Engineering topics
Publications and source records attributed to Wai, J. C..
Two approaches to calculate turbulent vortical flows over delta wing configurations are illustrated. The first is for a simple delta wing at low speeds using the boundary layer approximation to treat the effects of the secondary separation. The second is for the supersonic case of a generic fighter using the NASA Ames parabolized Navier/Stokes method. Test/theory comparisons are given in both cases.
The NASA/Ames parabolized Navier/Stokes computer code was used to calculate the turbulent flow over the wing/fuselage for a generic fighter at M = 2.2. 18 deg angle-of-attack, and 0 and 5 deg yaw. Good test/theory agreement was achieved in the zero yaw case. No test data were available for the yaw case.
A method using a transonic small disturbance code with successive line over-relaxation is described for treating wing/fuselage configurations with a nacelle/pylon/powered jet. Examples illustrating its use for the NASA transport research model are given. Reasonable test/theory comparisons were obtained.
A transonic small perturbation method has been developed for the analysis of general wing-fuselage-nacelle-pylon configurations with powered jet exhausts. Finite difference successive line relaxation algorithm is used to solve the small disturbance potential equation in conservative form. The nacelle tangency condition and the jet exhaust plume contact conditions are fulfilled in a quasi-cylindrical fashion on a surface fitting the Cartesian grid. The pylon tangency condition is treated in a quasi-planar manner as for the wing. Viscous displacement effects on the wing are modeled by suitable shape changes including the placement of a viscous ramp at the base of the shock. Computed results of a transport configuration show satisfactory correlation with test data.
Nonlifting 10% biconvex airfoils are mounted in a 30 x 40 cm Ludwieg-tube-driven transonic test-section and the flow field recorded with a holographic interferometer. Nitrogen, argon, and carbon dioxide are used as the principal test gases. Experiments are conducted with Reynolds number based on chord of (0.5-3.5) x 10 to the 6th with Mach numbers of 0.70, 0.75, and 0.80. Supporting calculations use inviscid transonic small-disturbance and full-potential computer codes coupled with simple integral boundary-layer modeling. Systematic studies show that significant gamma-effects can occur due to shock-induced separation.
Pressure and speed of sound measurements are used to obtain averaged gas properties in a Ludwieg combustion tube in which the gas is burned just prior to use. Absorption of a laser beam in the exit flow is used to check the predicted temperature uniformity. The burning process and the heat transfer to the cold walls are modeled. Interferograms of Ludwieg tube boundary layers are utilized to develop a new semiempirical boundary layer expression referred to as the integral density thickness, which is then applied to a recently developed linearized analysis of the effects of this layer. Pressure-time histories are determined for a wide range of tube flow Mach number and aspect ratio. Viscous and nonviscous effects are considered. It is shown that although the influence of the integral density thickness on the flow can be large, a modified linear theory can be used to estimate the effects for arbitrary tube flow Mach number and values of the aspect ratio that include cases where the boundary layers have merged.