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Aviation History
1933
1933 - 0456.PDF
60 SUPPLEMENT TO FLIGHT AUGUST 31, 1933 THE AIRCRAFT ENGINEER effect. They are of more value as a check on the corrections for wind tunnel static pressure drop. A fuller analysis of the data in relation to tunnel interference theory appears in R. & M. 1451. In the present report the data are considered mainly as a contribution t o existing information on the drag of streamline bodies. The degree of turbulence present in the boundary layer is found to have a marked effect on overall drag, but no appreciable effect on pressure distribution; whence it may b e inferred that the appreciable scale effect observed with a completely turbulent boundary layer is to be associated with a decrease in skin friction coefficient as Reynolds number is increased. For smooth, well-shaped bodies, the skin friction is closely the same as that on a flat plate of equal area. SPINNING CALCULATIONS ON SOME TYPICAL CASES. BY H. B. Irving, B.Sc, and A. S. Batson, B.Sc. R. A M. No. 1498. (26 pages and 12 diagrams.) February 20 1932. Price Is. 6d. net. The calculations described in this report were made in pursuance of the conception that certain combinations of aerodynamic and inertia] properties of an aeroplane should be good from the point of view of recovery from spinning ; others bad. One is helped to understand how this comes about by considering the behaviour of a model which is free to rotate about a vertical spinning axis and is also pivoted about an axis in the body permitting it to yaw. If the condition of the model is arranged so that the mass spread along the body is decidedly greater than the mass spread along the wings—or, in other words, if the difference, A-B, between the moments of inertia about the longitudinal and lateral axes is negative—it will be found that when the model is rotated about the vertical axis it will always tend to drop one or other wing tip. Now, If weights are added to the outer parts of the wings of the model until these preponderate over the body weights—thus making A-B positive—it will be found on spinning the model that it always tends to set its span horizontal. Calculations were made on a thin wing (R.A.F. 15) biplane with 0° and 30° stagger, and on a R.A.F. 15 monoplane. The outstanding features of the different combinations are given in tabular form in the report. Of these combinations, the one of most interest is perhaps that of wings stable in roll (0° yaw) with large negative A-B. In addition to giving a tendency towards a dangerous spin against the controls, it also has the following features :— (1) May be worse with e.g. forward. (2) May be better with floats. (3) May be worse at low altitudes. (4) Weathercock stability important. (5) May be worse when C-A is small. (6) Dihedral probably important. The dangerous spin may be either flat or steep, the monoplane (thin wing) having the greater liability to the flat spin. Whether this is true of the thick wing-tapered monoplane is not known. The ease of Btable wings and large positive A-B is also interesting in that good recovery is indicated when the elevators only are used. This ease is almost certainly that of the Pterodactyl. The small value of the rolling moment due to sideslips for this aeroplane (Mark IV) should also be in its favour. EXPERIMENTS ON THE REVERSAL OF AILEBON CONTBOL DUE TO WING TWIST. By W. J. Duncan, D.Sc., A.M.I.Mech.E., and G. A. McMillan, M.Eng. R. & M. No. 1499. (22 pages and 4 diagrams.) July 16, 1932. Price Is. Recently the question of loss of lateral control of aeroplanes due to the wing twist which accompanies operation of the ailerons at high flight speeds has come into prominence, more especially in relation to monoplanes with fabric-covered wings. The earliest theoretical study of the problem was made by Roxbee Cox and Pugsley,' who based their analysis on two-dimensional strip theory, and treated the wing as "semi-rigid." Slightly later, one of the present writers pointed out that the reversal speed could be expressed in terms of the aerodynamical derivatives of the theory of wing flutter.t More recently, Pugsley has shown how the formulae of the earlier theory must be modified when the aerodynamical reactions are deduced from the " Prandtl theory," while Pugsley and Brooke have shown§ that the flight speed for complete loss of lateral control of an elastic wing can be determined by a process of successive approximation. The present report gives an account of experiments which have been undertaken in order to test the theoretical formulae, and to investigate the effects of some modifications of the wing and aileron upon the reversal speed. The wing used in the tests was the light aeroplane wing previously used for experiments on wing flutter.