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Naigle, Shawn ChristopherFlow Control of Compressible Dynamic Stall using Vortex Generator Jets
Master of Science, The Ohio State University, 2016, Aeronautical and Astronautical Engineering
Dynamic stall is an airspeed and maneuver limiting event which occurs on helicopter retreating blades at high advance ratios and is associated with aerodynamic flutter and large negative pitch moments. The potentially violent dynamic stall sequence amplifies pitch link stress and can lead to loss of aircraft control. Steady Vortex Generator Jet (VGJ) blowing has proven to delay the onset of dynamic stall. This work presents results of an experimental investigation into active flow control of a Sikorsky SSC-A09 airfoil undergoing periodic pitching motion in a variety of flow conditions including, steady incompressible, steady compressible, and time-varying compressible freestream, representative of a helicopter rotor system in flight. The airfoil was evaluated at reduced pitching frequencies of 0.025 and 0.050 with a nominal angle of attack schedule, a=9.5°-10.5°cos(wt). Flow conditions were at steady Mach numbers of M=0.2 and M=0.4 and time-varying phase-locked freestream oscillations at Mach number M=0.4+0.07cos(wt), at Reynolds numbers Re=1.5 M and Re=3.0 M. Flow control was achieved through a spanwise row of jets located at 10% chord, oriented normal to the surface, with an effective activated control width of 75% airfoil span. Blowing flow control was evaluated at a jet mass flux ratios from Cq=0.002 to Cq=0.005. Flow control enhancements evaluated include stall penetration, lift and moment improvements, reduction in negative damping, and flow reattachment angle. Quantitative measurements of lift and moment coefficients were calculated through the integration of airfoil surface pressure taps. Qualitative, time-resolved Background Oriented Schlieren (BOS) supplemented surface pressure measurements to assess spanwise averaged dynamic stall vortex progression as well as shock interaction. No optimal mass flux ratio completely controls dynamic stall, but VGJs delayed boundary layer separation, consistently improved cycle average moment, and increased cycle average lift. VGJs triggered an earlier flow reattachment which reduced hysteresis and circuits of clockwise rotation on the CM curve related to negative damping. BOS imagery confirmed the presence of leading edge shock formation and showed VGJ capability to delay shock-induced flow separation. The effectiveness of VGJ flow control is primarily a function of maximum angle of attack, pitching frequency, and freestream compressibility. A comparison of VGJ flow control evaluated on a pitching airfoil in a steady compressible freestream at M=0.4 versus a pitching airfoil in a time-varying compressible freestream at M=0.04+0.07cos(wt) at matched mean reduced frequency and Reynolds number, experienced similar quantitative improvements. Comparison of BOS imagery reveal the same physical VGJ to shear layer interaction between the steady and time-varying freestream cases. Thus, performance measurements based on active VGJs in a steady compressible freestream provide a good prediction of the expected performance measurements when blowing is applied to an airfoil in a low amplitude, time-varying compressible freestream. At low freestream oscillations, airfoil pitching frequency is the dominant factor influencing VGJ effectiveness.

Committee:

Jeffrey Bons (Advisor); James Gregory (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

dynamic stall; flow control; vortex generator jets

Frankhouser, Matthew WilliamNanosecond Dielectric Barrier Discharge Plasma Actuator Flow Control of Compressible Dynamic Stall
Master of Science, The Ohio State University, 2015, Aero/Astro Engineering
Dynamic stall is a performance-limiting phenomenon experienced by rotorcraft in directional and maneuvering flight. Dynamic stall occurs on the retreating blade due to the high angles of attack that are experienced by the blades. Increasing the angle of attack is required to overcome the asymmetry of lift across the rotor disk that is a result from the velocity disparities between the advancing and retreating blade. This works sets out to study and improve the performance of a dynamically pitching NACA 0015 airfoil. The airfoil is subjected to both an incompressible and compressible flow field to simulate the dynamics of a rotor blade with cyclic pitching. In this experimental investigation of dynamic stall flow control, the effectiveness of nanosecond dielectric barrier discharge (NS-DBD) plasma actuation will be evaluated as a means to exert control authority. The NS-DBD plasma actuation is generated by a high-voltage magnetic compression pulsed power supply that was designed and built at The Ohio State University. To measure the influence of plasma actuation on the flow, surface pressures on the airfoil were measured through discrete pressure taps located on both the suction and pressure surfaces. The surface pressures are used to calculate the lift and moment during the dynamic pitching cycle. To visualize the compressibility effects in the outer flow, shadowgraph imagery was used to capture features in the flow around the leading edge of the test article. Tests were conducted at static and oscillating angles of attack at both Mach 0.2 and 0.4, and Reynolds numbers of 1.2 million and 2.2 million respectively. Pitch oscillations were conducted at reduced frequencies of k = 0.05. Actuation frequencies varied from non-dimensional frequencies (F + ) of 0.78 to 6.09. Surface pressures acquired at Mach 0.2 without actuation applied agreed with historical data at static angles of attack, validating that the application of the actuator had limited intrusiveness to the flow. When subjected to pitch oscillations, plasma actuation reduced the severity of lift and moment stall by altering the development of the dynamic stall vortex at Mach 0.2. At Mach 0.4, marginal improvements were gained through actuation. Excitation resulted in a strong dynamic stall vortex that convected more slowly in comparison to the baseline case. Shadowgraph imagery revealed lambda shock waves forming over the first 15 percent of the airfoil chord in the same proximity of th

Committee:

James Gregory, PhD (Advisor); Jeffrey Bons, PhD (Committee Member); Mo Samimy, PhD (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

dynamic stall; nanosecond dielectric barrier discharge plasma actuation; flow control

Singhal, Achal SudhirUnsteady Flow Separation Control over a NACA 0015 using NS-DBD Plasma Actuators
Master of Science, The Ohio State University, 2017, Mechanical Engineering
Flow field surrounding a moving body is often unsteady. This motion can be linear or rotary, but the latter will be the primary focus of this thesis. Unsteady flows are found in numerous applications, including sharp maneuvers of fixed wing aircraft, biomimetics, wind turbines, and most notably, rotorcraft. Unsteady flows cause unsteady loads on the immersed bodies. This can lead to aerodynamic flutter and mechanical failure in the body. Flow control is hypothesized to reduce the load hysteresis, and is achieved in the present work via nanosecond pulse driven dielectric barrier discharge (NS-DBD) plasma actuators. These actuators have been effective in the delay or mitigation of static stall. The flow parameters were varied by Reynolds number (Re=167,000-500,000), reduced frequency (k=0.025-0.075), and excitation Strouhal number (Ste=0-10). It was observed that the trends of Ste were similar for all combinations of Re and k, and three major conclusions were drawn. It was first observed that low Strouhal number excitation (Ste<0.5) results in oscillatory aerodynamic loading in the stalled stage of dynamic stall. At high Strouhal number excitation (Ste>2), this behavior is not observed, as in the static stall cases. Second, all excitation resulted in earlier flow reattachment. Lastly, it was shown that excitation resulted in reduced aerodynamic hysteresis and dynamic stall vortex strength. The decrease in the strength of the dynamic stall vortex is achieved by the formation of excited structures that bleed the leading edge vorticity prior to the ejection of the dynamic stall vortex. At sufficiently high excitation Strouhal numbers (Ste˜10), the dynamic stall vortex was suppressed.

Committee:

Mo Samimy (Advisor); Datta Gaitonde (Committee Member); James Gregory (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

Flow Control; Aerospace; Dynamic Stall; Plasma Actuators