Experimental investigation of an airfoil equipped with a mid-chord slot (MCS) on suction side

Measurements of heat transfer and pressure in a trailing edge cavity of a turbine blade

This experimental study examines the downstream wake development influenced by attached turbulent boundary layers under adverse and favorable pressure gradients (APG, FPG) on the suction and pressure sides of a National Advisory Committee for Aeronautics 4412 airfoil. Experiments were performed at a chord-based Reynolds number of Rec=400 000 and angles of attack of 5° and 8°. Surface pressure scans and hot-wire anemometry were used to characterize boundary layer and wake profiles, extending three chord lengths downstream of the trailing edge. The key contributions are: (i) demonstrating a strong memory effect whereby upstream boundary-layer characteristics persist into the very-near- and near-wake; (ii) establishing a unified inner/outer velocity scaling framework that enables direct comparison of boundary-layer and wake profiles; and (iii) identifying distinct suction- vs pressure-side routes to self-similarity driven by the sign of the pressure gradient. Inner scaling of wake velocity profiles using friction velocity and viscous length scale proves valid only within the very-near-wake. In contrast, outer-scaled velocity-defect profiles using the Zagarola–Smits velocity reveal self-similarity on the suction side from the trailing edge across the entire wake region, whereas pressure-side profiles retain significant upstream influence initially, with self-similarity emerging from the near-wake onward. Turbulence analysis via streamwise velocity variance and premultiplied power spectral densities further shows distinct development patterns on the suction and pressure sides of the wake. The APG on the suction side promotes early development of wake-type structures and self-similarity, while the FPG on the pressure side preserves boundary-layer features deeper into the wake. These findings demonstrate how upstream pressure gradients modulate the transition from boundary-layer-dominated to wake-dominated behavior, providing new physical insight into wake development and supporting improved aerodynamic modeling and wake prediction for aeronautical and wind energy applications.

The process of film cooling is known to severely disturb the boundary layer around a turbine airfoil. Since most film-cooled airfoils have more than one injection station, the flow field approaching a row of film cooling holes could be altered by the presence of an upstream cooling station. To investigate this possibility, an experimental investigation was conducted on the suction side of a scaled-up turbine vane. Adiabatic effectiveness measurements were made downstream of a single row of cooling holes both with and without the upstream showerhead holes operating. A range of suction side blowing ratios, 0.3 ≤ M ≤ 1.3, were investigated with large-scale mainstream turbulence intensities of Tu∞ = 0.5% and Tu∞ = 21%. The effects of the showerhead coolant were evaluated at an engine-typical showerhead blowing ratio of Msh = 1.6, with three of the six rows of cooling holes in the showerhead directed towards the suction side of the airfoil. Experiments were conducted with a coolant-to-mainstream density ratio of DR = 1.6. An infrared camera was used to obtain spatially-resolved surface temperature measurements, which were corrected for conduction effects and converted to adiabatic effectiveness. The results showed that showerhead coolant had a strong impact on suction side adiabatic effectiveness levels under low mainstream turbulence. Although effectiveness levels increased with the showerhead operating, the suction side coolant jets increased dispersion of the showerhead coolant. Under high mainstream turbulence conditions, there was very little interaction between the showerhead coolant and the suction side coolant jets. Adiabatic effectiveness levels were considerably lower than those for the low turbulence case, which was partially due to increased dispersion of the showerhead coolant upstream of the suction side holes. The superposition model over-predicted adiabatic effectiveness levels under low mainstream turbulence conditions, but was very effective in predicting the combined performance of the showerhead and the suction side cooling holes under high mainstream turbulence conditions.