|j It Is of reactangular plan form, of 9 ft. span and 3 ft. chord, and the spars are uniform. It was mounted horizontally as a cantilever in a wind tunnel, and the reversal speed was determined from observations of the flexural displacements at the tip and near the root corre sponding to various wind speeds and settings of the aileron. The principal results of the tests can be summarised as follows :— (1) The measured " reversal speed " for the wing without external bracing is in good agreement with the reversal speed calculated by the method ol Roxbee Cox and Pugsley ,1! and with that deduced from the known values of the flutter derivatives. • " Loss of Lateral Control in Aeroplanes due to Elastic Deformation of the Wings."—H. Roxbec Cox and A. G. Pugsley. September, 1931. t " Note on the Beversal of Aileron Control."—W J.Duncan. September, 1931. t " The Aerodynamical Characteristics of a Semi-Rigid Wing, relevant to the Problem of Loss of Lateral Control due to Wing Twisting."—A. G. Pugsley. May, 1932. § '' The Calculation by Successive Approximations of the Critical Reserval Speed for an Elastic Wing."—A. G. Pugsley and G. R. Brooke. September, 1932 ]| See R. and M. 1155, I 50. "The Flutter of Aeroplane Wings."— Frazer and Duncan. August, 1928. H It is to be remarked that the distribution of twist under tip load for this wing is very nearly linear, as assumed in the simplest form of the theory. The theoretical discussion leads to the following conclusions :— (a) The reversal speed is proportional to the square root of the torsional elastic stiffness of the wing. (6) The reversal speed IB independent of the position of the flexural centre.** (c) The reversal speed falls as the centre of pressure of the aileron load moves aft. *• This conclusion was first reached in Ref. 1. It is discussed at length In § 3 of the present report. (2) For a range of variation of aileron hinge position the change in the reversal speed was small. (3) The reversal speed is highest when the aileron control operates at the inboard end of the aileron. (4) Covering the aileron gap raises the reversal speed slightly. (5) When the wing is externally braced at the inboard end of the aileron the reversal speed is raised by a greater amount than would correspond to the increase of torsional stiffness as measured at the wing tip. This result suggests that it will be advisable to select the "reference section " of the semi-rigid theory near the midspan of the aileron. CALCULATIONS OP THE RESISTANCE DERIVAIUVES OF FLUTTER THEORY. PART I. By W. J. Duncan, D.Sc, A.M.I.Mech.E., and A. R. Collar, B.A., B.Sc. R. & M. No. 1500. (14 pages and 2 diagrams.) October 8, 1932. Price 9d. net. In R. & M. 1242,* H. Glauert has calculated the reactions upon a rectilinear aerofoil in two-dimensional accelerated motion in an infinite fluid when the influence of the vorticity in the wake is taken into account, and he arrives at results equivalent to those previously obtained by H. Wagner.t Glauert applies the general formulae to find expressions for the force and moment ujion an aerofoil performing a simple harmonic oscillation in pitch about a point in the chord which advances through the fluid with uniform velocity. In particular he examines the influence of frequency on the damping derivative for pitching motion. In the present paper the analysis is extended to the case where the aerofoil has two independent oscillatory motions, namely, a pitch and a translation at right angles to the general direction of motion. A complete set of derivatives is obtained which are the two-dimensional analogues of the fiexural-torsional derivatives of the theory of wing flutter. It is found that the values of all the derivatives are