Surface Oil Flow Visualization Research Articles

Fluid-Structure Interactions of a Fin over a Compliant Panel in Supersonic Flow

The fluid-structure interactions of a blunt fin mounted over a compliant panel in Mach 2 flow were investigated. The test geometry consisted of a thin, flexible panel fixed on all edges with a blunt fin with a hemispherically rounded leading edge mounted in a manner that the leading edge of the fin overhung the panel. The blunt fin was also pivoted to several angles of attack with respect to the oncoming flow. This allowed for the study of the dynamic system created by the interactions between the compliant panel and the blunt fin geometries. Both rigid and compliant panel models were used in order to provide a comparison between the flow with and without the dynamics of the compliant panel. High-speed schlieren and surface oil flow visualizations showed that the time-averaged flow remained almost entirely unchanged between the rigid and compliant panels. The instantaneous flow fields, however, showed a highly dynamic system where the compliant panel induced oscillatory shock waves that were both stationary and freely moving. Simultaneous high-speed schlieren visualization and stereo digital-image correlation showed a strong linkage between the motion of the shock waves and the deformation of the panel. This was demonstrated through frequency analysis and modal decomposition of the panel deformation and shock motions.

Experimental Optimization of the SG6043 Airfoil for Horizontal Axis Wind Turbines Using Schmitz Equations

This study presents the experimental optimization of the SG6043 airfoil for horizontal axis wind turbines (HAWTs) using the Schmitz equation, focusing on enhancing power output and elucidating the surface flow structure. Two blade models, M1 (conventional) and M2 (optimized), were designed and tested at rotational speeds of 400 rpm and 600 rpm across a range of tip speed ratios (TSR). The M2 model, optimized using Schmitz equations, demonstrated significantly improved performance compared to the M1 model at both rotational speeds. At 400 rpm, the maximum power coefficient (CP) for M1 was 0.274, while M2 reached 0.419, indicating a 52.91% improvement. At 600 rpm, M1 achieved a maximum CP of 0.293, whereas M2 attained 0.458, representing a 56.31% enhancement. The M2 model also showed superior performance at higher TSRs, with the highest percentage increase in CP recorded at 4.9 TSR, reaching 574.54%. Additionally, dynamic surface oil-flow visualization experiments were conducted to examine flow behavior on the blade surfaces. Results indicated better flow attachment in the M2 blade due to its optimized twist angle and chord length, particularly in the mid-section, leading to delayed flow separation. The reattachment observed on the suction side of the M2 model, following the laminar separation bubble (LSB), which was absent in the M1, contributed to its higher aerodynamic efficiency and overall power performance. These findings confirm that the optimized SG6043 airfoil design, guided by Schmitz equations, offers significant improvements in HAWT performance, particularly under varying operational conditions.

Experimental study on flow field and performance of bionic-wavy leading edge in an axial compressor with positive bowed blades

To enhance the aerodynamic performance of compressors in advanced aeroengines, a compound flow control method combining positively bowed blades and bionic-wavy leading edges is proposed for improving the aerodynamic performance of compressor cascades with controlled diffusion airfoils. This study verified the effectiveness of the compound control method through low-speed wind tunnel experiments using five-hole probe measurements and surface oil-flow visualization techniques. Additionally, the flow field structure was analyzed, and vortex models were established to thoroughly discuss the mechanism of the compound flow control method. The results show that within the incidence angle range of 0°–4° studied in this paper, the composite control method achieved significantly effective control, with a maximum reduction in overall total pressure loss of 25.8% compared to straight blade cascades. Three vortex models were established. The positive bowed blade cascade induced a complex vortex structure in the concentrated shedding vortex region, increasing losses in the concentrated shedding vortex (CSV) region but reducing profile losses. The coupled method further reduced profile losses and optimized the flow field in the CSV region. This study not only validates the feasibility of the compound method but also provides guidance for applying flow control methods to bowed blade cascades.

