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Aviation History
1929
1929 - 1348.PDF
62 TO JUNE 27, 192«FLIGHT THE AIRCRAFT ENGINEER Substituting (8) in (7) and re-arranging gives Sw C :*« = S( L letting Ki = -' Sw and -p = 6a • —[(x - e) + Km0TJ + a (b - 1) (9) TAIL SETTING ANGLES TO TRIM -..-*; TECHNICAL LITERATURE SUMMARIES OF AERONAUTICAL RESEARCH COMMITTEE REPORTS These Reports are published by His Majesty's Stationery Offioe, London, and may be purchased directly from H.M. Stationery Office at the following addresses : Adastral House, Kingsway, W.C. 2; 28, Abingdon Street, London, S.W.I; York Street, Manchester; 1, St. Andrew's Crescent, Cardiff; or 120, George Street, Edinburgh; or through any book- seller. •. • atr o CO to <*. - INCIDENCE IN DEGREES •: V: . . For neutral stability the slope of the a* curve againBt incidence is zero, as shown by Fig. 3, which shows the curves of Fig. 2 plotted against incidence instead of against V. It follows, therefore, that da at L a( aw (a: — e) (b~l) o^ .u aw V* - e; Q Sw c o( (1 — 6) The slope of the lift coefficient curve a*, and at are obtained by means of a correction to the slope of the model lift curve for some known aspect ratio, the correction being dependent upon aspect ratio only for a monoplane. The aspect ratio is taken as being span* -*- area, and the correction is based* upon Prof. Prandtl's approximate formula for obtaining the slope of the lift curve for two-dimensional motion (infinite wing) from the slope of the lift curve for a finite wing of known aspect ratio (see translation in N.A.C.A. 116). Re-writing the formula with notation in conformity with the remainder of the article and with the slope of the lift curve given per degree instead of per radian gives :— Aw 104 A, a. + 37-82 (11) Aw = aspect ratio of wing o0 = slope of lift curve for infinite wing ow = slope of lift curve for wing of aspect ratio Aw> so that if a6 = the slope for aspect 6, the slope of the lift curve for any other aspect ratio Aw is :— 1 ' " .;;".- ' . " . . •.• (12)1 B nnn 37-826-303 +——«• A w THE AIRFLOW AROUND A CIRCULAR CYLINDER IN THE REGION WHERE THE BOUNDARY LAYER SEPARATES J-ROM THE SURFACE. By A. Fage, A.R.C.Sc. R. & M. No. 1179 (Ae. 343). (18 pages.) August, 1928. Price 9d. net. The boundary layer around a circular cylinder was located from observa-tions of total head taken near the surface with an exceedingly small tube. The diameter of the cylinder and the range of wind speed selected allowedthe particular range of V,,D/u from 105 to 3 • 5 x 105, through which a marked change in the character of the flow around a cylinder occurs, to be covered inthe experiments. The analysis shows that the distinctive traits of the boundary layer around the cylinder resemble those common to the layeralong a plane surface : that is, there is a critical point on the cylinder where a transition from the laminar to a turbulent state of flow in the boundarylayer begins, and also the transitional region is marked by a rapid opening-out of the layer. The critical value of Reynolds number for the cylinder is ofthe same order of magnitude as that obtained by Van der Hegge Zijnen for a plane surface. The separation of the boundary layer from the surface oceurajust beyond the critical point. WIND-TUNNEL EXPERIMENTS WITH INFINITE CASCADES OF AEROFOILS. By R. G. Harris, D.Sc, F.R.S.E., and R. A. Fairthorne. Presented by the Director of Scientific Research, Air Ministry. R. & M. No. 1206 (Ae. 367). (18 pages and 12 diagrams.) September, 1928. Price 1*. net. The object of the experiments was to provide data for a theory of turbines, in the form of (a) forces on the blades and (b) deflection of the fluid by the. blades.' The experiments were carried out in the 4-ft. wind tunnel from April to July 1928, on the basis of the familiar cascade theory (cf. R. & M. 620*)..A single row of blades: in a turbine is represented by a row of parallel aerofoils placed across the tunnel at various angles to the tunnel axis.Measurements were made of the forces on and the flow behind the central aerofoil of a cascade fitted up in the 4-ft. tunnel. Deflection, total head,pressure, and wind speed over a line in the plane of the centre section covering • one gap were measured. The arrangements tested were: cascade axis at45°, 22i°, and 0° to the plane of the normal cross-section of the tunnel (corresponding to staggers — 22°, 0-5°, and + 2S° respectively), and also10° on each side of these positions ; spacing of the aerofoils, i, 1, and li chords. These arrangements cover the cases of reaction and impulse turbineand compressor rotors and stators. Lift and drag also were measured on the central aerofoil with the cascade removed. The result* can be regarJ.'d as giving approximate values of the forceson an infinite cascade and should be useful for the object for which they were obtained, viz., turbine blade design. The general trend of the forcesis in agreement with the equations of motion of an infinite cascade, but there are considerable discrepancies in detail, probably to be ascribed to theimitations of the experimental method. The results may be of value for other purposes besides turbine-blade designand cascade theory ; for example, in designing guide blades for the bends of a wind tunnel. It will be noted that the minimum loss of energy per degreedeflection occurs in the case *lc = H, <f> = 22i°, where the deflection produced by the 45° arc aerofoil is only 29°. • Preliminary investigation of multiplane interference applied to propellertheory.—R. M.Wood and H. Glauert, A.C.A. 1918-19, p. 541. ON THE EFFECT OF AIR COMPRESSION ON DRAG AND PRESSURE DISTRIBUTION IN CYLINDERS OF INFINITE ASPECT RATIO. By T. E. Stanton. R. & M. No. 1210 (Ae. 370) (5 pages and 2 diagrams.) November, 1928. Price Qd. net. Before attempting the determination of the distribution of pressure round aerofoils at speeds in the neighbourhood of that of sound, it was thought advisable to undertake the relatively simple problem of a cylinder of infinite aspect ratio.In order to eliminate the effect of viscosity, the experiments have all been carried out at the same value of vd/v. From data furnished by the Aero-dynamics Department, it appeared that the value of the drag coefficient of cylinders was nearly constant for values of vdlv in the neighbourhood of20,000, and the present experiments have all been made at approximately this value of the Reynolds number in the 3-in. high-speed wind channel.The range of speed has varied from 0-25o to 2 -0o, and the determinations made have been: (1) the pressure distribution over the section of thecylinder passing through the axis of the wind channel, and (2) the drag ol the whole cylinder from wall to wall of the channel measured on the aero-dynamical balance described in R. <fe M. 1130.* The general form of the pressure distribution curve at the low speeds issimilar to that found by A. Fage (R. & M. 106).T As the speed of sound is approached, the value of the maximum pressure (# = 0) increaaes accordingto the Rayleigh law from 0-5p»2 up to 0-82pt« at u/a = 2-04. At 0 = 180 the value of the pressure coefficient k increases numerically from — 0-51 at«/o = 0-25 to — 0-60 at uja = 0-7. For values of w/a greater than unify the value of k at 180 falls progressively to - 0-08 at uja --• 2-04. (To be concluded.) * A high-speed wind channel for tests on aerofoils.—T. E. Stanton. \ Determination of the pressure distribution round a cylinder.—A. Fage. 522/t
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