Kyle Gompertz

The Application of Flow Control to an Aft-Loaded Low Pressure Turbine Cascade With Unsteady Wakes

The synchronous application of flow control in the presence of unsteady wakes was studied on a highly loaded low pressure turbine blade. At low Reynolds numbers, the blade exhibits a nonreattaching separation bubble under steady flow conditions without upstream wakes. Unsteady wakes from an upstream vane row are simulated with a moving row of bars. The separation zone is modified substantially by the presence of unsteady wakes, producing a smaller separation zone and reducing the area-averaged wake total pressure loss by more than 50%. The wake disturbance accelerates transition in the separated shear layer but stops short of reattaching the flow. Rather, a new time-averaged equilibrium location is established for the separated shear layer. The focus of this study was the application of pulsed flow control using two spanwise rows of discrete vortex generator jets. The jets were located at 59% Cx, approximately the peak cp location, and at 72% Cx. The most effective separation control was achieved at the upstream location. The wake total pressure loss decreased 60% from the wake-only level and the cp distribution fully recovered its high Reynolds number shape. The jet disturbance dominates the dynamics of the separated shear layer, with the wake disturbance assuming a secondary role only. When the pulsed jet actuation was initiated at the downstream location, synchronizing the jet to actuate between wake events was key to producing the most effective separation control. Evidence suggests that flow control using vortex generator jets (VGJs) will be effective in the highly unsteady low pressure turbine environment of an operating gas turbine, provided the VGJ location and amplitude are adapted for the specific blade profile.

Modification of a Transonic Blowdown Wind Tunnel to Produce Oscillating Freestream Mach Number

This paper details the design, construction, and evaluation of a modification to Ohio State’s 6” × 22” Transonic Wind Tunnel to enable modulation of the freestream Mach number. The test section Mach number in this blowdown facility is dynamically set by varying the choke area in a harmonic fashion. Disturbances from the modulated throat area propagate upstream to vary the test section Mach number. Flow properties such as wave propagation speed were evaluated using 1D theory, 2D CFD, and experiments in order to understand the fundamental mechanisms of the Mach oscillation process and subsequent performance limitations of the facility. The current configuration of the Mach oscillation feature enables Mach oscillations between 0.4 and 0.6 at a controlled frequency up to 25Hz and a Reynolds number range of 2 – 16 million per foot.

Active Control of Flow Separation on a Laminar Airfoil

Detailed investigation of the NACA 643−618 is obtained at a Reynolds number of 6.4×104 and an angle of attack sweep of −1≤α≤20  deg via particle image velocimetry and hot-film anemometry. The baseline flow is characterized by four distinct regimes, depending on angle of attack, with each exhibiting unique flow behavior. Active flow control via steady normal blowing is employed at four representative angles; blowing ratio (BR) is optimized by maximizing the lift coefficient with minimal power requirement. Suggestions are made with regard to the flow mechanisms whereby control is effective at each angle. A figure of merit is employed to quantify the efficiency of the actuation. The optimal BRs obtained at Reynolds number equal to 6.4×104 are kept constant to observe the scalability of the control with respect to Reynolds number, ranging over nearly two orders of magnitude. It is observed that this approach is nonideal and that the BR should be optimized with Reynolds number as well. The optimization procedu...

Separation Control Authority of Vortex Generating Jets in a Low-Pressure Turbine Cascade with Simulated Wakes

