J. A. Jr. Blackwell

Abstract: Nomenclature b = wing span c = wing chord c , = section drag coefficient d c = section lift coefficient c = section normal force coefficient n c = slope of section normal force vs angle-ofa attack curve c = section pitching moment coefficient C = aircraft drag coefficient C = aircraft lift coefficient LI C = pressure coefficient Cju = blowing coefficient M = Mach number P = total pressure q , q = freestream dynamic pressure

Abstract: The transonic flow around a low-aspect-ratio infinite wing in a wind tunnel is investigated by means of numerical simulations, with a focus on shock-induced separated flows. A coarse global grid with far-field boundaries matching those of the test section is subdivided into zones: the flow in the clustered zones near the wing is analyzed by solving the Reynolds-averaged Navier-Stokes equations, while that farther from the wing is modeled with the Euler equations. The results are presented graphically, and it is shown that a mushroomlike separated flow with two counterrotating vortices can be simulated when the correct shock strength is imposed (by careful selection of the artificial dissipation, the boundary conditions, the grid refinement, the algebraic turbulence model, and the geometry representation).

Abstract: An investigation was carried out on two CAST 10-2 airfoil models with chords of 3 in. and 6 in. To evaluate the extent of sidewall influence on airfoil tests at transonic Mach numbers. The tests were conducted in the Langley 0.3-m Transonic Cryogenic Tunnel two-dimensional test section equipped with an upstream sidewall boundary layer removal system which reduces the boundary layer displacement thickness to about 1 percent of model halfspan from an initial 2 percent without boundary layer removal. Test results have shown the changes in the location of the shock on the upper surface of the airfoil to be about the same for both models with and without sidewall boundary layer removal. Even though large differences were noted in the high lift characteristics of the two models, the sidewall boundary layer removal had little effect on the differences. These tests also served to validate the boundary layer removal technique and the associated Mach number correction required with upstream boundary layer removal.

Abstract: A simplified fourwall interference assessment method has been described, and a computer program developed to facilitate correction of the airfoil data obtained in the Langley 0.3-m Transonic Cryogenic Tunnel (TCT). The procedure adopted is to first apply a blockage correction due to sidewall boundary-layer effects by various methods. The sidewall boundary-layer corrected data are then used to calculate the top and bottom wall interference effects by the method of Capallier, Chevallier and Bouinol, using the measured wall pressure distribution and the model force coefficients. The interference corrections obtained by the present method have been compared with other methods and found to give good agreement for the experimental data obtained in the TCT with slotted top and bottom walls.

Abstract: The National Research Council 5ft Trisonic Wind Tunnel enables testing of half-span models at a high Reynolds number cost effectively, however there is the possibility of experiencing relatively larger wall-interference effects compared to full span tests. Subsonic wall-interference effects due to the partially open boundaries at the NRC 5ft wind tunnel are corrected by the “one-variable” method, resulting in corrections in Mach number and angle of attack for the measured quantities. This dissertation introduces the need for adequately correcting for wall interference in subsonic half-model wind-tunnel testing in both pre-stall and post-stall conditions. The one-variable method is described and several aspects that require validation and improvement in this method are pointed out based on specific experimental and numerical results obtained for different test articles. An initial assessment of the influence of the different possible tunnel configurations and computational parameters is defined based on experimental data from a test performed on a scaled model of the Bombardier Global 6000 business jet. This initial study unveils some of the weaknesses of the one-variable method. While this method provides accurate wall corrections during tests in pre-stall conditions; it is unable to generate reliable corrections in stall. Partially, this is due to the development of flow separation on a model tested at subsonic flow conditions reaching stall, not currently accounted for by the potential theory model representation used by the wall-correction methodology at the NRC 5ft TWT. Possible improvements to this singularity representation, paramount to the one-variable method, are then investigated. This revision of the model representation is based on high fidelity free-air CFD solutions for the NASA Common Research Model, validated by experimental data sets obtained from a semispan version of the CRM tested at the NRC 5ft Trisonic Wind Tunnel. The improved potential representation is tested on solid-wall experimental wind-tunnel data to present more reliable behavior of the wall-interference correction estimates when substantial wing stalling is encountered.