Tuft Visualization Research Articles

Aerodynamic and RSM Analysis of Wingsuit Stability

A Wingsuit is a Skydiving Jumpsuit that generates more lift for longer flights. This study examined the effects of side slip angles on a beginner wingsuit at 106 Reynolds number. Experimental tests were determined by using the length of the model scale at angles of attack ranging from 0° to 40° and sideslip angles of up to 20°. Force and moment coefficients were analyzed using variations in angles of attack and sideslip. Despite the absence of significant effects of sideslip angles on the lift and drag coefficients, side force and rolling/yawing moments were highly nonlinear. Flow structure visualization and numerical simulation show that surface stalls only occur on the lower side when slip angles are lower. In individual aviation sports, wingsuits are more advantageous when they have less sideslip. With Tuft visualization on the wingsuit model, the best aerodynamic coefficient under different flight conditions was determined by comparing the Response Surface Methodology performance under different flight conditions.

Experimental and numerical simulation of shock train characteristics in an isolator with incident shocks

To better comprehend the steady-state structure and dynamic characteristics of shock train flow fields under interactions with incident shocks, wind tunnel experiments investigating the shock train flow fields in a supersonic isolator with incident shocks were conducted. High-speed schlieren visualization with multiple knife edges, surface tuft visualization, and high-frequency wall pressure measurements were used to analyze the shock train flow fields in the primary and corner regions. The understanding of the flow mechanism was enhanced based on additional three-dimensional steady numerical simulations. Flow field asymmetry was observed when the shock train interacted with parallel incident shocks. The shock train leading edge comprised a longer shock, which traversed the entire channel, and a shorter shock, with flow separation extending downstream of the longer shock. The flow downstream of the shorter shock passed through a small separation bubble before reattachment. The flow field was symmetrical when the shock train interacted with crossed incident shocks. Stronger oscillations were observed in the symmetric shock train owing to the adverse pressure gradient created by the incident shock and the inconsistent flow parameters downstream of the shock. The downstream propagation of the pressure fluctuation caused by the oscillation of the shock train leading edge was revealed based on coherence and phase analyses. The flow separation initiated further upstream in the corner region than in the primary region of the shock train flow field, and the separation flow covered a larger proportion of the flow field. These differences were enhanced by increasing the incoming Mach number.

Experimental and CFD Investigation of Directional Stability of a Box-Wing Aircraft Concept

This study aimed to explore the directional stability issues of a previously studied light box-wing aircraft model with a pusher propeller engine in the fuselage aft section. Earlier configurations have included the use of fuselage together with a lifting system consisting of two wings joined together at their wingtips with vertical stabilizers. However, these side vertical surfaces failed to provide the aircraft with sufficient directional stability, thus prompting the quest in this study for novel solutions that would exclude the need for a fuselage extension and a typical fin. Solutions included the use of a ducted propeller and few configurations of small “fishtail” vertical fins, which formed part of the aft fuselage itself and coupled with vortex generators on the fuselage surface to improve their interference and heal flow separation at the fuselage aft cone. The results of wind tunnel testing were supported with CFD simulations to explain the flow behavior of each of the studied solutions. Tuft visualization and computed flow patterns allowed identification of the sources of the observed low efficiency in terms of directional stability of the fishtail against a simple idle duct without a propeller. A final configuration with a duct and a modified version of the fuselage fins was achieved that provides enough yaw stability margins for a safe flight.

Pressure and velocity measurements of air flow past a proposed generic Aircraft carrier geometry

Research on the air flow disturbances in the aircraft carrier environment has gained prominence in recent times. However, there is presently no representative carrier model analogous to the Simplified Frigate Shape (SFS) which is generic naval frigate for air flow investigation. In the present study, a Generic Aircraft Carrier (GAC) model is proposed, as a simplified, benchmark model for aerodynamic research. With the motivation to provide validation data for future CFD studies, baseline experimental data is generated in the wind tunnel, in terms of pressure distribution over the deck, for two variants, namely, a complete flat deck configuration with no island and secondly, with the island in the baseline position of the GAC. Effect of the island in modifying the flow is discussed by a comparison between the two variants. Particle Image Velocimetry (PIV) is employed to record velocity and turbulence levels in the GAC environment, highlighting regions of velocity deficits, and unsteady flow which may hinder the landing procedure of an approaching pilot. Comprehensive database of experimental data is presented as baseline data for future work and for validation of numerical models. Traditional tuft and smoke visualization studies are also conducted to provide corroboratory qualitative insights.

