The purpose of this work is to design an aerodynamically shaped vortex generator set-up to delay flow separation on a swept wing with dihedral of an unmanned aerial vehicle being designed at KTH Royal Institute of Technology. Therefore, a 2.5D CFD study of the wing was performed using the Spalart-Allmaras turbulence model. The optimization of the vortex generator set-up followed a multipoint Pareto strategy to establish an optimum design of the vortex generator vanes including its airfoil cross-section. The resulting vortex generator setup achieved a respective improvement of the maximum lift coefficient and stall angle of attack with respect to the baseline wing of 26.34% and 3 deg as a counter-rotating arrangement and 24.02% and 2 deg as a co-rotating set-up. The optimization procedure showed that the optimum cant angle of the vanes, a geometric parameter not tested in the available literature, contributed to 2.45% of the overall improvement of the maximum lift coefficient. The optimization procedure also showed that the flow separation control performance of the vortex generators is sensitive to its airfoil cross-section, and among all the airfoils tested, the S1223 cross-section showed a superior performance. Finally, the optimum height-toboundary-layer-thickness ratio obtained was 1.301 and a further numerical flow visualization demonstrated that the aerodynamically shaped vortex generators produced a vortex system similar to that of a delta wing, with the difference of the influence of the wing’s wall on the axial flow, that generated a primary vortex submerged in the boundary layer. Because of the resulting leading-edge separation vortex system, the penalty drag of the optimized aerodynamically shaped vortex generators was comparable to that of a conventional, flat-plate vortex generator. Nonetheless, the airfoil-shaped vanes produced higher maximum lift coefficients than the flat-plate vanes configurations.

The purpose of this work is to design an aerodynamically shaped vortex generator set-up to delay flow separation on a swept wing with dihedral of an unmanned aerial vehicle being designed at KTH Royal Institute of Technology. Therefore, a 2.5D CFD study of the wing was performed using the Spalart-Allmaras turbulence model. The optimization of the vortex generator set-up followed a multipoint Pareto strategy to establish an optimum design of the vortex generator vanes including its airfoil cross-section. The resulting vortex generator setup achieved a respective improvement of the maximum lift coefficient and stall angle of attack with respect to the baseline wing of 26.34% and 3 deg as a counter-rotating arrangement and 24.02% and 2 deg as a co-rotating set-up. The optimization procedure showed that the optimum cant angle of the vanes, a geometric parameter not tested in the available literature, contributed to 2.45% of the overall improvement of the maximum lift coefficient. The optimization procedure also showed that the flow separation control performance of the vortex generators is sensitive to its airfoil cross-section, and among all the airfoils tested, the S1223 cross-section showed a superior performance. Finally, the optimum height-to-boundary-layer-thickness ratio obtained was 1.301 and a further numerical flow visualization demonstrated that the aerodynamically shaped vortex generators produced a vortex system similar to that of a delta wing, with the difference of the influence of the wing's wall on the axial flow, that generated a primary vortex submerged in the boundary layer. Because of the resulting leading-edge separation vortex system, the penalty drag of the optimized aerodynamically shaped vortex generators was comparable to that of a conventional, flat-plate vortex generator. Nonetheless, the airfoil-shaped vanes produced higher maximum lift coefficients than the flat-plate vanes configurations.

This paper presents a numerical aerodynamic study of the stall characteristics and flow separation mechanismsof a blended wing body unmanned aerial vehicle being developed at KTH Royal Institute of Technologyand proposes a bio-inspired leading-edge modification to control the separation mechanism andimprove the aerodynamic performance of the aircraft at high angles of attack. A numerical study of theaircraft was performed at cruise speed, corresponding to Reynold’s number of 1.3x106, employing an UnsteadyReynolds Averaged Navier-Stokes solver with the Spalart-Allmaras turbulence model. Numericalresults indicated that the aircraft is characterized by the presence of an unstable longitudinal vortex – visibleat the stall angle of 9 deg – which breaks up at an angle of attack of 10 deg, resulting in an unsteady,full-chord stall cell in the mid-span region of the wing section. To mitigate this phenomenon, a modificationto the leading edge between 0.4 m and 1.8 m wing spans was implemented inspired by the geometryof the nose of a porpoise whale, effectively generating a porpoise (hump) leading-edge inboard section.Preliminary numerical results indicate an increase in stall angle of attack to ∼13 deg and an in maximumlift coefficient to ∼1.0. Furthermore, the porpoise hump allowed controlling the stall behavior of the aircraftby enforcing a wing tip, trailing-edge separation stall achieved by the generation of an extended flowacceleration region at the leading edge.

This paper focuses on the development of a model to represent the longitudinal aerodynamics of a wing and/or flight vehicle in ground effect when height is a function of time including high angles of attack. A general aerodynamic model of a wing and/or airplane in ground effect that includes high angles of attack was created. The aerodynamic coefficients of wings studied herein were obtained by the unsteady vortex-lattice method with Kirchhoff-based correction (UVLM-K). The wind-tunnel measurements presented in the literature were used to validate the UVLM-K. Then, a rectangular wing was simulated at high angles of attack in takeoff and flare, and the aerodynamic characteristics at different heights above ground were obtained along with the h derivatives during these maneuvers. The mathematical model presented herein is capable of modeling unsteady aerodynamic phenomena in ground effect at high angles of attack. When this model was used in the static ground effect, values of R2 equal to or greater than 0.999, 0.981, and 0.993 were obtained for lift, induced drag, and pitching moment coefficients, respectively. In the dynamic ground effect, the model can adjust the aerodynamic coefficient during the maneuvers.