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A novel passive flow control concept for subsonic and transonic flows over 2D airfoils is proposed and examined via CFD. The control concept is based on the local modification of the airfoil geometry via a newly proposed Surface-based Trapped Vortex

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A novel passive flow control concept for transonic flows over airfoils is proposed and examined via computational fluid dynamics. The control concept is based on the local modification of the airfoil's geometry. It aims to reduce drag or to increase lift without deteriorating the original lift and/or drag characteristics of the airfoil, respectively. Such flow control technique could be beneficial for improving the range or endurance of transonic aircraft or for mitigating the negative effects of transonic flow on the advancing blades of helicopter rotors. To explore the feasibility of the concept, two-dimensional computational fluid dynamics simulations of a NACA 0012 airfoil exposed to a freestream of Mach 0.7 and Re = 9 × 106 as well as of a NASA SC(3)−0712(B) supercritical airfoil exposed to a freestream of Mach 0.78 and Re = 30 × 106 were conducted. The baseline airfoil simulations were carefully verified and validated, showing excellent agreement with wind tunnel data. Then, 32 various local geometry modifications were proposed and systematically examined, all functioning as a trapped-vortex generator. The surface modifications were examined on both the upper and lower surfaces of the airfoils. The upper surface modifications demonstrated remarkable ability to reduce the strength of the shockwave on the upper surface of the airfoil with only a small penalty in lift. On the other hand, the lower surface modifications could significantly increase the lift-to-drag ratio for the full range of the investigated angles of attack, when compared to the baseline airfoil.

A novel passive approach for controlling the flow in a 2D dynamic stall at variabl freestream is investigated. 2D computational fluid dynamics simulations of an SC1095 airfoil with surface-based trapped vortex generator (STVG) type passive flow control were conducted. The airfoil was exposed to a fluctuating freestream of Mach 0.537 ± 0.205 and Re = 6.1 × 106 (based on the mean Mach number) and experienced a 10° ± 10° pitch oscillation with a frequency of 4.25 Hz. These conditions were selected as an approximation to the flow experienced by a UH-60A helicopter rotor airfoil section in an actual fast forward flight test case. The baseline simulations were cautiously validated with experimental data for both transonic flow and dynamic stall under the variable freestream. Then, 20 different local STVGs type geometry modifications were investigated as a means of passive flow control. Modifications were examined on both the airfoil’s upper and lower surfaces. Results showed that the STVGs were able to mitigate the negative effects of shock-induced dynamic stall. The best geometries could reduce the peak negative pitching moment by as much as 9–23% during the transonic phase of a cycle and by as much as 19–71% during the dynamic stall phase. Also, they were able to reduce peak drag by 8–20% in the transonic phase and by 15–44% in the dynamic stall phase. On the other hand, the lift-to-drag ratio was significantly increased by 3–28% per one rotor cycle. All the above advantages came at virtually no penalty in the lift.

When the rotor blades are at a high advance ratio and/or a high thrust coefficient, the onset of dynamic stall makes accurate prediction of airloads on the rotor blades difficult. Comprehensive rotor analysis codes used in the industry rely on semi-empirical dynamic stall models to generate the aerodynamic coefficients of the blade sections undergoing dynamic stall. However, these models neglect the unsteady nature of the freestream seen by the blade sections compromising the accuracy of the analysis at high speed forward flight and high blade loading conditions. Thus, this thesis aims to investigate the impact of including the unsteady freestream effects in dynamic stall for the prediction of the airloads on the rotor blades. To study the impact of including the unsteady nature of the freestream in dynamic stall, Computational Fluid Dynamics (CFD) was used to generate the unsteady 2D dynamic stall aerodynamic data. The CFD data then served as inputs to the in-house rotor analysis code called Qoptr to generate blade airload results. The flight test data from a steady-level flight case (C T /σ = 0.129, µ = 0.24) from the UH-60A Airloads program was used for validation. The Qoptr blade airload results generated with the unsteady CFD dynamic stall data showed considerably better agreement with the flight test data than the results generated with semi-empirical dynamic stall models, especially in the sectional moment results.