AERODYNAMIC FLOW CONTROL AND ADVANCED DIAGNOSTICS (AFCAD) RESEARCH GROUP: Research
Accordions
If the cross section of a body perpendicular to the flow is significant, it is called a bluff body. Wind moving past skyscrapers and flow over a moving truck or car are some examples of bluff-body flow. Because of the shape of these bodies, there exists a region of considerable size behind them where the flow is slowed down. This region is termed the wake. Due to the existence of a large wake, bluff bodies experience pressure drag. Sometime, large vortices are shed periodically behind the body. This phenomenon is called vortex shedding. It creates an unsteady, periodic force which makes the bluff body vibrate and may lead to structural failure. Two kinds of control can be implemented to control bluff-body flows. One is called passive control where the geometry of surface is modified, for example by putting tabs, streaks or dimples on the surface. In the other method, external energy is input to some type of electronic actuator which operates at a specific condition. This is called active control. Some examples are blowing and suction, synthetic jets, and plasma actuators.
Flow past a cylinder is often used as a canonical case of bluff-body flow. Natural vortex shedding is clearly visible in this flow visualization image using a smoke wire technique.
In our group, research is focused on active control of bluff-body flow using dielectric barrier discharge (DBD) plasma actuators mounted on a circular cylinder. DBD actuators are constructed by separating two electrodes (called the exposed and buried electrodes) with a dielectric material. Figure 2 shows a schematic of a DBD actuator. When high voltage is applied between them, plasma is created in front of the exposed electrode. The ionized particles collide with the neutral air molecules and create a body force. This whole action gets revealed in the form of an air jet. This jet can be used for flow control purpose.
Schematic of dielectric barrier discharge (DBD) plasma actuator.
Segmented Plasma Actuator
Segmented DBD plasma actuator implemented on cylinder.
The current research employs the concept of segmented actuation (Gregory et al., 2008). To achieve this, buried electrodes have been mounted on the cylinder only at specific spanwise locations. The main idea behind this kind of actuator design is to promote three-dimensionality in the wake. Greater three-dimensionality encourages greater momentum transport in the wake and consequently less pressure drag.
Flow visualization image showing attenuation of vortex shedding for high power forcing (13 W) with segmented actuator. No signatures of vortical structures are observed in the near-wake.
Initial results from this research indicated considerable drag reduction achieved with segmented actuation compared to conventional two-dimensional forcing (with a continuous buried electrode rather than a segmented one). Vortex shedding was greatly attenuated at a power level of 13 watts. At a lower power level of 5 watts, the wake responded differentially to the forcing based on the spanwise location across the cylinder. In the region behind cylinder locations without plasma formation, the vortices migrated closer to the wake centerline whereas behind regions of plasma formation on the cylinder, they were displaced away from the centerline.
Measurements of cylinder wake with low-power forcing: region behind cylinder without plasma formation (left) and region behind plasma formation (right).
Attention is being focused on understanding the relation between actuator performance and the wavelength of actuation, as well as the effect on wake modes. The preliminary mechanism for two-dimensional forcing has been assumed by some researchers as a phase-mismatch in the Karman vortex street. However, it is yet to be confirmed whether the same mechanism is responsible in the present case of three-dimensional forcing or if it is related to some other cause such as modification of the separation point. Hot-wire anemometry and particle image velocimetry are being deployed to understand the complicated flow behavior in the wake of the circular cylinder.
Representative Publications
- Gregory, J.W., Porter, C.O., Sherman, D.M., and McLaughlin, T.E., 2008, “Circular Cylinder Wake Control using Spatially Distributed Plasma Forcing,” AIAA 2008-4198, 4th AIAA Flow Control Conference, Seattle, WA.
