This work is focused on the aerodynamics for a proprietary laminar flow airfoil for Unmanned Aerial Vehicle (UAV) applications. The two main focuses are (1) aerodynamic performance at Reynolds number on the order of 10,000, (2) the effect of a conventional hot-wire probe on laminar separation bubbles. For aerodynamic performance, pressure and wake velocity distributions were measured at Re = 40,000 and 60,000 for a range of angles of attack. The airfoil performed poorly for these Reynolds numbers due to laminar boundary layer separation. 2-D boundary layer trips significantly improved the lift-to-drag ratio. For probe effects, three Reynolds numbers were investigated (Re = 100,000, 150,000, and 200,000), with three angles of attack for each. Pressure and surface shear distributions were measured. Flow upstream of the probe tip was not affected. Transition was promoted downstream due to the additional disturbances in the separated shear layer.

This work is focused on the aerodynamics for a proprietary laminar flow airfoil for Unmanned Aerial Vehicle (UAV) applications. The two main focuses are (1) aerodynamic performance at Reynolds number on the order of 10,000, (2) the effect of a conventional hot-wire probe on laminar separation bubbles. For aerodynamic performance, pressure and wake velocity distributions were measured at Re = 40,000 and 60,000 for a range of angles of attack. The airfoil performed poorly for these Reynolds numbers due to laminar boundary layer separation. 2-D boundary layer trips significantly improved the lift-to-drag ratio. For probe effects, three Reynolds numbers were investigated (Re = 100,000, 150,000, and 200,000), with three angles of attack for each. Pressure and surface shear distributions were measured. Flow upstream of the probe tip was not affected. Transition was promoted downstream due to the additional disturbances in the separated shear layer.

The later stages of separated shear layer transition within separation bubbles developing over a NACA0018 airfoil operating at a chord Reynolds number of 105 and at angles of attack of 0, 5, 8, and 10 degrees were investigated experimentally in a wind tunnel. Several experimental tools, including a rake of six boundary-layer hot-wire anemometers, were used to perform measurements over the model. Novel high-speed flow visualization performed with a smoke-wire placed within the separated shear layer showed that roll-up vortices are shed within separation bubbles forming on the suction side of the airfoil. The structures were found to convect downstream and eventually break down during laminar-to-turbulent transition. Top view visualizations revealed that, at angles of attack of 0, 5, and 8 degrees, roll-up vortices form coherently across the span and undergo significant spanwise deformations prior to breaking down. At angles of attack of 5 and 8 degrees, rows of streamwise-oriented structures were observed to form during vortex breakdown. Statistics regarding the formation and development of shear layer roll-up vortices were extracted from high-speed flow visualization sequences and compared to the results of boundary layer measurements. It was found that, on the average, roll-up vortices form following the initial exponential growth of unstable disturbances within the separated shear layer and initiate the later stages of transition. The onset of these nonlinear stages was found to occur when the amplitude of velocity disturbances reached approximately 10% of the free-stream velocity. The rate of vortex shedding was found to fall within the frequency band of the unstable disturbances and lie near the central frequency of this band. The formation of vortices has been linked to the generation of harmonics of these unstable disturbances in velocity signals acquired ahead of mean transition. Once shed, vortices were found to drift at speeds between 33% and 44% of the edge velocity. Vortex merging at an angle of attack of 5?? was investigated. It was found that the majority of roll-up vortices proceed to merge with either one or two other vortices. Vortex merging between two and three vortices was found to occur periodically in a process similar to vortex merging in plane mixing layers undergoing subharmonic forcing of the most amplified disturbance. The flapping motion of the separated shear layer was investigated by performing a cross-correlation analysis on the high-speed flow visualization sequences to extract vertical displacement signals of the smoke within the shear layer. The frequency of flapping was found to correspond to the unstable disturbance band. At an angle of attack of 5??, it was found that the separated shear layer has a low-frequency component of flapping that matches a strong peak in velocity and surface pressure spectra that lies outside the unstable disturbance frequency band. The spanwise development of disturbances was assessed in the aft portion of the separation bubbles by performing a cross-correlation analysis on signals acquired simultaneously across the span with the rake of hot-wires. The spanwise correlations between signals was found to be well-correlated ahead of shear layer roll-up, after which disturbances became rapidly uncorrelated ahead of mean reattachment. These results were found to be linked to the coherent roll-up and subsequent breakdown of roll-up vortices.

