Browsing by Author "Robinson, A."
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Item Open Access The aerodynamic derivatives with respect to sideslip for a delta wing with small dihedral at supersonic speeds(College of Aeronautics, 1947-12) Robinson, A.; Hunter-Tod, J. H.Expressions are derived for the sideslip derivatives on the assumptions of the linearised theory of flow for a delta wing with small dihedral flying at supersonic speeds. A discussion is included in the appendix on the relation between two methods that have been evolved for the treatment of aerodynamic force problems of the delta wing lying within its apex Mach cone. Continues…Item Open Access Aerofoil theory for swallow tail wings of small aspect ratio(College of Aeronautics, Cranfield., 1950-10) Robinson, A.A method is developed for the calculation of the aerodynamic forces acting on a ‘swallow tail’ wing of small aspect ratio. Lift, induced drag, and aerodynamic centre position of simple swallow tail wings (Fig. 1 (b)) are computed as an application. For a given incidence, lift and induced drag are, within the limits of the theory, proportional to aspect ratio and independent of speed. The chordwise life distribution rises linearly from zero at the apex, drops rapidly in the region of the root chord trailing edge, and then decreases gently to zero.Item Open Access Bound and trailing vortices in the linearised theory of supersonic flow, and the downwash in the wake of a delta wing(College of Aeronautics, Cranfield, 1947-10) Robinson, A.; Hunter-Tod, J. H.The field of flow round a flat aerofoil at incidence can/may be regarded in linearised theory as the result of both bound and trailing vortices for supersonic as well as for low speed flight. This leads to a convenient method, given the lift distribution over an aerofoil, for calculating the flow round it at supersonic speeds. As an application of the results the downwash is calculated in the wake of a delta wing lying within the Mach cone emanating from its apex. The downwash is found to be least just aft the trailing edge and is everywhere less that the downflow at the aerofoil. It increases steadily to a limiting value which is attained virtually within two chord lengths of the trailing edge. The ratio of the downwash at any point in the wake to the downflow at the aerofoil decreases with increasing Mach number and apex angle.Item Open Access The effect of the sweepback of delta wings on the performance of an aircraft at supersonic speeds(College of Aeronautics, Cranfield., 1947-03) Robinson, A.; Davies, F.T.The variation with sweepback of total drag of an aircraft in level flight at supersonic speeds is calculated. It is shown that sweepback is not uniformly beneficial, but that in general the optimum amount of sweepback depends in the design speed and altitude.Item Open Access Interim report on the research into the use of formula for calculating economic batch sizes(College of Aeronautics, 1959-12) Robinson, A.This report analyses the replies and subsequent conversations held with companies. Based on this analysis it is shown that various factors have to be considered Then deciding what is an Economic Batch size, and that the batch size given by the formula Q =√200AR UtI is not necessarily the size which provides the best economic gain to a company. A relationship is formulated between the percentage increase in total unit cost for given variations from the batch size: so that companies may determine what alteration to this batch size is acceptable or desirable under given operating conditions. The factors to be included in the terms A (set up cost) Ut (unit manufacturing cost), and. I (holding charges expressed as a Percentage of total manufacturing costs) are discussed, and practical methods of calculating Economic Batch Quantities discussed. Finally a review is made of further research work required.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 On some problems of unsteady supersonic aerofoil theory(College of Aeronautics, Cranfield, 1948-05) Robinson, A.Unsteady supersonic flow round an aerofoil of infinite span is considered in the first part of the paper.It is shown that the pressure at any given point of an aerofoil under forward acceleration can be analysed into three components, one of which is the steady (Ackeret)pressure due to the instantaneous velocity, while of the other two, one depends directly on the acceleration, and one on the square of the velocity, during a limited time interval preceding the instant under consideration. However, the difference between the total pressure and the "steady pressure component" is such that it can be neglected in all the definitely supersonic conditions which are likely to occur in practice. The oscillatory supersonic flow round a Delta wing inside the Mach cone emanating from its apex is considered in the second part of the paper. Particular "normal" solutions are obtained by means of a special system of curvilinear coordinates. It is shown that the velocity potentials corresponding to vertical and pitching oscillations of the wing can be represented by series of such normal solutions. The assumptions of linearised theory arc adopted throughout.Item Open Access On the integration of hyperbolic differential equations(College of Aeronautics, Cranfield, 1948-07) Robinson, A.See full text for abstract.Item Open Access Source and vortex distributions in the linearised theory of steady supersonic flow(College of Aeronautics, Cranfield, 1947-10) Robinson, A.The hyperbolic character of the differential equation satisfied by the velocity potential in linearised supersonic flow entails the presence of fractional infinities in the fundamental solutions of the equation. Difficulties arising from this fact can be overcome by the introduction of Hadmard’s ‘finite part of an infinite integral’. Together with the definition of certain counterparts of the familiar vector operators this leads to a natural development of the analogy between incompressible flow and linearised supersonic flow. In particular, formulae are derived for the field of flow due to an arbitrary distribution of supersonic sources and vortices. Applications to Aerofoil theory, including the calculation of the downwash in the wake of an aerofoil, are given in a separate report.Item Open Access Wave reflection near a wall(College of Aeronautics, Cranfield., 1950-05) Robinson, A.The field of flow due to a shock wave or explanation wave undergoes a considerable modification in the neighbourhood of a rigid wall. It has been suggested that the resulting propagation of the disturbance upstream is largely due to the fact that the main flow in the boundary layer is subsonic. Simple models were produced by Howarth, and Tsien and Finston, to test this suggestion, assuming the co-existence of layers of uniform supersonic and subsonic main stream velocities. The analysis developed in the present paper is designed to cope with any arbitrary continuous velocity profile which varies from zero at the wall to a constant supersonic velocity in the main stream. Numerical examples are calculated and it is concluded that a simple inviscid theory is incapable of giving an adequate theoretical account of the phenomenon. The analysis includes a detailed discussion of the process of continuous wave reflection in a supersonic shear layer.Item Open Access Wing body interference at supersonic speeds(College of Aeronautics, Cranfield, 1947-04) Kirkby, S.; Robinson, A.The increment in lift due to wing-body interference at supersonic speed is calculated approximately for an untapered wing without sweepback.