Experimental and Computational Comparisons of Fan-Shaped Film Cooling on a Turbine Vane Surface (original) (raw)
Related papers
International Journal of Geomechanics, 2015
This research was done to analyze the effects of two different blowing ratios of BR=1.25 and BR=3.18 on the film cooling effectiveness at the combustor outlet, whereas the cylindrical and row trench holes with alignment angle of +60 degrees were considered. In the current research, a three-dimensional representation of a Pratt and Whitney gas turbine engine was simulated and analyzed with a commercial finite volume package FLUENT 6.2.26. This study has been carried out with Reynolds-averaged Navier-Stokes turbulence model (RANS) on internal cooling passages. The combustor combines the interaction of two rows of dilution jets, which are staggered in the stream-wise direction and aligned in the span-wise direction, with that of film cooling along the combustor liner walls. The entire findings of the study showed that trenched holes performed much more efficiently at both blowing ratios, especially at BR=3.18.
Mechanical Engineering Journal
The reliability of turbine blades and vanes of modern high temperature gas turbines is assured by turbine blade cooling technologies. Among the various cooling methods, film cooling has been a key technology to ensure the long-term operation of turbine blades and vanes that are exposed to hot gas-path flows. Therefore, many papers have been published aiming at the improvement of film cooling effectiveness by optimization of film cooling hole geometries. Although the turbulence intensity of the mainstream generated in the gas turbine combustor is very high and may reduce the film cooling effect on the turbine vane and blade, there are few papers investigating quantitatively the effect of the mainstream turbulence on the film cooling. For this reason, the influence of mainstream turbulence intensity on film cooling effectiveness was investigated with an active turbulence generator equipped with electric-motor driven propellers for circular and fan-shaped film cooling holes. Spatial distributions of the turbulent mixing field between turbulent mainstream and film coolant jet were measured with quantitative measurement methods, such as PIV and LIF. As a result, it was revealed that when the mainstream turbulence is high the counter-rotating vortex pair is weakened, the film cooling air spreads in the span-wise direction and the lateral-averaged film cooling effectiveness decreases about 10%.
Proceeding of 5th Thermal and Fluids Engineering Conference (TFEC)
The higher temperature of the combustion chamber of a gas turbine yield higher efficiencies for the turbines but can affect the blade's life. As a preventative method, Film-cooling is a useful technique to enhance the performance of it by injecting coolant jets that the metal surfaces can be protected against the hot main flow. To improve cooling efficiency and increase the life of these components, several cooling strategies have been introduced. In the present study, the effects of the laidback fan-shaped hole geometry, in particular, are examined against the conventional cylindrical hole geometry using Computational Fluid Dynamic simulations. Computations are carried out based on three-dimensional Large Eddy Simulation. The open-source software OpenFOAM was utilized to solve the filtered governing equations for mass, momentum, energy conservation, and heat transfer. The mixing mechanism between hot and coolant fluids, nondimensional adiabatic film cooling effectiveness, and dynamics of vortices are presented and discussed.
50th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition, 2012
An experimental investigation was performed to study the cooling performance and convective heat transfer on a film cooled turbine blade surface. A 2D linear cascade model of the first stage turbine rotor blade of a land-based gas turbine was employed in the study. The film cooling configuration on the blade comprises of 2 rows of fan-shaped holes on the pressure side (PS), and 1 row of fan-shaped hole on the suction side (SS). The tests were performed in the Virginia Tech transonic wind tunnel facility, which simulates engine representative conditions of high turbulence intensity at the inlet and transonic Mach numbers at the exit, with matching Reynolds number to that of the real engine. All the tests were performed at inlet turbulence intensity of 12% with integral length scale of 0.26 normalized by cascade blade pitch. Exit Mach number of 0.67, 0.84, and 1.01 were chosen for the tests. Two combinations of blowing ratios at different rows of cooling holes were tested. (Nominal blowing ratio settings are: suction side injection BR=1.2 and 1.6; pressure side row 1 BR=2.8 and 3.8; pressure side row 2 BR=2.2, and 2.8.) The shock wave/boundary layer interaction effect on Nusselt number and adiabatic effectiveness was observed on the suction side (SS) with a cascade exit Mach number of 1.01. The trends for Nusselt number and adiabatic effectiveness across the shock impinged region suggest the existence of a separation bubble at the location of shock impingement. The detailed discussion of the shock effect on film cooling is presented.
Film cooling of a contoured endwall nozzle vane through fan-shaped holes
International Journal of Heat and Fluid Flow, 2010
The present paper shows the results of an experimental investigation into a nozzle vane cascade with endwall contouring and film cooling. In a film cooled endwall the front area, especially close to the leading edge, is one of the most critical regions to be cooled; in case of a contoured endwall the thermal protection becomes even more difficult because of the endwall curvature effect. Two film cooling hole geometries with four rows of holes were tested: cylindrical holes and conical expanded holes. Tests were performed at low speed (M 2is = 0.2) and low inlet turbulence intensity level, with coolant to main stream mass flow ratio varied within the 0.5-2.5% range. Aerodynamic and thermal performances were evaluated in every injection condition. Compared to cylindrical holes, results showed that shaped film cooling holes generate slightly higher secondary losses and provide a better thermal coverage, at the expense of a higher coolant flow injection.
