Experimental and Numerical Study of the Thermal Performance of a Film Cooled Turbine Platform

Detailed surface measurements of the thermal performance of a film cooling system have been performed on the endwall of a nozzle guide vane (NGV) mounted in a linear cascade facility at EPFL. An external cooling scheme including several rows of fan-shaped and cylindrical cooling holes has been designed. By testing different cooling flow rates at a NGV exit Reynolds number of 1.7E+06 and Mach number of 0.88, detailed aerodynamic and heat transfer values were obtained destined to assess the design tools for film cooled platforms. The surface static pressure distribution and the film cooling effectiveness on the endwall surface have been experimentally determined. The measurements were obtained applying the pressure sensitive paint technique measuring the coolant gas concentration. An engine representative density ratio between the coolant and the external hot gas flow was achieved by the injection of CO2. The working conditions of the test case similar to realistic engine conditions allow for the validation of in-house CFD codes and the investigation of the reliability of modern commercial tools in such a complex cooling system. The numerical campaign has been performed on the same numerical grid, using the commercial codes FLUENT and CFX, used by EPFL and MTU respectively. A detailed analysis of the grid effects on the obtained results has been previously realised as well as the study of the influence of the modelling approximations. Three cooling mass flows have been simulated and the performance parameters of the film cooling system have been compared to the experimentally obtained data. Special emphasis has been put on the jet penetration effects and on the interaction of secondary flows with the coolant flow. The experimental and numerical efforts were part of the EU funded research project TATEF2 (Turbine Aero-Thermal External Flows 2).

Experimental study on a heavy film cooled nozzle guide vane with contoured platforms

Gregory Vogel

Nowadays, efficiency improvement of the modern gas turbines is usually achieved by increasing the pressure ratio which leads to an increase of the gas temperature in the combustion chamber. As a consequence, the temperature conditions imposed on the first stages of the turbine are well beyond the maximum allowable material of the engine. Therefore, it is necessary to effectively cool on these parts. Some of the highest thermal loading is applied to the nozzle guide vane located just behind the combustion chamber. Its function is to efficiently direct the hot and highpressure gas on the rotor blades of the turbine in order to extract maximum work. From the aerodynamic side, the loss level generated during the hot gas deviation defines the nozzle guide vane efficiency. From the thermodynamic side, an efficient nozzle guide vane is characterized by its cooling protection against the hot gas coming out of the combustion chamber. For this reason, the vane is internally cooled by coolant fluid circulating in channels; but it is also externally cooled by injection of coolant fluid on the surface. This process is usually called "film-cooling". If it is well designed, the film efficiently protects the vane airfoil and platforms of the engine; but if badly designed (unfavorable holes injections positions or inadequate blowing ratios), then the overall performance of the engine is reduced. Recently, studies have shown a decrease in the aerodynamic losses for nozzle guide vanes with contoured platforms. Moreover, it has also been noticed that temperature profiles at the exit of the combustion chamber are more and more flattened and consequently require a better cooling of the platforms. The problem is to efficiently film-cool a nozzle guide vane airfoil and its contoured platforms. A numerical approach of this problem is yet not possible as the numerical codes are not accurate enough to take into account all the physics governing such situations. Therefore, an experimental approach is required. So far, many film cooling studies have been performed but on relatively simple geometries (flat plates, cylinders etc.). The innovative part of this present work was to study film cooling on a complex nozzle guide vane geometry equipped with contoured platforms. For this, as well the aerodynamic side and the thermodynamic side related to film cooling were considered. The experiments were divided into two parts. The first part was dedicated to the study of the vane airfoil. The second part was dedicated to the film-cooling study on the contoured platforms. In both cases, heat transfer coefficients and film-cooling effectiveness were determined by transient measurement techniques using liquid crystals. For the vane airfoil, a preconditioning system followed by a rapid insertion was used. This installation was already used during previous projects, but in the frame of this research work, the liquid crystal technique was improved. For this, a new signal analysis and image processing system was developed. For the measurements on the platforms, a new measurement technique based on the use of a thin electrical heater-foil was developed. The innovation of this technique comes from the fact that it can be applied on complex geometries for which inhomogeneous heat fluxes are produced (curved surfaces, introduction of cooling holes). This novel measurement technique was first tested and validated on a simple geometry consisting of a film-cooled flat plate. The same technique was then applied to the film-cooled contoured platform and allowed obtaining interesting results of heat transfer coefficient and film-cooling effectiveness distributions. The latter result was also compared to values obtained from another innovative measurement technique based on Pressure Sensitive Paint with Nitrogen injection. The results obtained in the frame of this experimental work dedicated to film cooling applied on a complex geometry were obtained thanks to the development of novel measurement techniques. These measurement techniques provide the opportunity to perform systematic tests in order to gain further insight into the physics of the film cooling process on complex geometries such as contoured platforms.