Experimental and theoretical investigations on the effect of a single leading-edge protuberance on airfoil performance

Geometrical modification of airfoils by leading-edge protuberances has attracted plenty of attention as a passive separation control method. In this research, the effect of a single leading-edge protuberance with different amplitudes on the performance of a two-dimensional airfoil (NACA 634-021) is investigated through wind tunnel experiment and theoretical analysis. Force measurement experiment shows that all the modified airfoils demonstrate a special two-step stall performance. When the first step of stall occurs, the lift coefficient of the modified airfoils drops to a medium value between the pre-stall and post-stall lift coefficients of the baseline airfoil and keeps nearly constant within a range of AOA (angle of attack). Tuft visualization reveals that during the first step of stall, the flow pattern on different sides of the protuberance is distinctly different, respectively, showing leading-edge separation and trailing-edge separation, which are compartmentalized by the attached flow originated from the peak of the protuberance. The flow control mechanism is contributed to the combined effect of the early occurrence of the leading-edge separation around the protuberance root, the inherent hysteretic performance, and the flow compartmentalization effect. A theoretical model based on the lifting line theory, which takes the amendment of the hysteresis and compartmentalization effects into consideration, is proposed to analyze the performance of the modified airfoils. The sectional and overall performances by theoretical analysis are highly consistent with the tendency of the experiment results. Based on the theoretical model, the spanwise circulation gradient induced by the protuberance is considered to be responsible for the earlier separation around the protuberance roots at pre-stall AOAs. Further investigation on the effect of the geometrical parameters of the protuberance reveals that the amplitude-to-wavelength ratio plays a dominant role on the performance of the modified airfoils. The special flow phenomenon induced by the single leading-edge protuberance discovered in this research and the proposed mechanism provide an essential point of view for fully understanding the combined mechanism of multiple protuberances.

Active Separation Control at the Pylon-Wing Junction of a Real-Scale Model

The effect of active flow control on local flow separation behind the installation location of an ultra-high-bypass-ratio nacelle is investigated in a real-scale experiment at the TsAGI T-101 wind tunnel. The investigated model represents a swept 2.5D section of the pylon-wing junction in landing configuration. A flow control system employing periodic excitation is integrated into the unprotected leading edge inboard of the pylon. Tuft visualization as well as pressure and force measurements are used to investigate the vortex-dominated base flow and the effect of active flow control. It is shown that the flow control effect is governed by the normalized parameters of flow control, the momentum coefficient, and the velocity ratio, and is largely independent of the free-stream Mach and Reynolds number within the investigated range. The influence of a variation of the momentum coefficient on the lift gain is investigated. At the highest Reynolds number of an application of active flow control with fully eliminates local separation and increases total lift by approximately 2% of the respective baseline value across a range of 5 deg in angle of attack.

Vortex Shedding and Frequency Lock in on Stand Still Wind Turbines—A Baseline Experiment

During the commissioning and stand-still cycles of wind turbines, the rotor is often stopped or even locked leaving the rotor blades at a standstill. When the blades are at a standstill, angles of attack on the blades can be very high, and it is therefore possible that they experience vortex-induced vibrations. This experiment and analysis helps to explain the different regimes of flow at very high angles of attack, particularly on moderately twisted and tapered blades. A single blade was tested at two different flow velocities at a range of angles of attack with flow tuft visualization and hotwire measurements of the wake. Hotwire wake measurements were able to show the gradual inception and ending of certain flow regimes. The power spectral densities of these measurements were normalized in terms of Strouhal number based on the projected chord to show that certain wake features have a relatively constant Strouhal number. The shedding frequency appears then to be relatively independent of chord taper and twist. Vortex generators (VGs) were tested but were found to have little influence in this case. Gurney flaps were found to modify the wake geometry, stall onset angles, and in some cases the shedding frequency.