Progress Of Film Cooling In Industrial Gas Turbine Vanes And Blades
2017
Cooling air can be chilled for industrial gas turbines. By taking advantage of this feature, a deep digging fan-shaped film cooling hole equipped the function of a trench has been devised. The characteristics of this film cooling hole were investigated using quantitative measurement methods. It was revealed that this film cooling hole had high film cooling effectiveness at high mass flux ratio M of 11.5. It was also revealed that film cooling effectiveness could be greatly improved when the film cooling air was swirled. Cooled turbine blades may be manufactured with the additive manufacturing method in the future, and the relationship between the positions of the film cooling hole and internal cooling structure can determined. In this case, it becomes possible to design turbine blades utilizing the swirling flow generated by the internal cooling structure to attain high film cooling effectiveness.
Sweeping Jet Film Cooling at High Blowing Ratio on a Turbine Vane
Volume 5B: Heat Transfer, 2019
Experimental and numerical investigations were conducted to study the effects of high blowing ratios and high freestream turbulence on sweeping jet film cooling. Experiments were conducted on a nozzle guide vane suction surface in a low-speed linear cascade. Experiments were performed at blowing ratios of 0.5–3.5 and freestream turbulence of 0.6% and 14.3%. Infrared thermography was used to estimate the adiabatic cooling effectiveness. Thermal field and boundary layer measurement were conducted at a cross-plane (x/D = 12) downstream of the hole exit. Results were compared with a baseline 777-shaped hole and showed that sweeping jet hole has a better cooling performance at high blowing ratios. The Thermal field data revealed that the coolant separates from the surface at high blowing ratios for the 777-shaped hole while the coolant remains attached for the sweeping jet hole. Boundary layer measurement further confirmed that due to the sweeping action of the jet, the jet momentum of t...
Effects of Realistic Inflow Conditions on the Aero-Thermal Performance of a Film-Cooled Vane
A high-pressure vane equipped with a realistic film-cooling configuration has been studied. The vane is characterized by the presence of multiple rows of fan-shaped holes along pressure and suction side while the leading edge is protected by a showerhead system of cylindrical holes. Steady three-dimensional Reynolds-Averaged Navier-Stokes simulations have been performed. A preliminary grid sensitivity analysis has been performed with uniform inlet flow to quantify the effect of the spatial resolution. Turbulence model has been assessed in comparison with available experimental data. The effects of a realistic inflow condition on the thermal behaviour of the cooled vane are then investigated by means of comparison between two conjugate heat transfer simulations. The first one is characterized by a uniform inlet flow while the second one presents a temperature distortion and a superimposed aggressive swirl derived from the EU-funded TATEF2 project. The effect of the swirling flow in d...
Numerical Simulation of Film Cooling a Turbine Blade Through a Row of Holes
Journal of Thermal Engineering, 2017
We undertake a numerical three-dimensional study of the interaction of a row of discrete jets in a wall with a transversal compressible flow for different injection rate (M=0.3, 0.5, 0.7 and 1.4). This simulation is applied to the stator blade of the CFM56 engine and is performed using the computational fluid dynamics (CFD) simulation tool, with CFX.13 software. Reynolds averaged Navier-Stokes equations were solved using a finite volume method. Turbulence closure was achieved using the Shear-Stress Transport model (SST). The velocity and temperature distributions and the film cooling effectiveness are presented and discussed. We found that the best cooling effectiveness occurs at M= 0.7. More, for higher injection rate (M=1.4), the results show the existence of two counter-rotating vortices. These vortices transport the hot gas in the jet and thus degrade the protective wall.
Film Cooling from Two Staggered Rows of Compound Angle Holes at High Blowing Ratios
International Journal of Rotating Machinery, 1996
Experimental results are presented which describe the development and structure of flow downstream of two staggered rows of film-cooling holes with compound angle orientations at high blowing ratios. These film cooling configurations are important because they are frequently employed on the first stage of rotating blades of operating gas turbine engines. With this configuration, holes are spaced 3d apart in the spanwise direction, with inclination angles of 24 degrees, and angles of orientation of 50.5 degrees. Blowing ratios range from 0.5 to 4.0 and the ratio of injectant to freestream density is near 1.0. Results show that spanwise averaged adiabatic effectiveness, spanwise-averaged iso-energetic Stanton number ratios, surveys of streamwise mean velocity, and surveys of injectant distributions change by important amounts as the blowing ratio increases. This is due to injectant lift-off from the test surface just downstream of the holes which becomes more pronounced as blowing rat...