Detailed pressure and velocity measurements were acquired at Rec = 20,000 with 3% inlet free stream turbulence intensity to study the effects of position, phase and forcing frequency of vortex generator jets employed on an aft-loaded low-pressure turbine blade in the presence of impinging wakes. The L1A blade has a design Zweifel coefficient of 1.34 and a suction peak at 58% axial chord, making it an aft-loaded pressure distribution. At this Reynolds number, the blade exhibits a non-reattaching separation region beginning at 60% axial chord under steady flow conditions. Wakes shed by an upstream vane row are simulated with a moving row of cylindrical bars at a flow coefficient of 0.91. Impinging wakes thin the separation zone and delay separation by triggering boundary layer transition, reducing area-averaged wake total pressure loss by more than 75%. One objective of this study was to compare pulsed flow control using two rows of discrete vortex generator jets (VGJs). The VGJs are located at 59%Cx, approximately the peak Cp location, and at 72% Cx. Effective separation control was achieved at both locations. In both cases, wake total pressure loss decreased 35% from the wake only level and the Cp distribution recovered its high Reynolds number (attached flow) performance. When actuating at 59%Cx the VGJ disturbance dominates the dynamics of the separated shear layer, with the wake disturbance assuming a secondary role only. On the other hand, when actuating at 72%Cx, the efficacy of VGJ actuation is derived from the relative mean shear layer position and jet penetration. When the pulsed jet actuation (25% duty cycle) was initiated at the 72%Cx location, synchronization with the wake passing frequency (8.7Hz) was critical to produce the most effective separation control. A 20% improvement in effectiveness was obtained by aligning the jet actuation between wake events. A range of blowing ratio was investigated at both locations to maximize separation reduction with minimal mass flow. The optimal control parameter set for VGJ actuation at 72%Cx does not represent a reduction in required mass flow compared to the optimal parameter set for actuation at 59%Cx. Differences in the fundamental physics of the jet interaction with the separated shear layer are discussed.

Reynolds Number Scalability for Separation Control on a Laminar Airfoil

Wind tunnel tests are performed on a NACA 643-618 airfoil over a Reynolds number range of 0.06-4.0x10 in order to study several aspects of a laminar airfoil. Studies of blowing flow control and the effect of Reynolds number are the major topics of this effort. The tools used for investigation are surface pressure measurements for lift and wake surveys for drag. Preliminary testing at Re = 64,000 determined that four distinct flow regimes exist with respect to angle of attack: weak laminar separation, moderate laminar separation, laminar separation bubble, and strong leading edge laminar separation. A portion of the study investigates the cause of such dynamic flow physics. Attempts are then made to employ blowing to induce or imitate the laminar separation bubble. By creating the laminar separation bubble, significant lift increase and drag reduction are realized over a broader range of angles of attack. Normal, steady blowing is used because it is a well-characterized device that limits the number of parameters to be varied. Lift is increased significantly and separation is delayed for most cases, where attempts are made to describe the physical mechanisms that induce change. It is observed that the optimal blowing ratio changes between angle of attack regimes because different flow physics are required to induce a change. Studies of Reynolds number scaling found that the lift increased and drag decreased as Reynolds number increased. Importantly noted is that the laminar separation bubble becomes naturally effective at most angles of attack by Re = 180,000. Therefore, the value of flow control diminishes except in regions where strong leading edge separation is the limiting element of the airfoil. This research suggests that the laminar airfoil can be controlled in an energy efficient manner such that high performance is gained across all flight regimes with straightforward actuation.

Measurement Techniques for Shock Movement Capture on a NACA 0012 in Unsteady Compressible Flow

The Ohio State 6” × 22” Unsteady Transonic Wind Tunnel was used to study different methods of experimentally capturing the time-varying location of a shock wave on a NACA 0012 airfoil. The unsteady wind tunnel enables a freestream Mach number oscillation of 0.51 ± 0.1 for a Reynolds number range of 17 – 43 million per meter. Fast-response pressuresensitive paint (PSP) and particle image velocimetry (PIV) were evaluated for their ability to capture time-accurate location of the oscillating shock wave. Unsteady surface pressure and temperature measurements were made using a fast-response bi-luminophore pressuresensitive paint with a single-shot intensity-based technique. This method allows for high spatial resolution surface pressure and temperature maps. Phase-locked planar particle image velocimetry data were also acquired in this experiment, demonstrating an ability to resolve a moving shock. These advanced diagnostic tools enable detailed understanding of the impact of time-varying compressibility both on and off the airfoil surface. The information obtained with each of these techniques is compared in terms of spatial resolution and accuracy with data acquired via surface pressure taps. A NACA 0012 airfoil was tested at 9, 10 and 11 degrees angle of attack and at low reduced frequencies.