Wind Tunnel Tests of a Thick Wind Turbine Airfoil

ABSTRACTThis article reports about a wind‐tunnel experiment carried out in the ONERA F2 low‐speed wind tunnel on a model of the DU 97‐W‐300Mod airfoil designed for wind turbine application. The wind tunnel, the airfoil model, and experimental techniques used are presented, with special emphasis on the data processing and corrections required to derive airfoil forces and pressure distribution. To better document the flow physics at play, the results are illustrated by infrared thermography and surface oil flow visualization. The test allowed investigating Reynolds number effects between 1 and 3.8 millions. To ameliorate the understanding of the benefits and limitations of such airfoil testing, one section is devoted to the comparison of present results with previous experiments in other wind tunnels. Some of the difficulties arising in airfoil testing are evidenced and discussed to contribute to the improvement of test methods.

Experimental study of flow control on tip leakage flow in variable geometry linear turbine cascades with different pivot layouts

Variable geometry turbines are essential for adjusting operational conditions in industrial gas turbines and variable cycle engines. These adjustments necessitate partial gaps at both ends of the variable guide vanes to alter the turning angle, consequently introducing an aerodynamic performance penalty. Moreover, the pivot layout profoundly influences aerodynamic losses. Research on turbine cascades that considers various partial gap layouts is limited, particularly in terms of experimental studies, which are rarely conducted. This study aims to diminish aerodynamic losses and augment the efficiency of gas turbines by examining the impact of pivot layouts on partial gap clearance and secondary flow. It further investigates the effectiveness of flow control strategies at the blade tip across different pivot configurations within a variable geometry turbine cascade, utilizing pneumatic probe scanning and surface oil flow visualization techniques. The results reveal that employing a cavity at the tip can significantly reduce aerodynamic losses in schemes both with and without a pivot, achieving maximum loss reductions of 15.8% and 3.7%, respectively. Additionally, a narrower squealer width can further decrease these losses. However, with a pivot located at the tip, the resulting separation flow and wake vortex become predominant sources of losses. The presence of the pivot weakens the tip leakage flow rate and the intensity of the tip leakage vortex (TLV), thus diminishing the effectiveness of cavity tip flow control. The cavity moderates TLV and enhances the interaction between TLV and the wake vortex, leading to increased aerodynamic losses.

On incident shock wave/boundary layer interactions under the influence of expansion corner

Cowl-induced incident shock wave/boundary layer interactions (ISWBLIs) under the influence of shoulder expansion represent one of the dominant phenomena in supersonic inlets. To provide a more comprehensive understanding of how an expansion corner affects the ISWBLI, a detailed experimental and analytical study is performed in a Mach 2.73 flow in this work. Pressure measurement, schlieren photography and surface oil-flow visualisation are used to record flow features, including the pressure distribution, separation extent and surface-flow topological structures. Our results reveal three types of ISWBLIs influenced by the expansion corner. When the shock intensity is weak, the separation is small scale with the expansion waves emanating from the expansion corner. This is the first type of expansion-corner-affected ISWBLI (EC-ISWBLI). When the incident shock wave is strong, large-scale separation occurs, accompanied by the disappearance of expansion waves, forming the second type of EC-SWBLI. The expansion corner induces a ‘lock-in’ effect in which the separation onset is consistently locked near the expansion corner regardless of the incident shock intensity and impingement position. The third type of EC-ISWBLI occurs when the shock is sufficiently strong and the impingement point is close to the expansion corner. In this interaction, the ‘lock-in’ effect ceases to manifest. Moreover, a shock polar-incorporating inviscid model is employed to elucidate the shock patterns. Two criteria are established by combining free interaction theory with this model. The first criterion provides valuable insights into the evolution of separations with a minimal overall pressure rise and the second criterion determines the threshold for the occurrence of the ‘lock-in’ effect.

Propeller Ceiling Effect in Forward Flight

The study extends the investigation of the fixed-pitch small-scale propeller in the ceiling effect to forward flight conditions at different propeller incidence angles. Force-based experiments, phase-locked particle image velocimetry (PIV), and surface oil flow visualization were conducted on two APC propellers at the University of Dayton Low Speed Wind Tunnel. For propellers in edgewise flight, the power required at constant thrust decreases at small advance ratios for each h/D and then increases with a further increase in the advance ratio. Tilting the propeller forward reduces the ceiling effect benefits in both thrust and power, particularly at higher advance ratios. Performance similarity in the propeller ceiling effect at different h/D is observed, and a performance prediction method is proposed. Phase-locked PIV showed an increase in the propeller inflow angle in the ceiling effect at small advance ratios, resulting in higher thrust generation. This effect reduces with an increase in the advance ratio due to minimized interactions with the ceiling plate. Hence, the measured propeller in-ceiling-effect (ICE) propeller performance cannot represent the propeller ICE performance at a higher forward flight speed. Additionally, PIV and surface flow visualization indicated the presence of a stagnation point on the ceiling plate near the trailing side of the propeller disk at higher advance ratios, leading to a reduction in propeller thrust generation.