The accuracy of existing methods for estimating the drag based on experimental flow field measurements were assessed for two-dimensional bodies. The effects of control volume boundary placement and inherent simplifying assumptions were also investigated. Wind tunnel experiments were performed on a circular cylinder operating at a Reynolds number of 8,000 and 20,000, and on a NACA 0018 airfoil operating at a chord Reynolds number of 100,000 for three angles of attack (α), specifically, 5◦, 10◦, and 15◦. The circular cylinder experiments fall within the the shear layer transition flow regime. Airfoil investigations span both types of flow development common to low Reynolds number airfoil operation. For α = 5◦ and 10◦, a separation bubble forms on the upper surface of the airfoil, while, for α = 15◦, the flow separates without reattaching, resulting in a stalled airfoil. Wake velocity and pressure measurements were performed at several downstream locations to investigate the impact of control volume boundary placement. Wake profiles were measured between 3 and 40 diameters downstream from the circular cylinder axis and between 1 and 4.5 chord lengths from the trailing edge of the airfoil. In addition to wake profiles, the outer flow velocity variation was quantified to investigate the appropriate location to measure freestream flow characteristics in a test section with streamwise-varying outer flow conditions. The results show that drag estimates are strongly dependent on the streamwise position of the measured wake profile for all methods investigated. Drag estimates improved, and streamwise variation decreased, with increasing streamwise position of the flow measurements. For the pressure based method examined, wake measurements should be taken at least 10 times the projected model height downstream of the model. In the case of the circular cylinder, this is equivalent to 10 diameters and, for the airfoil investigated, it is approximately 1 chord length from the trailing edge. For the methods relying on velocity measurements, acceptable estimates of drag were possible when based on measurements taken at least 30 projected heights downstream, i.e., 30 diameters for the circular cylinder and 3 chord lengths for the airfoil model investigated. The findings highlight the importance of providing a detailed description of the methodology and experimental implementation for drag estimates based on flow field measurements. Finally the study offers guidelines for implementing momentum integral based drag calculations in future investigations.

Shear layer development over a NACA 0018 airfoil at a chord Reynolds number of 100,000 was investigated experimentally. The effects of experimental setup and analysis tools on the results were also examined. The sensitivity of linear stability predictions for measured separated shear layer velocity profiles to both the analysis approach and experimental data scatter was evaluated. Analysis approaches that are relatively insensitive to experimental data scatter were identified. Stability predictions were shown to be more sensitive to the analysis approach than to experimental data scatter, with differences in the predicted maximum disturbance growth rate and corresponding frequency of approximately 35% between approaches. A parametric study on the effects of experimental setup on low Reynolds number airfoil experiments was completed. It was found that measured lift forces and vortex shedding frequencies were affected by the end plate configuration. It was concluded that the ratio of end plate spacing to projected model height should be at least seven, consistent with the guideline for circular cylinders. Measurements before and after test section wall streamlining revealed errors in lift coefficients due to blockage as high as 9% and errors in the wake vortex shedding frequency of 3.5%. Shear layer development over the model was investigated in detail. Flow visualization images linked an observed asymmetry in wake velocity profiles to pronounced vortex roll-up below the wake centerline. Linear stability predictions based on the mean hot-wire profiles were found to agree with measured disturbance growth rates, wave numbers, and streamwise velocity fluctuation profiles. Embedded surface pressure sensors were shown to provide reasonable estimates of disturbance growth rate, wave number, and convection speed for conditions at which a separation bubble formed on the airfoil surface. Convection speeds of between 30 and 50% of the edge velocity were measured, consistent with phase speed estimates from linear stability theory.

Shear layer development over a NACA 0018 airfoil at a chord Reynolds number of 100,000 was investigated experimentally. The effects of experimental setup and analysis tools on the results were also examined. The sensitivity of linear stability predictions for measured separated shear layer velocity profiles to both the analysis approach and experimental data scatter was evaluated. Analysis approaches that are relatively insensitive to experimental data scatter were identified. Stability predictions were shown to be more sensitive to the analysis approach than to experimental data scatter, with differences in the predicted maximum disturbance growth rate and corresponding frequency of approximately 35% between approaches. A parametric study on the effects of experimental setup on low Reynolds number airfoil experiments was completed. It was found that measured lift forces and vortex shedding frequencies were affected by the end plate configuration. It was concluded that the ratio of end plate spacing to projected model height should be at least seven, consistent with the guideline for circular cylinders. Measurements before and after test section wall streamlining revealed errors in lift coefficients due to blockage as high as 9% and errors in the wake vortex shedding frequency of 3.5%. Shear layer development over the model was investigated in detail. Flow visualization images linked an observed asymmetry in wake velocity profiles to pronounced vortex roll-up below the wake centerline. Linear stability predictions based on the mean hot-wire profiles were found to agree with measured disturbance growth rates, wave numbers, and streamwise velocity fluctuation profiles. Embedded surface pressure sensors were shown to provide reasonable estimates of disturbance growth rate, wave number, and convection speed for conditions at which a separation bubble formed on the airfoil surface. Convection speeds of between 30 and 50% of the edge velocity were measured, consistent with phase speed estimates from linear stability theory.