function of (y) V == forward speed V = — s cos 6 REFERENCES: Glauert, "Aerofoil and Airscrew Theory." Van Werner Schmeidler, " Untersuchungen iiber Flugzenge mit veranderli' Flachen." Zeilschrift /Or Flugtechnik uni Motorluftschiffarl, 1931.
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