PANAIR uses high-order panel method to predict inviscid subsonic or supersonic flows about arbitrary configuration. Panel method solves linear partial differential equation numerically by approximating configuration surface with panels on which unknown singularity strengths are defined. PANAIR includes advanced software technology as well as advanced aerodynamic technology.

In solving a mixed-type (elliptic-hyperbolic) differential equation in an unbounded region, which is elliptic near infinity, some way must be found to transfer the boundary conditions at infinity to a finite artificial boundary in order to keep the discretized problem finite. The common example of this is transonic flow over an airfoil or wing with subsonic freestream. Here we present an approach which is in many ways analogous to the adaptive wind-tunnel wall concept. Iterative revision of a Dirichlet condition on the common or boundary of the near and far fields results in convergence to a far-field solution that matches the discretized near-field solution in potential and normal derivative across the matching boundary. The far-field equation is either a first-order (FO) Prandtl-Glauert, or a second-order (SO) Poisson-type approximation to the transonic equation. A parameter is easily calculated which gives a good estimate of the accuracy of the far-field solution in either case. Two-dimensional results are given showing the success of the method in reproducing the circulation and Cp for a lifting airfoil. Accurate solutions are given using far-field matching boundaries which are much closer to the airfoil than is permissible with Klunker-type far fields based on multipole expansions. The results are shown to be invariant with the location of the vortex representing the far-field circulation. Thus, we significantly reduce computer time by factors of 3 (FO) and 7 (SO) for mesh density and accuracy equivalent to those of a fixed asymptotic far-field representation. Nonlifting FO calculations for a three-dimensional rectangular wing similarly yield accurate results for a much reduced near field, cutting computer time by more than a factor of 2 in an unoptimized case where the minimum boundary size has not yet been established.

A self-consistent version of the compressible boundary-value problem for configurations with leading-edge vortex separation is formulated, based on the assumption that the compressible flow field is controlled by the linearized potential equation. The stream surface boundary condition and the zero pressure jump condition of the compressible free vortex flows are analyzed; application of the Goethert rule permits the compressible nonlinear boundary-value problem for the subsonic flow domain to be transformed into an equivalent nonlinear incompressible problem. The compressibility corrections developed are used in numerical calculations of subsonic leading-edge vortex flows about planar wing geometries. The sample calculations, employing an inviscid flow model in which the wing and vortex sheets are represented by piecewise continuous quadratic doublet sheet distributions, are applicable to high subsonic Mach numbers.

The application of a new, general, potential flow computational technique to the solution of the subsonic, three-dimensional flow over wings with leading-edge vortex separation is presented. The present method is capable of predicting forces, moments, and detailed surface pressures on thin, sharp-edged wings of rather arbitrary planform. The wing geometry is arbitrary in the sense that leading and trailing edges may be curved or kinked and the wing may have arbitrary camber and twist. The method employs an inviscid flow model in which the wing, the rolled-up vortex sheets, and the wake are represented by piecewise continuous quadratic doublet sheet distributions. The Kutta condition is imposed along all wing edges. Strengths of the doublet distributions as well as shape and position of the free fortex sheet spirals are computed in iterative fashion starting with an assumed initial sheet geometry. The method is verified by numerous computed results.

A higher-order panel method is described for numerical solution of boundary-value problems relating to steady inviscid irrotational incompressible subsonic fluid flow in a domain. Both Neumann and Dirichlet boundary conditions are treated; two types of auxiliary conditions are used to remove the degrees of freedom that arise from specifying only the derivative of the perturbation velocity potential. Four general network types and two expansions of the induced potential kernel are employed in the numerical solution. Some results are presented which illustrate the modeling options and numerical characteristics of the method.