Browsing by Author "Young, A. D."
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Item Open Access Application of the linear perturbation theory to compressible flow about bodies of revolution(College of Aeronautics, Cranfield, 1947-09) Young, A. D.; Kirkby, S.The linearised theory is developed in some detail in order to clarify the differences between two-dimensional and axi-symmetric flow. In agreement with other authors it is concluded that the perturbation velocity on a thin body of revolution in compressible flow is 1/β2 times the perturbation velocity in incompressible flow on a thinner body at reduced incidence obtained by reducing the lateral dimensions of the original body in the ratio (3:1). This result is applied to a representative family of streamline bodies of revolution at zero incidence. Continues…Item Open Access The calculation of the profile drag of aerofoils and bodies of revolution at supersonic speeds(College of Aeronautics, Cranfield, 1953-04) Young, A. D.The effects of viscosity on the aerodynamic characteristics of wings and bodies at supersonic speeds can be assessed if we can calculate… Continues...Item Open Access The equations of motion and energy and the velocity profile of a turbulent boundary layer in a compressible fluid(College of Aeronautics, Cranfield., 1951-01) Young, A. D.As far as the author is aware the derivation of the equations of motion and energy for a turbulent boundary layer in a compressible fluid have not yet been given in detail in any publication. To meet a possible need in this connection this paper puts on record the analysis underlying the equations quoted by the author in Chapter x of the forthcoming Vol. 111 of Modern Developments in Fluid Dynamics.Item Open Access Note on the application of the linearised theory for compressible flow to transonic speeds(College of Aeronautics, Cranfield, 1947-01) Robinson, A.; Young, A. D.It is shown that for finite aspect ratio the linearised theory of compressible flow remains theoretically consistent in the region of transonic speeds, although tis predictions may deviate appreciably from experimental results in that region. The variation of the theoretical lift curve slope of an aerofoil of finite span is considered as the mach number increases from below unity to above unity, and it is shown that the lift curve slope remains finite and continuous.Item Open Access Note on the limits to the local Mach number on an aerofoil in subsonic flow(College of Aeronautics, 1948-10) Young, A. D.Item Open Access Note on the limits to the local Mach number on an aerofoil in subsonic flow(College of Aeronautics, 1948-04) Young, A. D.It has been noted in some experiments that the local Mach number just ahead of a shock wave on an aerofoil in subsonic flow is limited, values of the limit of the order of 1.4 are usually quoted. This note presents two lines of thought indicating how such a limit may arise. The first starts with the observation that the pressure after the shock will not be higher than the rain stream pressure. Fig.1 shows the calculated relation between local Mach number ahead of the shock (M„ 1 ), shock inclination (S), mainstream Mach number (M1) and pressure coefficient just aft of the shock. • (Cp) It is noted that, for given M1 , Cp and .5 ,two shocks are possible in general, a strong one for which Ms , > 1.48, and a weak one for which MS1 < 1.48, and it is argued that the latter is the more likely. The second approach is based on the fact that a relation between stream deflection (8) and Mach number for the flow in the limited supersonics regions on a number of aerofoils has been derived from some. experimental data. Further analysis of experimental data is required before this relation can be accepted as general. If it is accepted, however, then it indicates that the Mach numbers increase above unity for a given deflection is about one-third of that given by simple wave theory (Fig.2). An analysis of the possible deflections on aerofoils of various thicknesses (Fig.3) then indicates that deflections corresponding to local Mach numbers of the order of 1,5 or higher are unlikely except at incidences of the order of5 ° or more, and may then be more likely for thick wings than for thin wings. Flow breakaway will make the attainment of such high local Mach numbers less likely.Item Open Access Note on the velocity and temperature distributions attained with suction on a flat plate of infinite extent in compressible flow(College Of Aeronautics, Cranfield, 1947-08) Young, A. D.The problem considered by Griffith and Meredithl for incompressible flow is here considered for compressible flow, it being assumed that there is no heat transfer by conduction at the plate. Essentially, the method consists of establishing a correspondence between the velocity and temperature profiles for incompressible flow and those for compressible flow, the lateral ordinated being scaled by factors which are functions of the ordinates and of Mach number. The results of calculations covering a range of Mach numbers up to 5.0 are shown in Figs. A and 2.Item Open Access The profile drag of yawed wings of infinite span(College of Aeronautics, Cranfield., 1950-05) Young, A. D.; Booth, T. B.A method is developed for calculating the profile drag of a yawed wing of infinite span, based on the assumption that the form of the spanwise distribution of velocity in the boundary layer, whether laminar or turbulent, is insensitive to the chordwise pressure distribution. The form is assumed to be the same as that accepted for the boundary layer on an unyawed plate with zero external pressure gradient. Experimental evidence indicated that these assumptions are reasonable in this context. The method is applied to a flat plate and the NACA 64-012 section at zero incidence for a range of Reynolds numbers between 106 and 108. Angles of yaw up to 45°, and a range of transition point positions. It is shown that the drag coefficient of a flat plate varies with yaw as Cos½ ˄ (where ˄ is the angle of yaw) if the boundary layer is completely laminar, and it varies as Cos4/5 ˄ if the boundary layer is completely turbulent. The drag coefficient of the NACA 64-012 section, however, varies very closely as Cos˄ foe transition point positions between 0 and 0.5 c. Further calculations on wing sections of other shapes and thicknesses and more detailed experimental checks of the basic assumptions at higher Reynolds number are desirable.Item Open Access Skin friction in the laminar boundary layer in compressible flow(College of Aeronautics, 1948) Young, A. D.