An experimental and numerical investigation of boundary layer suction (BLS) on a transonic compressor blade has been conducted. The objective of the present work is to identify the possible benefit of boundary layer suction via a slot in the region of the shock/boundary layer interaction (SBLI). The study has two major parts: First, an investigation of the flow on an isolated airfoil without and with BLS on the suction side (SS). This investigation is intended to optimise the suction slot geometry in terms of location and width. Second, a study on a transonic cascade with supersonic inlet flow has been conducted with the goal to show the influence of the BLS on the pressure loss coefficient and the surface Mach number distribution. An initial literature study revealed that only few work has been carried out in the field of BLS in transonic compressors and no data close to the suction slot was available. The main steps of the present study are: The design of an isolated airfoil for testing in a Laval nozzle, representative for the flow at a compressor root section, with enough space inside to evacuate the aspirated air. A numerical study of the flow on the isolated airfoil without and with BLS on the SS . An experimental investigation of the isolated airfoil without and with BLS on the SS using static surface pressure taps, PSP and the Schlieren method for data acquisition. Design and numerical investigation of a suitable transonic compressor cascade with enough space inside to evacuate the suction mass flow. An experimental investigation of the designed transonic cascade with an inlet Mach number of 1.23 with two sets of blades in the non-rotating annular cascade test-rig; one without and one with BLS at 40% chord. The main results are: The numerical study of the flow on the isolated airfoil without and with BLS on the SS shows, that the onset of shock-induced separation could be shifted to an increased inlet Mach number. Steady 2D-NS simulations show that the reference configuration exhibits an attached flow behind the shock up to an inlet Mach number of up to 0.70 for an inlet flow angle of +4°. The configuration with BLS shows attached flow behind the shock for an an inlet Mach number of 0.725. Unsteady 2D-NS simulations show for an inlet flow angle of +4° at an inlet Mach number of 0.725 for the reference configuration a shock, which moves periodically on the SS due to a separated boundary layer created by the shock/boundary layer interaction. The same configuration with BLS shows a stable shock position. The experimental investigation of the isolated airfoil without and with BLS on the SS shows that BLS at 15% chord upstream of the initial shock location for an inlet Mach number of 0.725 at a flow angle of +4° leads to a shock location 10% chord downstream of the initial position. Boundary layer suction suppresses flow separation and stabilises the shock. Due to the suction a stagnation point is created at the downstream edge of the suction slot. Downstream of this stagnation point, the flow is re-accelerated to the isentropic Mach number level of the reference configuration without BLS. The design and 3D-numerical investigation of a suitable transonic compressor cascade shows, that BLS with a suction mass flow of 2% can increase the maximum pressure ratio and diffusion factor of approximately 10%. This is the consequence of the stabilisation of the shock in the blade passage due to the suction. The experimental investigation of the designed transonic cascade with an inlet Mach number of 1.23 with 2% suction mass flow shows, that BLS acts locally and that suction leads to an increased isentropic surface Mach number in front of the suction slot and due to the suction related stagnation point, to a locally decreased velocity immediately downstream of the suction slot. Then the flow re-accelerates to the same level as the non-suction configuration attains. The pressure loss coefficient is decreased by up to 3.6% for the most positive incidences. The results of this investigation confirm that boundary layer suction on the suction side of an airfoil is a suitable way to increase the working range of a compressor as the shock position can be stabilised (shock-trapping) and the increase of the pressure loss coefficient can be shifted to higher incidences.

An experimental investigation of three-dimensional flow field in a film-cooled turbine model is carried out by using particle image velocimeter (PIV) in a low-speed wind tunnel. The effects of different blowing ratios (M=1.5, 2) on the flow field are studied. The experimental results reveal the classical phenomena of the formation of kidney vortex pair and secondary flow in wake region behind the jet hole. And the changes of the kidney vortex pair and the wake at different locations away from the hole on the suction and pressure sides are also studied. Compared with the flow field in stationary cascade, there are centrifugal force and Coriolis force existing in the flow field of rotating turbine, and these forces bring the radial velocity in the jet flow. The effect of rotation on the flow field of the pressure side is more distinct than that on the suction side from the measured flow fields in Y-Z plane and radial velocity contours. The increase of blowing ratio makes the kidney vortex pair and the secondary flow in the wake region stronger and makes the range of the wake region enlarged.