largely influenced by the frequency of the oscillation. Since the values of the derivatives depend on the frequency in a Rimple harmonic motion, they depend also on the logarithmic increment (or decrement) in a growing (or damped) oscillation. Explicit formulae for the derivatives have also been found for the case of a non-oscilliatory divergence. * " The Force and Moment on an Oscillating Aerofoil." March, 1929. t " Uber die Entstehung des dynamischen Auftriebes von Tragflugefn." Zeit.fiir angeivan'lle Math. u. Meet!., Vol. 5, p. 36 (1925). TESTS OF FLOATING AILERONS ON A BRISTOL FIGHTEK AEROPLANE. PART I. ROLLING BALANCE TESTS ON MODEL WINGS. By F. B. Bradfield, Math, and Nat. Sci. Tri poses, and G. F. Midwood. PART 11. FULL SCALE TESTS. By A. V. Stephens, B.A. Communicated by the Dirdetoi of Scientific Research, Air Ministry. R. & M. No. 1501. (26 pages and 42 diagrams.) January 29, 1932. Price Is. 9d. net. The following model experiments have been made in connection with full-scale tests of floating ailerons of a Bristol Fighter. The port and starboard ailerons are interconnected in such a way that they have freedom to move in the same sense, while they are moved differentially by moving the control column. The port and starboard ailerons are constrained to float at a common setting when the column is central; and when the aeroplane has a rate ot roll and the effective incidence varies along the wing, a mean floating angle is taken up. At zero rate of roll, rolling and yawing moments and the floating anule of the ailerons were measured from a = 0° to 36°, for from ± 15" aileron angles. For a range of ps/V from 0 to 0-5, and angles of incidence irom 2° to 60°, the same quantities were measured for aileron angles of 0° and ± 1 •) • both with the ailerons floating and in the standard position. Floating the ailerons reduces, but does not entirely eliminate, the auto-rotation range. , The aileron control with Uoating ailerons is better than with stanuara ailerons in that the yawing moment due to the ailerons is considerably less positive, but there is little increase in rolling moment, and under Nin\ conditions the rolling moment is less than for standard ailerons, being reverse in sign at a very large values of a when there is a rate of roll. WIND TUNNEL TESTS ON A BRISTOL " BULLDOG " FITTED AVITH A THIN TOWNEND RING. By W. G. A. Perring, R.N.C. Communicated by the Director of Scientific Research, Air Ministry. R. & M. No. 1504. (18 pages and 8 diagrams.) August 4, 1932. Price Is. net. The present report continues the investigation commenced in R. <v J" Zj into the effect of a thin Townend ring on the performance of the i» Bulldog. The tests have been made on a one-fifth scale model represem» of a single-seater Bristol Bulldog fitted with a Jupiter VII engine. «"• thf ment of the drag of the model without airscrew, and the resuitarn de] thrust and drag of the model with airscrew have been made IS) tor u n without ring, and (ft) for the model fitted with a thin polygonai '<.>_ tbe Townend ring, the tests being carried out for three angular settin r Tests without, airscrews showed that the drag reduction due to_ f^_6» was greatest with the ring at -9°, was slightly less for a ring an^H w,^ _j». and was approximately one-half the best value when the ring aiU-'*. ,' ll(.tion The addition of the slipstream had only a slight effect on the drat • • (hoge due to the ring at -6° and -9°, when the conditions corresponaeu wffi for level flight, but it improved the performance of the ring J"?5" cticaUy at -3°. The tests also showed that the addition of the ring had P'- no effect on the airscrew torque. . . ,.,,lupared The tests result in a predicted rate of climb of 460 ft. per "Wjuw^ m0(lel with 360 ft. per minute achieved full scale; the diKeience™1" ,, jn the and full scale being probably partlv, or wholly, due to a dinei form. scale effect with and without the ring. The tests also show_ that.UKI^ 0| ance of the full-scale aircraft might be improved by a furtner i" ^ t]uougn the improvement already achieved if the ring full scale was cnaiw 3° to a larger negative angle. ^_-—• * " Some wind-tunnel experiments on the cowling of air-cooled en Perring. 870 A
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