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A supersonic airfoil is a cross-section geometry designed to generate lift efficiently at supersonic speeds. The need for such a design arises when an aircraft is required to operate consistently in the supersonic flight regime. Supersonic airfoils generally have a thin section formed of either angled planes or opposed arcs (called "double wedge airfoils" and "biconvex airfoils" respectively), with very sharp leading and trailing edges. The sharp edges prevent the formation of a detached bow shock in front of the airfoil as it moves through the air.〔Courant & Friedrichs. ''Supersonic Flow and Shock Waves''. Pages 357:366. Vol I.New York: Inter science Publishers, inc, 1948〕 This shape is in contrast to subsonic airfoils, which often have rounded leading edges to reduce flow separation over a wide range of angle of attack.〔Zucker, Robert & Biblarz, Oscar. ''Fundamentals of Gas Dynamics'', pages 226:229. Second Edition.ISBN 0-471-05967-6 John Wiley & Sons, Inc.〕 A rounded edge would behave as a blunt body in supersonic flight and thus would form a bow shock, which greatly increases wave drag. The airfoils' thickness, camber, and angle of attack are varied to achieve a design that will cause a slight deviation in the direction of the surrounding airflow.〔Bertin, John & Smith, Michael. ''Aerodynamics for Engineers''. Third Edition. Prentice Hall. ISBN 0-13-576356-8. Prentice Hall.〕 However, since a round leading edge decreases an airfoil's susceptibility to flow separation, a sharp leading edge implies that the airfoil will be more sensitive to changes in angle of attack. Therefore, to increase lift at lower speeds, aircraft that employ supersonic airfoils also use high-lift devices such as leading edge and trailing edge flaps. ==Lift and Drag== At supersonic conditions, aircraft drag is originated due to: * Skin-friction drag due to shearing * The wave drag due to thickness (or volume) or zero-lift wave drag * Drag due to lift Therefore the Drag coefficient on a supersonic airfoil is described by the following expression: CD= CD,friction+ CD,thickness+ CD,lift Experimental data allow us to reduce this expression to: CD= CD,O + KCL2 Where CDO is the sum of C(D,friction) and C D,thickness, and k for supersonic flow is a function of the Mach number.〔 The skin-friction component is derived from the presence of a viscous boundary layer which is infinitely close to the surface of the aircraft body. At the boundary wall, the normal component of velocity is zero; therefore an infinitesimal area exists where there is no slip. The zero-lift wave drag component can be obtained based on the supersonic area rule which tells us that the wave-drag of an aircraft in a steady supersonic flow is identical to the average of a series of equivalent bodies of revolution. The bodies of revolution are defined by the cuts through the aircraft made by the tangent to the fore Mach cone from a distant point of the aircraft at an azimuthal angle. This average is over all azimuthal angles.〔Woodhull, John. Supersonic Aerodynamics: Lift and Drag. University of Colorado.''Paper presented at the RTO AVT course on Fluid Dynamics Research on Supersonic Aircraft''〕 The drag due-to lift component is calculated using lift-analysis programs. The wing design and the lift-analysis programs are separate lifting-surfaces methods that solve the direct or inverse problem of design and lift analysis. 抄文引用元・出典: フリー百科事典『 ウィキペディア（Wikipedia）』 ■ウィキペディアで「Supersonic airfoils」の詳細全文を読む スポンサード リンク
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