DETERMINING LIFT OF A WING PROFILE AT INSTANTANEOUS CHANGE IN ANGLE OF ATTACK (OPREDELENIE PODEMNOI SILY PROFILYA KRYLA PRI MGNOVENNOM IZMENENII UGLA ATAKI),
Abstract
Formulas for the lift coefficient C sub y of an airfoil and its increase dC sub y/d(alpha) caused by a sudden change in the angle of attack alpha are derived by using the following model of lift generation in a plane-parallel flow: At the instant t = o, when alpha assumes instantaneously a new value alpha sub 1 > alpha, an initial Prandtl vortex leaves the sharp trailing edge of the airfoil. The circulation of this vortex increases with distance from the trailing edge, and at any instant t > o, the circulation must have such a value that the Zhukovskiy hypothesis (about the finite velocities at the points where the airstream leaves the trailing edge) will be fulfilled. The axis of this trailing-edge vortex can be visualized as a small-radius cylinder on which the boundary layer leaving the airfoil surface is wrapped, thus gradually increasing the circulation of the trailing vortex. The circulation of the boundary layer itself (per unit length) is considered as an infinitesimal quantity of the second order; the velocity field induced by the vortex layer is neglected. The problem is solved by using the methods of the theory of functions of a complex variable, assuming that at the instant t = o the lift equals zero. The formulas for C sub y and dC sub y/d(alpha) are presented in terms of a parameter tau = U sub infinity tcos alpha/4, where U sub infinity is the flow velocity at infinity. (Author)
Document Details
- Document Type
- Technical Report
- Publication Date
- Sep 21, 1967
- Accession Number
- AD0674159
Entities
People
- A. A. Kosmodemyanskii
Organizations
- National Air and Space Intelligence Center