Validation of a RANS Turbulence Model for a S833 Wind Turbine Airfoil With a Trailing Edge Flap Using Oil Visualization and Pressure Taps

Abstract The aerodynamics of a small wind turbine blade was captured using a γ–Reθ k–ω shear stress transport transitional turbulence model tuned with production limiter coefficients at a Reynolds number of 1.70×105. The computational fluid dynamics simulations were validated against wind tunnel experiments that included airfoil pressure tap measurements and surface oil flow visualization (SOFV) to capture the flow field. The uniqueness of this blade included a trailing edge flap that was 20% of the chord controlled using a servomotor. The test matrix included angles of attack (AOA) between 1 deg and 7 deg with flap angles of 10 deg in the upward and downward position. Two locations were always observed on the airfoil: a leading edge region of high shear and a midsection of flow separation. Within the flow separation section, two distinct regions existed: a complete detachment of flow from the airfoil surface creating a stagnation region which was followed by a reverse flow region. A third location of flow reattachment near the trailing edge was observed for all cases excluding a downward angled trailing edge flap. The utilization of the flap resulted in changes to the size of the separation zone and the movement of the separation zone along the chord. The numerical skin friction coefficient, oil residue profiles from the SOFV, and pressure tap measurements all showed onset of separation locations on the chord within 10%. The computational fluid dynamics model also predicted the coefficient of pressure across the chord of the airfoil within 10% in comparison to the experimental measurements.

Experimental investigation of secondary flow losses in a variable geometry linear turbine cascade with winglet tip

Variable geometry turbines (VGT) hold a crucial role in gas turbines and variable cycle engines (VCE) by allowing adjustments to different operational conditions. In order to realize the adjustable function of the VGT stator blade, there will be partial gap structure at both ends of the stator blade, which will lead to partial leakage flow and serious secondary flow loss at the blade end that cannot be ignored, especially in a large angle adjustment range. At present, there are few researches on turbine cascades considering the real partial gap structure and winglet tip, and the numerical simulation calculation is mainly used, while the in-depth study using wind tunnel test is rarely seen. Therefore, this study attempts to analyze the internal flow characteristics of variable geometry turbine cascade by using five-hole pneumatic probe scanning and surface oil flow visualization, and discusses the mechanism of controlling partial clearance leakage flow and reducing secondary flow loss in the cascade by using winglet tip flow control technology. The winglet offset distance and the change of the relationship between the pivot and the clearance position on the flow field and flow loss is studied experimentally for the first time. The results show that the influence mechanism of winglet tip applied in VGT partial gap is quite different from that of typical blade tip clearance application, and more consideration should be given to the influence of pivot structure. Reasonable parameter selection can achieve a maximum loss reduction of 13.4 % and 20.2 % with and without pivot, respectively, while the unreasonable flow control scheme may lead to the loss or even increase with pivot. The interaction between the new secondary flow structure induced by pivot in the gap and other gap leakage flows is analyzed to explore the new mechanism of using winglet tip to reduce the loss of gap leakage flow.

Effects of herringbone riblets on shock-wave/turbulent boundary-layer interactions

In the experiment reported in this paper, a spanwise array of herringbone riblets strips is applied upstream of the interaction region between an impinging shockwave and a turbulent boundary layer developing along a flat plate boundary layer at a freestream Mach number of 1.85. High-speed schlieren photography, surface oil flow visualization and PSP are used to examine the effect of herringbone riblets on the characteristics of the shockwave-boundary layer interactions. It shows that despite that the height of the riblets is less than 2 % of the local thickness of the undisturbed boundary layer, a profound modification to the shockwave-boundary layer interaction zone is observed, and the overall size of the separation bubble is decreased. The observed control effects are attributed to the large-scale secondary flow structures produced by the directional riblets. To the best knowledge of the author, this is the first study evaluating the control effect of herringbone riblets on shock wave and turbulent boundary layer interactions in supersonic flow using the experimental method.