Effect of the laminar separation bubble induced transition on the hydrodynamic performance of a hydrofoil

Experimental investigations of the film cooling heat transfer coefficient of a Micro-Tangential-Jet scheme on a gas turbine vane

This paper experimentally investigates the film cooling performance of an enlarged turbine guide vane with full-coverage cylindrical hole film cooling in short duration transonic wind tunnel which can model realistic engine aerodynamic conditions and adjust inlet Reynolds number and isentropic exit Mach number independently. The effects of mass flow rate ratio (MFR=4.83%∼8.83%), inlet Reynolds number (Rein= 1.7×105∼5.7×105), and isentropic exit Mach number (Mais=0.81∼1.01) are investigated. There are five rows of cylindrical film cooling holes on the pressure side and four such rows on the suction side respectively. Another four rows of cylindrical holes are provided on the leading edge to obtain a showerhead film cooling. The surface heat transfer coefficient and adiabatic film cooling effectiveness are derived from the surface temperatures measured by the thermocouples mounted in the middle span of the vane surface based on transient heat transfer measurement method. Mass flow rate ratio is shown to have a significant effect on film cooling effectiveness. The increase of mass flow rate ratio increases film cooling effectiveness on pressure side, while increasing this factor has opposite effect on film cooling effectiveness on the suction side. At the same mass flow rate ratio, increasing the Reynolds number can enhance the film cooling performance, the expectation is that at low mass flow rate ratio condition increasing the Reynolds number decreases film cooling effectiveness on the pressure side. The heat transfer coefficient increases with the mass flow rate ratio increasing on both pressure and suction side. At middle and high inlet Reynolds number condition, in the region of 0.4<s<0.6 on suction side, the coolant weakens heat transfer adversely.

ABSTRACTWater‐suspended sediments wear down hydro turbine components by erosion, reducing their lifetime. Nonetheless, even heavily sediment‐loaded rivers are a valuable renewable and clean energy resource. Forecasting sediment erosion and optimizing the hydraulic turbine design for extended durability by numerical flow simulations became imperative to operate the hydropower plant safely and economically in sediment‐laden environments. This strategy is challenged by the unknown model parameters, that is, the lack of reliable validation data. Therefore, in this study, a 2 kW Francis‐type model turbine is tested in a non‐recirculating sediment‐laden test facility. The Francis turbine runner and the guide vanes are coated with four different colors to visualize locations of surface degradation qualitatively due to erosion during operation in representative sediment concentrations. The turbine is tested at two different operating conditions OP1 and OP2 for visualizing the locations of erosion. For operating condition OP1, with low rotational speed, erosion is primarily observed at the leading and trailing edge of the suction and pressure side, respectively. The Francis turbine runner is particularly eroded in the transition between the hub and the blades on the suction side towards the trailing edge. Meanwhile, for operating condition OP2, with higher rotational speed, the trailing edge of the pressure side of the blade and the region of the shroud close to the trailing edge of the blades are found to be vulnerable to erosion. Test for the material erosion of the runner is conducted at OP3 conditions with high sediment concentration over 45 h, reporting the weight loss in intervals of 15 h. It is observed that the cumulative erosion rate of the Francis runner made of brass material is 0.016 mg g−1 h−1 after 45 h of operation. Similarly, the percentage loss of the runner with respect to the hours of operation is calculated to be 0.0065%, 0.04%, and 0.07% for 15, 30, and 45 h of operation, respectively. This data set can also be useful for qualitative and quantitative validation of computational fluid dynamic simulations for erosion prediction.