Unsteady Compressible Flow on a NACA 0021 Airfoil

,This paper details a modification to Ohio State’s 6” × 22” Transonic Wind Tunnel to enable modulation of the freestream Mach number and application of this oscillatory flow on a NACA 0021 airfoil. The test section Mach number in this blowdown facility is set dynamically by varying the choke area in a harmonic fashion. Flow properties such as wave propagation speed were evaluated using 1-D theory and experiments in order to understand the fundamental mechanisms of the Mach number oscillation process. The current configuration of the facility can produce a Mach number oscillating between 0.44 and 0.65 for a Reynolds number range of 17 – 43 million per meter at sufficiently low oscillation frequencies. The control frequency can be varied up to 25Hz, though the wind tunnel emulates a low-pass filter with a -3dB cutoff frequency of approximately 8Hz. Unsteady measurements in this oscillating freestream flow were then made on a NACA 0021 airfoil. Unsteady surface pressure measurements were made using a fast-acting Pressure-Sensitive Paint with a single-shot lifetime-based technique. Oscillatory standing shocks and variations in surface pressure coefficients due to compressibility effects are observed on the airfoil.

Numerical Simulations of Vortex Generating Jets on Low Pressure Turbine Blades

The combination of low Reynolds number (~20,000) and adverse pressure gradient can induce laminar separation in low pressure turbine sections of a gas turbine propulsion system. Total pressure loss generated by the separation zone has a negative impact on overall engine efficiency. Vortex generating jets have been shown to mitigate laminar separation and offer the ability to actively adapt to varying flight conditions. Various fluid dynamics mechanisms have been observed in jet-affected flow fields such as primary vorticity and turbulent transition. Understanding the appearance and behavior of these mechanisms will aid future implementation of this flow control system into flight hardware. Direct numeric simulations were evaluated for use in analyzing these flow fields. Benchmark data in the blade wake and near the jet hole were obtained from hot film anemometry, particle image velocimetry, and various pressure measurements. Uncontrolled blade wakes behave comparably to experimental observations in terms of t ime-averaged velocity and pressure fields. Steady state and pulsed flow control simulations are shown to successfully mitigate separation effects. Separated shear layer dynamics occurring in response to pulsed jets are observed and discussed. The results s how the presence of non-linear shear layer distortion just prior to separation zone elimination. Short duration, low blowing ratio pulsed jets were shown to be just as effective as steady blowing jets.

Fluid Dynamics of Impinging Wakes and Separation Control on a Low-Pressure Turbine Profile

Detailed pressure measurements were acquired at an axial chord Reynolds number of 20,000 with 3% inlet free stream turbulence intensity to study the effects of position and phase of vortex generator jets employed in an aft-loaded low-pressure turbine cascade in the presence of impinging wakes. Two-dimensional particle-image-velocimetry measurements were made at selected phases of the wake-passing period to investigate the temporal and spatial evolution of coherent spanwise structures and assess their roles in achieving optimal separation flow control. Impinging wakes thin the separation zone and delay separation by triggering transition in the separated shear layer. Though the frequencies differ by orders of magnitude, the vortex shedding from within the separation aligns with the wake forcing frequency. Effective separation control was achieved by pulsing a row of discrete vortex generator jets (VGJs) at a jet penetration threshold location buried within the uncontrolled separation. Synchronization with the periodic wake passing frequency was critical to produce the most effective separation control and recover the cascade’s high Reynolds number performance.

Rotorcraft Airfoil Performance Near Critical Mach Number

6M∞. The Mach number was varied to span a range of subcritical and supercritical conditions, 0.3 ≤ M∞ ≤ 0.7. Differences in compressibility effects between the two airfoils are discussed. The thicker of the two airfoils has a lower drag divergence due to the increased local flow velocities. Three compressibility correction factors (PrandtlGlauert, Karman-Tsien, and Laitone) are compared for their accuracy in predicting airfoil flow behavior in subsonic Mach numbers. The Prandtl-Glauert rule proved to be the most accurate for both airfoils by closely predicting critical Mach number and airfoil performance changes with increasing Mach number.