An experimental investigation on the aerodynamic characteristics and vortex dynamics of a flying wing

AbstractIn this paper, we present a detailed experimental investigation mainly on the vortical flow fields and the associated vortex breakdown phenomena over a non-slender flying wing (sweep angle, ${\rm{\Lambda }}$ = 53°). In the process, the aerodynamic coefficients were also determined using a six-component force balance. Surface oil flow visualisation, surface pressure measurements and particle image velocimetry (PIV) measurements, in various crossflow planes and in a longitudinal plane passing through the leading-edge vortex core, were carried out at various Reynolds numbers to understand the flow field over the non-slender flying wing. Aerodynamic characteristics of the flying wing show local peaks and valleys in the pitching moment coefficient. The surface flow visualisation reveals that the nonlinearity of the pitching moment curve is due to the complex nature of vortical flow structures. The flow visualisation also demonstrates the presence of a wave-like surface pattern, and its size is found to reduce with increasing Reynolds numbers. The present PIV measurements confirm that this wave-like surface pattern is associated with vortex breakdown phenomena. These measurements also reveal that the vortex breakdown has not reached the apex of the wing, even at post-stall angle-of-attack. For pre-stall ( $\alpha $ = 20°) flow regimes, it is observed that the location of the vortex breakdown moves downstream as the Reynolds number increases, but this influence is minimised at near-stall ( $\alpha $ = 25°) and post-stall ( $\alpha $ = 30°) flow regimes. Reconstructed velocity field using the first 10 dominant proper orthogonal decomposition (POD) modes reveals that the nature of the vortex breakdown over the flying wing is a spiral-type vortex breakdown.

Aerodynamic Characterization of a Generic High-Speed Projectile Configuration

A combined experimental and numerical study was conducted on a generic projectile configuration with low-aspect-ratio fins. The main objectives of this study were to characterize the aerodynamic behavior and validate numerical simulations for a range of Mach numbers (0.4–4) to facilitate an understanding of major flow features such as forebody and fin-generated shock waves, crossflow shear layer vortex, and fin vortex interactions. Measurements included forces and moments, surface oil flow visualization, and high-speed shadowgraph imaging. Numerical simulations were performed using the CFD++ solver. The results showed an excellent match between the experimental and numerical force and moment data. Pressure contours obtained using numerical simulations were integrated to obtain the contributions of individual components toward the total normal force on the body. Flow visualization results show a few complex and interesting flow features, such as shear layer roll-up, crossflow and fin tip vortex interactions, and shock-wave–boundary-layer interactions. The effects of vortex strength and location were analyzed to determine their contributions to the overall forces on the model. The database generated will be very useful for further validation of the numerical tool and a better understanding of vortex-dominated supersonic flows.

Effect of aero-shaped vortex generators on NACA 4415 airfoil

A parametric study was carried out to investigate the effect of aero-shaped vortex generators (AsVGs) on NACA 4415 airfoil by comparing conventional vortex generators (CVGs). Force and flow visualization experiments were carried out for the CVGs and AsVGs at Re = 1.4 × 105. Six different configurations were selected for the positioning of the CVGs to investigate the best one. The best configuration was applied to the Aero-shaped VGs type, which is generated with the S1223 airfoil model. The results indicated that even if AsVGs did not show better performance to delay the stall angle when compared with the CVGs, in some positions they produced the same stall performance as the CVGs. Moreover, aero-shaped VGs increased the lift coefficient (CL) more than the conventional ones in the pre-stall region. At x/c = 0.1, CL and CD improve by up to 107.7% and 23.39% compared to baseline, respectively. At α = 18° and Re = 0.8 × 105, AeM5 at x/c = 0.1 increased the CL by 102.45% compared to the baseline model. While the CVGs increased the lift coefficient (CL) and postponed the stall (α) angle up to 4° according to the baseline model, the CVGs also caused an overall increase in the drag coefficient (CD). AsVGs decreased the drag coefficient (CD) more than conventional ones. Lift to drag ratio (L/D) of AsVGs is higher than that of the CVGs. The results of the surface oil flow visualization showed that both CVGs and AsVGs segmented the laminar separation bubble (LSB) and formed attached flow zones on the airfoil surface. As a result, AsVGs can be used as an alternative to CVGs in flow control applications.