- Bhattacharya, S., and Gregory, JW, 2012, "Effect of Three-Dimensional Plasma Actuation on the Wake of a Circular Cylinder," AIAA 2012-0907, 50th AIAA Aerospace Sciences Meeting, Nashville, TN. doi: 10.2514/6.2012-907
- Bhattacharya, S., and Gregory, JW, 2012, "Study of the Wake of a Circular Cylinder under Spatially and Temporally Modulated Plasma Actuation," AIAA 2012-2957, 6th AIAA Flow Control Conference, New Orleans, LA. doi: 10.2514/6.2012-2957
Aerodynamic bodies subjected to pitching motions or oscillations exhibit a stalling behavior different from that observed when the flow over a wing at a fixed angle of attack separates. The latter phenomenon is referred to as static stall, since the angle of attack is fixed. In the case of a dynamically pitching body, such as an airfoil, the shear layer near the leading edge rolls up to form a leading-edge vortex (LEV) which provides additional suction over the upper airfoil surface as it convects downstream. This increased suction leads to performance gains in lift and stall delay, but the LEV quickly becomes unstable and detaches from the airfoil. The LEV detachment is accompanied by a dramatic decrease in lift and sudden increase in pitching moment. Dynamic stall is not a well-understood phenomenon despite its importance to the performance and operational limits of helicopters, flapping wings, and wind turbines. In fact, dynamic stall can lead to violent vibrations and dangerously high loads in these aerodynamic applications, leading to fatigue and structural failure.
Mach number map of dynamic stall flow field near leading edge of an airfoil. "Warmer" colors indicate higher Mach number.
Where it concerns helicopters specifically, many experimental, computational, and theoretical investigations have been undertaken in order to grasp the fluid physics at play when an airfoil executes pitching oscillations in a steady low-speed airstream. However, the scenario is complicated when considering a helicopter travelling at especially high forward-flight speeds. In this case, the impact of the time-varying relative velocity seen by the rotor plays a significant role in the dynamics stall process. At sufficiently high speeds, compressibility effects are encountered and shock waves can form on the advancing rotor and they can induce stall. The image at left shows one instantaneous Particle Image Velocimetry (PIV) measurement acquired at the onset of shock-induced stall. This picture was acquired in the 6” × 22” Transonic Wind Tunnel at Ohio State University. This is a very unique facility which allows experimentalists to model and understand the unsteady fluid dynamics associated with compressible dynamic stall with enhanced fidelity. In addition to investigating the effects of airfoil pitching motions, the tunnel was upgraded in 2008 to produce a time-varying compressibility condition by oscillating the free stream Mach number. This is accomplished by rotating a set of oblong vanes at the wind tunnel choke point downstream of the test section. As they rotate, the area ratio (between the choke point and the test section) varies harmonically, as does the test-section Mach number. Dynamic airstream oscillations can be generated over a range of Mach number and frequency relevant to advanced rotorcraft. These features make this facility suitable for experimental modeling of compressible dynamic stall as well as exploration of novel flow control strategies to expand the operating envelope of rotorcraft.
In addition to conventional measurement techniques, the wind tunnel is equipped with optical access. This makes it possible to acquired data using PIV and the fast-responding pressure-sensitive paint (PSP) formulations developed and applied by our research group. These advanced diagnostics may be utilized in dynamic stall investigations to better understand the time-accurate flow topology both on- and off-body. Experimental data gleaned from these investigations are all important for validation of computational models for dynamic stall, which could lead to improved performance gains and more defined operating envelopes as a matter of safety for helicopters.
Representative Publications
- Gompertz, K., Kumar, P., Jensen, C.D., Peng, D., Gregory, J.W., and Bons, J.P., 2011, "Modification of a Transonic Blowdown Wind Tunnel to Produce Oscillating Freestream Mach Number," AIAA Journal, vol. 49, no. 11. doi: 10.2514/1.J051090
- Juliano, T.J., Peng, D., Jensen, C.D., Gregory, J.W., Liu, T., Montefort, J., Palluconi, S., Crafton, J., and Fonov, S., 2011, "PSP Measurements on an Oscillating NACA 0012 Airfoil in Compressible Flow," Proceedings of the 41st AIAA Fluid Dynamics Conference and Exhibit, Honolulu, HI (AIAA 2011-3728).
A typical PSP measurement system to acquire two-dimensional surface pressure distributions. PSP is based on the interaction between oxygen and special luminescent molecules.
Knowledge of the time-varying pressure field acting on an aerodynamic surface is insightful in understanding complex flow structures. For example, interactions between the fan and downstream stator inside a jet engine contribute to high-cycle fatigue and engine tonal noise, and unsteady inlet distortions have a significant impact on engine operability and efficiency. Helicopter blades operate under the influence of their own unsteady wake, which complicates the accurate prediction of lift and induces undesired noises and vibrations. Conventional techniques for measuring surface pressures in these instances, such as pressure taps or transducers, can be quite cumbersome to install on very thin leading edges and rotating surfaces. In addition, pressure information obtained by these conventional techniques is limited to point-wise measurements which may not capture subtle flow phenomena with sufficient spatial resolution.