The demand for significantly higher performance gas turbine engines has led to the exploration and identification of "Out of the Box" innovative engine design concepts. These demands include increased thrust-to-weight ratio goals that can primarily be met by substantial engine performance increases such as specific thrust, engine weight and size reductions, and repackaging of engine components to create compact engines. Concepts of an Ultra-Compact-Combustor (UCC) for use as a main combustor, or as an Inter- Turbine Burner (ITB) to boost engine work output, reduce pollutant emissions and engine weight are being explored. The available experimental results and observations indicate that UCC/ITB can operate at 95­99% combustion efficiency over a wide range of operating conditions and with flame lengths up to 50% shorter than those of conventional combustors. In the present study the radial curved vane ITB design concept has been modeled using three-dimensional computational fluid dynamics (CFD). The objectives are to predict ITB flow field and combustion characteristics, guide ITB experimental investigations, identify the key design parameters driving performance, and use the results to optimize ITB design configurations. The CFD predictions demonstrated that intense burning in a high-g loaded cavity occurred which resulted in high combustion efficiency. Models with the radial vane cavity located in both the suction and pressure side have been developed. The circumferential cavity air is injected through the air injection tubes into the circumferential cavity. The orientation of this injection is used to create both a clock-wise (CW) and a counter-clock-wise (CCW) direction of circumferential flow in the outer cavity, when looking upstream from the aft end of the ITB configuration. The resulting five candidate configurations have been simulated and analyzed in detail. This study indicates improved exit profile characteristics for the curved radial vane (CRV) with the cavity in the suction side and the air injected in the CCW direction, compared to the pressure side cavity with air injected either in CCW or CW direction and CRV with no cavity.

Experimental investigations of the film cooling effectiveness of a micro-tangential-jet scheme on a gas turbine vane

At high Reynolds number, the flow of an incompressible viscous fluid over a lifting surface is a rich blend of fluid dynamic phenomena. Here, boundary layers formed at the leading edge develop over both the suction and pressure sides of the lifting surface, transition to turbulence, separate near the foil's trailing edge, combine in the near wake, and eventually form a turbulent far-field wake. The individual elements of this process have been the subject of much prior work. However, controlled experimental investigations of these flow phenomena and their interaction on a lifting surface at Reynolds numbers typical of heavy-lift aircraft wings or full-size ship propellers (chord-based Reynolds numbers, $Re_C {\sim} 10^7{-}10^8$) are largely unavilable. This paper presents results from an experimental effort to identify and measure the dominant features of the flow over a two-dimensional hydrofoil at nominal $Re_C$ values from near one million to more than 50 million. The experiments were conducted in the US Navy's William B. Morgan Large Cavitation Channel with a solid-bronze hydrofoil (2.1 m chord, 3.0 m span, 17 cm maximum thickness) at flow speeds from 0.25 to 18.3 m s$^{-1}$. The foil section, a modified NACA 16 with a pressure side that is nearly flat and a suction side that terminates in a blunt trailing-edge bevel, approximates the cross-section of a generic naval propeller blade. Time-averaged flow-field measurements drawn from laser-Doppler velocimetry, particle-imaging velocimetry, and static pressure taps were made for two trailing-edge bevel angles (44$ ^\circ$ and 56$ ^\circ$). These velocity and pressure measurements were concentrated in the trailing-edge and near-wake regions, but also include flow conditions upstream and far downstream of the foil, as well as static pressure distributions on the foil surface and test section walls. Observed Reynolds-number variations in the time-averaged flow over the foil are traced to changes in suction-side boundary-layer transition and separation. Observed Reynolds-number variations in the time-averaged near wake suggest significant changes occur in the dynamic flow in the range of $Re_C$ investigated.