Wall Pressure Fluctuations on the Surface of a Fairing in the Shape of a Semi-Ellipsoid and in its Vicinity

The wall pressure fluctuation fields on fairing surfaces and in the surrounding turbulent boundary layer are experimentally investigated. The fairings were in the shape of semi-ellipsoids and were mounted on the wall of a subsonic low-noise wind tunnel. Their heights amounted to 25% of the oncoming boundary layer thickness. The main physical features of the flows under consideration are determined by means of the surface oil-flow visualization. The fluctuating wall pressure field is compared with the flow pattern. It is shown that the greatest pressure fluctuation levels are recorded in the nose region of the fairing surface. It is established that lengthening the model leads to a considerable reduction of the pressure fluctuation strength in the region of flow separation from the fairing surface.

Investigation of aerodynamic performance of Clark-Y airfoil with more realistic tubercle model and internal slots

An experimental study is conducted to research the combined effect of leading-edge (LE) tubercles and internal slots on the Clark-Y airfoil at Reynolds numbers (Re) of 5.0×104, 7.5×104, and 1.0×105 for angles of attack ranging from 0° to 30°. The Clark-Y is a cambered airfoil that performs well at low and medium Reynolds numbers, having a maximum thickness of 11.7% at 28% of the chord. Five different test models are the subjects of experiments, one of which is the baseline. The other test models are the Wavy model, which has LE tubercles, and the Wavy with three different slot positions: the Wavy-LES (leading-edge slot), the Wavy-MS (middle slot), and the Wavy-TES (trailing-edge slot). The LE tubercle geometry is formed by the sum of two sinusoidal wave functions with a realistic approach. In the present work, experimental studies are performed through force measurements, and detailed information on a 3-dimensional flow field is obtained by a surface-oil flow visualization technique. Force measurements indicate that the Wavy model delays stall and exhibits smoother stall behavior compared to the baseline. Moreover, the findings show that the lift coefficient (CL) of the Wavy model deteriorates in the pre-stall region with the inclusion of the internal slots but improves significantly in the post-stall region. Maximum improvement in CL in the post-stall region was achieved by 60% in Wavy-LES at Re of 5.0×104 as compared to the baseline. At Re of 5.0×104, Wavy and Wavy-MS present a better lift-to-drag ratio (CL/CD) than the other models in the pre-stall region, whereas the baseline is the best at Re of 1.0×105. The best CL/CD is achieved by Wavy-LES in the post-stall region, regardless of the Reynolds numbers. The variation of flow characteristics relevant to aerodynamic performance is revealed by surface oil flow visualization for all tested models.

Supersonic Flow Separation in Front of a Forward-Facing Step

A detailed experimental study has been done to understand the Shock Wave-Boundary Layer Interaction (SWBLI) over a Forward-Facing Step (FFS) of height (h) in a Mach 2.5 flow using Particle Image Velocimetry (PIV). In addition to PIV, unsteady wall pressure measurements, high-speed schlieren and surface oil flow visualization have been studied. Flow separation using surface oil flow visualization was observed around 4.1 step heights upstream of the step face. Shock motions extracted from the schlieren image sequence shows a peak frequency of approximately 1000 Hz, which corresponds to the dominant (low) frequency at which the shock oscillates; this being two orders in magnitude smaller than the characteristics frequency of the incoming boundary layer. Also, spectra of the unsteady wall pressure data shows that the shock oscillates with a peak frequency of approximately 1000 Hz. The effect of shock location (xs) from instantaneous velocity PIV fields in the cross-stream plane due to an upstream influence parameter (line-averaged boundary layer streamwise velocity, ul) and a downstream parameter (separation bubble or recirculation region area, A) is studied. Instantaneous cross-stream PIV measurements show significant variations of the separation shock location, incoming boundary layer velocity fluctuations and recirculation region. Besides, conditionally averaged velocity field based on incoming boundary layer streamwise velocity fluctuations and separation bubble area in the cross-stream plane shows that the shock location (xs) is correlated to the size of the separation bubble and not to boundary layer velocity fluctuations.