Pressure-sensitive paint (PSP) is an optical measurement technique that is particularly suited for acquiring surface pressure maps. PSP consists of a special luminescent coating that responds to local wall pressure changes after being excited with a short-wavelength illumination source (typically LEDs or lasers). The local emitted light intensity is inversely proportional to the local surface pressure; thus, by recording the light distribution over the surface with a scientific-grade camera, the corresponding pressure field can be calculated. Each pixel on the camera chip essentially acts as a pressure transducer by sampling the paint luminescence – a major advantage over conventional techniques consisting of relatively sparse point-wise measurements.
Research Contributions
In our research group, we continue to make developments in improving the frequency response of
PSP to enable 
Unsteady PSP applied to a hemispherical turret in Mach 0.6 flow, visualizing motion of the shock wave near apex. Pressure normalized by stagnation value.time-resolved measurements of the pressure map in unsteady flows. Our research group has played a major role in developing PSP data acquisition techniques which effectively enable “point-and-shoot” sampling of unsteady surface pressure maps. Some examples of our past work have included applying our fast-responding PSP to transonic and low-speed flows with significant unsteady flow structures. In particular, we have applied our PSP to the investigation of unsteady transonic flow past a hemispherical turret in the Air Force Research Laboratory’s Trisonic Gasdynamics Facility. The sample data at left emphasize the utility of PSP in detecting the three-dimensional structure of the shock wave formed at the turret apex. Movie files created from the PSP images are extremely useful in visualizing the unsteady motion of the shock wave – an advantage over transducer arrays whose output is a series of electrical signals.
Our fast-responding PSP has been applied to a helicopter rotor blade in forward flight conditions where the surface pressure is a rapidly-varying function of time and angular position of the blade. Our group has led the way in developing image processing tools to correct the PSP data for camera artifacts such as image blurring due to surface movement. These contributions have the potential to change the way that helicopter loads testing in industry is conducted by using an all-optical system.
Instantaneous PSP data acquired on a rotating helicopter blade in forward flight conditions (local pressure normalized by atmospheric value).
Another significant aspect of our PSP research is to improve the accuracy of PSP measurement by minimizing errors caused by temperature effects and model vibrations. PSP is inherently sensitive to temperature, and the temperature induced errors can be large for common fast PSPs. A 2-color fast PSP has been developed, which emits a red pressure signal and a green temperature signal simultaneously. The green signal is only sensitive to temperature and it is used to correct for the errors in red pressure signal due to temperature change. Either a color camera or a dual monochrome camera system is required for this paint. The same idea can be applied to removed the errors caused be model movements and vibrations.
Pressure field created by a sonic jet impinging obliquely on a flat plate: (a) no temperature correction, with false low pressure region is a result of jet cooling; (b) with temperature correction. Units in kPa.
An in-house suite of tools for paint mixing and characterization are available in our laboratory, including a full chemical preparation bench and spray hood. We have the capability of performing static PSP calibrations in a controlled sample chamber, as well as dynamic calibrations using acoustic resonance and shock tubes.
Representative Publications
- Gregory, J.W., Asai, K., Kameda, M., Liu, T., and Sullivan, J.P., 2008, "A Review of Pressure-Sensitive Paint for High Speed and Unsteady Aerodynamics," Proceedings of the Institution of Mechanical Engineers, Part G, Journal of Aerospace Engineering, vol. 222, no. 2, pp. 249-290. doi: 10.1243/09544100JAERO243
- Fang, S., Long, S.R., Disotell, K.J., Gregory, J.W., Semmelmayer, F.C., and Guyton, R.W., 2011, “Comparison of Unsteady Pressure-Sensitive Paint Measurement Techniques,” AIAA Journal, vol. 50, no. 1, 209-222. doi: 10.2514/1.J051167
- Juliano, T.J., Kumar, P., Peng, D., Gregory, J.W., Crafton, J.W., Fonov, S., 2011, "Single-Shot, Lifetime-Based Pressure-Sensitive Paint for Rotating Blades," Measurement Science and Technology, vol. 22, no. 8, 085403. doi: 10.1088/0957-0233/22/8/085403