Film cooling is widely used in modern gas turbines for the protection of the hot components against hot gases from the combustion process. Film cooling directly influences the thermal efficiency of the gas turbine, as the cooling gas is extracted from the compressor and mixed with the mainstream in the hot component. Huge efforts by industry as well as research organizations have been undertaken to improve the film cooling effectiveness. It can been concluded that there are two key points for the improvement of film cooling effectiveness, constraining the blow-off of cooling ejection and extending the lateral coverage of cooling gas. The paper presents a new cooling technology, which reaches high film-cooling effectiveness as a result of a well-designed cooling hole, named SYCEE film cooling technology (SFCT). Plate film cooling experiments of SYCEE tested by pressure sensitive paint (PSP) are carried out in this work, and traditional shape-hole are included as well for baselines. It is resulted that SFCT has a better film cooling performance than shape-hole in the same conditions, and the gap of the averaged film cooling effectiveness between them continuously enlarges as the blowing ratio increases. Furthermore, an application of SFCT on the first stage vane of an F-class gas turbine is studied as well. A two-dimension cascade has been employed to measure the cooling performance of SFCT using pressure sensitive paint (PSP) as well, and the tested vanes separately with round-hole and shape-hole are considered again for baselines. The different kinds of film holes separately locate on the pressure and suction side, while the showerhead in different cases are kept the same, arranged with round-holes. The cooling air is ejected at inclination angle 45° with compound-angle 90° in the showerhead and inclination angle 35°∼45° without compound-angle on the pressure side and suction side. The detailed local cooling effectiveness distributions as well as the span-averaged effectiveness over the vane surface are presented. As expected, the film cooling performance of round-hole is the worst due to the lift-off of the cooling ejection. SFCT has better film cooling performance than shape-hole on the pressure side, but the advantage decreases along the mainstream direction. However, the span-averaged film cooling effectiveness of SYCEE is similar with that of the shape-hole on the suction side. This may be due to enhanced impact of mainstream flow derived from the pressure gradient in the turbine passage, and consequently weakening the effect of film hole on the suction side.

This paper presents experimental investigations of the film cooling effectiveness performance of a Micro-Tangential-Jet (MTJ) Film cooling scheme on a gas turbine vane using transient Thermochromic Liquid Crystal (TLC) technique. The MTJ scheme is a micro-shaped scheme designed so that the secondary jet is supplied tangentially to the vane surface. The scheme combines the benefits of micro jets and tangential injection. The film cooling performance of one row of holes on both pressure and suction sides were investigated at a blowing ratio ranging from 0.5 to 1.5 on the pressure side and 0.25 to 0.625 on the suction side. The average density ratio during the investigations was 0.93, and the Reynolds Number was 1.4E+5, based on the free stream velocity and the main duct hydraulic diameter. The pitch to diameter ratio of the cooling holes is 5 on the pressure side and 6.5 on the suction side. The turbulence intensity during all investigations was 8.5%. Minor changes in the Mach number distribution around the airfoil surface were observed due to the presence of the MTJ scheme, compared with the case with no MTJ scheme. The investigations showed great film cooling performance for the MTJ scheme, high effectiveness values, and excellent lateral jet spreading. A 2-D coolant film was observed in the results, which is a characteristic of the continuous slot schemes only. The presence of this 2-D film layer helps minimize the rate of mixing between the main and coolant streams and provides uniform thermal loads on the surface. Furthermore, it was noticed that the rate of effectiveness decay on the suction side was less than that on the pressure side, while the lateral jet spreading on the pressure side was better than that of the suction side. The main disadvantage of the MTJ scheme is the increased pressure drop.

This paper aims to numerically and experimentally analyze the effect of the micro-jet on the suction side (SS) of the transonic high-pressure turbine cascade, which could enable the possibility of the mass matching for turbine of the variable cycle engine (VCE) when it transforms working conditions according to flight demands. The experimental study had been conducted in a transonic and supersonic cascade wind tunnel to observe the influence of this kind of flow control technique on aerodynamic performance, operated at different exit Mach numbers. The nozzle guide vane with convergent-divergent passage was investigated in this paper. Two-dimensional steady simulations were performed to investigate the influence of the flow control method to exploit the potential of the application of this method in the design of high-pressure (HP) turbine for VCE. Results of the numerical simulations and experiments indicated that micro-jet on the SS of the turbine blade would generate a separation zone near the jet slot, which would reduce the genuine throat area of the turbine passage and thus change the mass flow rate through the passage of the turbine cascade significantly. Also, velocity at the passage throat would be reduced due to the application of SS micro-jet, which would also contribute to the reduction of the mass flow rate through the turbine passage. However, extended total pressure loss was introduced with SS micro-jet.