Flow Field Investigation on a Double Delta Wing

In this paper flow field characteristics of double delta wing have been studied both experimentally and numerically at different angles of attack ranging from 0° to 30° at a Reynolds number of 2 x 10<sup>5</sup> based on the root chord length of the wing model. Surface oil flow visualization has been studied on the leeward surface of a sharp edged single beveled 76°/40° cropped double delta wing at subsonic wind tunnel facility. The flow physics is further validated with the help of surface pressure measurement in the context of surface flow pattern. Subsequently CFD (Computational Fluid Dynamics) analysis is carried out using ANSYS Fluent. Comparisons of both qualitative as well as quantitative measurement with experimental techniques and computational work have been carried out. For computational work, the computational domain is meshed with unstructured grids and SA-turbulence model is used.

Flow Structures and Aerodynamic Behavior of a Small-Scale Joined-Wing Aerial Vehicle under Subsonic Conditions

Flow behavior and aerodynamic performance of a small-scale joined-wing unmanned aerial vehicle (UAV) was studied experimentally and numerically under various pitch and yaw angle combinations in subsonic flow conditions. Selected numerical results are compared against experimental results obtained using surface oil flow visualizations and force measurements, with additional simulations expanding the range of combined pitch and yaw configurations. Under zero-yaw conditions, increasing the pitch angle leads to the formation of symmetric ogive vortex roll-ups close to the fuselage and their significant interactions with the fore-wing. Additionally, contributions to lift and drag coefficients under zero-yaw conditions by the key UAV components have been documented in detail. In contrast, when the UAV is subjected to combined pitch and yaw, no clear evidence of such ogive vortex roll-ups can be observed. Instead, asymmetric flow separations occur over the fuselage’s port side and resemble bluff-body flow behavior. Additionally, these flow separations become more complex, and they interact more with the fuselage and fore- and aft-wings when the yaw angle increases. Lift and drag variations due to different pitch and yaw angle combinations are also documented. Finally, rolling and yawing moment results suggest that the present UAV possesses adequate flight stability unless the pitch and yaw angles are high.

Assessment of Pseudoshock Models Against Experiment in a Low-Aspect-Ratio Isolator

A highly confined shock train is investigated in a direct-connect isolator facility with a Mach 2 inflow and a constant-area low-aspect-ratio rectangular test section. High-speed schlieren imaging, wall static pressure measurements, surface oil-flow visualization, and particle image velocimetry from this isolator are synthesized into a three-dimensional schematic of the shock train structure. Against this, the prevailing pseudoshock models in the literature are assessed to evaluate the validity of their underlying assumptions. None of the prevailing pseudoshock models are found to simultaneously model the pressure and Mach number profiles, indicating a gap in the model formation and underlying assumptions when applied to the experimental isolator of interest. The presence of distortion in the isolator flowfield, such as a wall-bounded vortex, is found to skew the structure of the shock train, altering the strength and distribution of the compressive pressure gradient. It is further observed that the separated flow morphology surrounding the shock train is not monolithic, as is typically assumed, adjusting the balance of compressive forces within the shock cells. These findings lead to the conclusion that existing flux-conserved modeling approaches require modification to be effective in distorted and highly confined cases, including closure models that capture the three-dimensional distorted structure of the approach flow and its evolution along the shock train.

Experimental and computational investigation of the vortical structures generated from a blended-wing-body UAV model

The current study presents a combination of experimental and computational investigations of a scaled-down Blended Wing Body (BWB) Unmanned Aerial Vehicle (UAV) model. The BWB model features highly swept wings. The experiments are conducted in a subsonic windtunnel in the Laboratory of Fluid Mechanics and Turbomachinery, in the Aristotle University of Thessaloniki. More specifically, surface oil flow visualization and Laser-Doppler Anemometry (LDA) are employed to investigate the vortical structures emanating from the leading edge and the wingtips of the model and provide the corresponding velocity and vorticity profiles. Using the windtunnel measurement data as a reference point, both steady-state and transient computational fluid dynamics (CFD) analyses are also conducted. The flow around the BWB scaled-down model is modeled by adopting four turbulence models, three of them based on the concept of the eddy-viscosity and one, on the Reynolds-stress transport equation modeling. The modeling results are compared to the experimental measurements, emphasizing on the streamwise and spanwise velocity profiles, axial vorticity profiles and separation patterns. The results of this study provide insight on the vortical phenomena dominating the flowfield over and around BWB configurations and conclusions concerning the tradeoffs between accuracy and computational efficiency are drawn.