Experimental Investigation of the Near Wall Flow Downstream of a Passenger Car Wheel Arch

The flow around and downstream of the front wheels of passenger cars is highly complex and characterized by flow structure interactions between the external flow, fluid exiting through the wheelhouse, flow from the engine bay and the underbody. In the present paper the near wall flow downstream of the front wheel house is analyzed, combining two traditional methods. A tuft visualization method is used to obtain the limiting streamline pattern and information about the near wall flow direction. Additionally, time resolved surface pressure measurements are used to study the pressure distribution and the standard deviation. The propagation of the occurring flow structures is investigated by cross correlations of the pressure signal and a spectral analysis provides the characteristic frequencies of the investigated flow. It is found that two main flow phenomena can be observed: one originates from flow exiting the upper wheelhouse and a second one resulting from a separation on the lower wheel house edge. The frequency spectrum reveals a dominant Strouhal number of 0.2. As the observed flow structures are attributed to the wheel-wheelhouse interaction, a closed wheelhouse configuration is also investigated and the results confirm that the fluctuations and observed flow structures are created by the flow interaction between the wheel, wheelhouse and the rotation of the wheel.

Periodic and aperiodic flow patterns around an airfoil with leading-edge protuberances

Recently leading-edge protuberances have attracted great attention as a passive method for separation control. In this paper, the effect of multiple leading-edge protuberances on the performance of a two-dimensional airfoil is investigated through experimental measurement of aerodynamic forces, surface tuft visualization, and numerical simulation. In contrast to the sharp stall of the baseline airfoil with large hysteresis effect during AOA (angle of attack) increasing and decreasing, the stall process of the modified airfoil with leading-edge protuberances is gentle and stable. Flow visualization revealed that the flow past each protuberance is periodic and symmetric at small AOAs. Streamwise vortices are generated on the shoulders of the protuberance, leading to a larger separation around the valley sections and a longer attachment along the peak sections. When some critical AOA is exceeded, aperiodic and asymmetric flow patterns occur on the protuberances at different spanwise positions, with leading-edge separation on some of the valley sections and non-stalled condition elsewhere. A combined mechanism, involving both the compartmentalization effect of the slender momentum-enhanced attached flows on the protuberance peaks and the downwash effect of the local stalled region with low circulation, is proposed to explain the generation of the aperiodic flow patterns. The influence of the number of protuberances is also investigated, which shows similar aperiodic flow patterns. The distance between the neighboring local stalled valley sections is found to be in the range of 4–7 times the protuberance wavelength. According to the proposed mechanism, it is speculated that the distance between the neighboring local stalled valley sections is inclined to increase with a smaller protuberance amplitude or at a larger AOA.

Two-Step Stall Characteristic of an Airfoil with a Single Leading-Edge Protuberance

An experimental investigation has been conducted to examine the influence of a single leading-edge protuberance on the performance of a baseline NACA airfoil. Static aerodynamic forces of the airfoils have been measured as the angle of attack increased and decreased. A hysteresis loop of the lift coefficient around the stall angle is observed for the baseline airfoil, with a sharp decline of the lift coefficient during angle-of-attack increase and a delayed recovery during angle-of-attack decrease. It was discovered for the first time that both the decline and the recovery of the lift coefficient of the modified airfoil with a single leading-edge protuberance included two steps. When the first step happened, the lift coefficient reached an intermediate value between the maximum value and the poststall value of the baseline airfoil, and it remained almost constant within a range of angles of attack. Surface tuft visualization revealed that, during the first step of stall, one side of the modified airfoil was stalled with leading-edge separation, whereas the other side remained in a nonstall condition. When the second step of stall happened, both sides of the modified airfoil were stalled, except for an attached flow on the protuberance peak. This one-sided stall phenomenon indicated that one important role of the leading-edge protuberance might be confining the local stall region from extending spanwise, which is similar to the effect of a wing fence.

Three-dimensional span effects of high-aspect ratio synthetic jet forcing for separation control on a low Reynolds number airfoil

The three-dimensional structure of the reattached flow caused by synthetic jet actuation on an airfoil was investigated using surface flow visualization. Without active control, the flow was stalled with laminar boundary layer separation occurring near the leading edge. Tuft and oil visualization showed the shape and spanwise extent of the attached flow due to a finite span synthetic jet where the effect of excitation frequency and blowing ratio was the focus. For all excitation frequencies tested, a similar contraction of the spanwise extent of the attached flow towards the trailing edge was observed due to edge effects of the finite span jet. Increasing the blowing ratio was found to decrease the amount by which the attached flow contracted.

