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An investigation to determine the low-speed lateral-directional characteristics of a generic 74-degree delta wing-body configuration employing the latest generation, gothic planform vortex flaps has been conducted. In addition, the theoretical estimates from VORSTAB were compared against experimental data to aid in documenting this new method. VORSTAB is an extension of the Quasi-Vortex-Lattice Method of Lan which empirically accounts for vortex breakdown effects in the calculation of longitudinal and lateral-directional aerodynamic characteristics. The experimental results indicated that leading-edge deflections of 30 and 40 degrees significantly reduce the magnitude of the wing effective dihedral relative to the baseline for a specified angle of attack or lift coefficient. For angles of attack greater than 15 degrees, these flap deflections reduce the configuration directional stability despite improved vertical tail effectiveness. Asymmetric leading edge deflections are shown to be inferior to conventional ailerons in generating rolling moments. Asymmetric leading-edge deflections are effective in producing side force at moderate to high angles of attack. VORSTAB lateral-directional calculations provide ballpark estimates at low to moderate angles of attack. The theory does not account for vortex flow induced, vertical tail effects at high angles of attack and should not be used for this angle of attack region. The empirical formulae for predicting vortex burst effects are not reliable in their present form. Although the basic trends are correct, the magnitude of the predicted vortex burst effect is typically over-estimated. / Master of Science

Steady, inviscid, supersonic flow past conical wings is studied within the context of irrotational, nonlinear theory. An efficient numerical method is developed to calculate cones of arbitrary section at incidence. The method is fully conservative and implements a body conforming mesh generator. The conical potential is assumed to have its best linear variation inside each cell; a secondary interlocking cell system is used to establish the flux balance required to conserve mass. In regions of supersonic cross flow, the discretization scheme is desymmetrized by adding the appropriate artificial viscosity in conservation form. The algorithm is nearly an order of magnitude faster than present Euler methods. It predicts known results as long as the flow Mach numbers normal to the shock waves are near 1; qualitative features, such as nodal point lift-off, are also predicted correctly. Results for circular and thin elliptic cones are shown to compare very well with calculations using Euler equations. This algorithm is then implemented in the design of conical wings to be free from shock waves terminating embedded supersonic zones adjacent to the body. This is accomplished by generating a smooth cross-flow sonic surface by using a fictitious gas law that makes the governing equation elliptic inside the cross-flow sonic surface. The shape of the wing required to provide this shock-free flow, if such a flow is consistent with the sonic surface data, is found by solving the Cauchy problem inside the sonic surface using the data on this surface and, of course, the correct gas law. This design procedure is then demonstrated using the simple case of a circular cone at angle of attack.

Thesis. 1977. M.S.--Massachusetts Institute of Technology. Dept. of Aeronautics and Astronautics. / This electronic version was submitted by the student author. The certified thesis is available in the Institute Archives and Special Collections. / MICROFICHE COPY AVAILABLE IN ARCHIVES AND AERONAUTICS. / Includes bibliographical references. / M.S.

A modified vortex lattice method is developed to solve for the nonlinear three-dimensional unsteady incompressible flow over delta wings. Symmetric motions (pitching, heaving) and asymmetric motions (roll, yaw) are considered. Then the method is generalized to treat the nonlinear three-dimensional steady flow for bodies and wing body combinations. Numerical examples include variety of shapes and comparison with existing experimental data and other numerical methods over a wide range of angle of attack shows good agreement. For bodies alone the results deteriorate downstream of the separation region, while for wing-body combinations the agreement with the experimental data is good as long as the body separation effect is not large. The developed computer codes should provide a useful tool in the design of aircraft and missiles. / Ph. D.

Ph. D.

Subsonic wind tunnel investigations were conducted on a 60° swept, flat plate, delta wing with a leading edge vortex flap. The pressure distributions were measured over a range of angles of attack starting from zero to 40° in 5° interval and flap deflection angles from zero to 45° with 5° increments at a Reynolds number of about 2.14 x 10‘ based on the root chord. The flow visualization experiments were performed from zero degree to the stall angle, with ten different flap deflection angles at the same Reynolds number. The mean flow field was measured at angles of attack l0° and 15° with the flap deflection angles of l0° and 30° at a Reynolds number of about 1.50 x 10°. The experimental results shows that the leading edge vortex flap is an effective means to control the vortex flow over a delta wing. The optimum flap deflection angles were found where the primary vortex was confined to the leading edge vortex flap, thus producing a thrust on the flap. It was found that flap deflection could be used to restore a vortex flow from burst vortex condition. / Ph. D. / incomplete_metadata

A Numerical Solution is presented for the problem of flow past a highly swept, slender wing with sharp leading edges. The lifting surface is modelled as a bound vortex sheet, while the wake is modelled as a force-free vortex sheet. The solution is obtained by the use of a unsteady Vortex-Lattice Method which includes the effect of leading edge separation. Numerical predictions for the aerodynamic loads and pressure distributions are compared with experimental data. A 75° Delta wing and a 60° Delta wing with Leading Edge Vortex flaps in uniform, symmetric and steady flow are studied. Uniform and cosine distributions are used to determine the effect of lattice shape on the solution. The results show that good aerodynamic load predictions are obtained by this Vortex-lattice method. The results also indicated that fewer cosine distribution control points predict pressures as well as the use of a larger number of uniform distribution control points. The numerical results for wings with LEVFs show good agreement with experimental data away from the trailing edge. This may be due to the viscous effects in the experiment not modelled in this method. It is also apparent that the size of the wake, trailing and leading edge wakes, is the important factor effecting computation times. / M.S.

Using a general, unsteady, nonlinear vortex lattice method, the aerodynamic loads have been found on a 75° delta wing with and without leading edge vortex flaps. The flap had an area approximately 26 percent of the wing area with a constant chord of 6.7 percent of the wing mean aerodynamic chord and was deflected at 30°. Results for lift, drag, axial force, and pitching moment coefficients are compared with experimental data and show very good agreement. Individual pressure difference coefficients along the wing and flap are also presented and compared with experimental data. Overall, the method shows the leading edge vortex flap to be very effective in reducing drag while maintaining lift comparable to that of the plain wing. / Master of Science