Non-intrusive detection of rotating stall in pump-turbines

When operated far from their optimum conditions, pump-turbines may exhibit strong hydrodynamic instabilities, often called rotating stall, which lead to substantial increase of vibration and risk of mechanical failure. In the present study, we have investigated the flow filed in a model of radial pump-turbine with the help of tuft visualization, wall pressure measurement and structure-borne noise monitoring. As the rotation speed is increased, the machine is brought from its optimum operation to runaway with zero torque on the shaft. The runaway operation is characterized by a significant increase of pressure fluctuation at the rotor–stator interaction frequency. As the speed is further increased, the flow exhibits sub-synchronous instability, which rotates at 70% of the rotation frequency. Tuft visualization clearly shows that, as the instability evolves, the flow in a given distributor channel suddenly stalls and switches to reverse pumping mode in periodic way. We have also investigated the monitoring of the rotating stall with the help of vibration signals. A specific signal processing method, based on amplitude demodulation, was developed. The use of 2 accelerometers allows for the identification of the optimum carrier frequency by computing the cyclic coherence of vibration signals. This non-intrusive method is proved to be efficient in detecting the rotating stall instability and the number of stall cells. We strongly believe that it could be implemented in full scale pump-turbines.

Effects of Sideslip on the Aerodynamics of Low-Aspect-Ratio Low-Reynolds-Number Wings

The growing interest in micro aerial vehicles has brought attention to the need for an improved understanding of the aerodynamics of low-aspect-ratio wings at lowReynolds numbers. In this study, flat plate wings with rectangular and tapered planforms were fabricated with aspect ratios of 0.75, 1, 1.5, and 3, and the aerodynamic loading was measured at Reynolds numbers between 5 10 and 1 10. Surface tuft visualization was used to observe the interactions between the tip vortices and the leading-edge vortex. The tests were initially conducted at a sideslip angle of 0 and were then repeated for 10, 20, and 35 with and without winglets. Measurements made with a sixcomponent force balance showed that a decrease in aspect ratio caused an increase in stall and CLmax due to the nonlinear lift induced by the interacting flow on the upper wing surface. In addition, the detachment of tip vortices after stall leads to a sudden decrease in drag coefficient as the magnitude of the induced drag drops significantly. At increasing sideslip angles, the effects of the crossflow still contribute to an increase in lift but significantly reduce the pitching moment about the quarter-chord, thus decreasing the wing’s ability to recover from angle-of-attack perturbations. These results show that, while the effects of tip vortices and the leading-edge vortex complicate the flowfield around a low-aspect-ratio wing, particularly at increased sideslip angles, their impact tends to improve the aerodynamic performance.

Mean Velocity Profile in Recovery Region of Pure Poiseuille, Pure Couette and Couette-Poiseuille Type Turbulent Flows over a Backward-Facing Step

Mean velocity profiles in velocity recovery region of pure Poiseuille, pure Couette and Couette-Poiseuille type turbulent channel flows over a backward-facing step are measured and investigated experimentally. The flows are turbulent channel flows between moving and step-side stationary walls with different non-dimensional upstream pressure gradients β0 (flow type parameter). The channel expansion ratio is 1.50. The upstream flows are fully developed pure Poiseuille (β0=-1), pure Couette (β0=0) and Couette-Poiseuille (β0=-0.5) turbulence plane channel flows. The reattachment region is specified by tuft visualization method. The central point of the reattachment region is less dependent on β0. Velocity profiles are measured by hot-wire anemometry for the flows at Reτ0=300, where Reτ0 is the Reynolds number based on the friction velocity and channel half width at the upstream flow. The logarithmic law of velocity is observed even for sections just after the reattachment region. The additive constant appeared in the logarithmic law decreases with increase of β0. The half power law is formed at the beginning of the recovery region for the flows with β0=-0.5 and 0, while it is not observed for the flow with β0=-1.

Rotor Configuration Effects on the Performance of a HAWT With Tip-Mounted Mie-Type Vanes

In this paper, the relationships between the power augmentation of a HAWT due to Mie-type tip vane application and other factors influencing the efficiency of a wind turbine such as the blade aspect ratio, number of blades, blade pitch angle and Reynolds number are investigated. Experiments were carried out in a wind tunnel with an open, 3.6 m diameter, test section. Rotor models included two- and three-blade upwind turbines with four sets of blades with different aspect ratios. With the rotor blades tested, a maximum power augmentation of about 14.5% was achieved due to Mie vane application. The relationships between power augmentation due to the Mie vane and the above factors are investigated. It is found that the application of a tip-mounted Mie vane results in a larger increase in maximum power coefficient for rotors with smaller aspect ratios and for lower wind speeds. Surface oil film and surface tuft visualization methods were used to detect the flow patterns at the blade tip. Addition of the Mie vanes causes significant changes in flow behavior near the blade tip, resulting in additional blade lift.