EPFL2002

Experimental investigations of showerhead film cooling on the leading edge of a turbine blade

Guillaume Wagner

The turbine components of a gas turbine need extensive cooling to withstand the high heat loads generated by the flow of the hot combustion gases. The leading edge of a turbine blade is one of the areas that faces the hottest gas flow conditions and is thus one of the most critical areas to be cooled. The stagnation region of turbine airfoils is generally cooled with showerhead cooling scheme, which consist in injecting coolant onto the stagnation region through closely spaced rows of film cooling holes. This provides convective cooling inside the holes and produces a layer of coolant film on the external surface. The objective of this study is to investigate experimentally the influence of mainstream incidence on showerhead film cooling. The film cooling effectiveness and heat transfer coefficient were thus measured on a series of symmetric blunt bodies with different showerhead configurations simulating the leading edge of turbine airfoils. Tests have been performed at 0° on-design and 5° off-design mainstream incidences at an approach Mach number of 0.26. Besides the mainstream incidence angle, the investigated parameters were the blowing ratio, the wedge angle, the row location, the spanwise pitch between the holes and the hole shape. In addition to the film cooling measurements, baseline heat transfer tests were performed on test models without film cooling holes in order to evaluate the film cooling performances. To characterize the mainflow conditions and the velocity profile on the test models, aerodynamic measurements have also been carried out. The heat transfer measurements have been performed using the transient liquid crystal technique. The transient experiments were generated by a pre-conditioning system and a rapid insertion mechanism. The data analysis was adapted to measurements on highly curved surfaces using the approximate analytical solution of the heat conduction into a solid cylinder with convective boundary conditions on the surface. Time-varying adiabatic wall temperatures were accounted for using Duhamel's superposition theorem. The first tested showerhead geometry was a triangular blunt body representative of a turbine rotor blade showerhead with a circular leading edge, a 21.1° wedge angle and two film cooling rows. Because of jet lift-off, the film cooling performances of this two-row configuration were relatively low and were almost unaffected by the increase of blowing ratio from 1.5 to 3.9. The heat transfer coefficient was found to be greatly increased at the row location in comparison to the case without film cooling. At 5° off-design incidence angle, the stagnation line moves passed the pressure side row. The two rows were then located on the same side of the stagnation line and were both blowing towards the suction side. On the suction side the film cooling protection was thus very good due to film accumulation but the pressure side was left almost unprotected resulting in highly asymmetric film cooling performances. Two-dimensional FEM calculations of the metal temperature in this second geometry have been performed at engine representative conditions for design and off design mainstream incidences. It was found that the contribution of the film cooling to cool the leading edge region was rather small in comparison to the contribution of the internal convective cooling. Showerhead film cooling was however required to effectively cool the stagnation region and to obtain a uniform surface temperature. The second tested showerhead geometry was similar to the first geometry but with an increased wedge angle of 55° and cooling rows located further downstream. The film cooling performances were found to decrease when the blowing ratio was increased from 1.2 to 3.0 because of jet lift-off. At middle blowing ratio (2.0), the film cooling effectiveness had similar values as for the first geometry. However, the heat transfer increase due to the presence of the film cooling jets was much higher for the second geometry because of higher local mainstream velocities at the injection locations. The resulting film cooling performances were thus lower for the second geometry than for the first geometry. Nevertheless, the larger wedge angle and the cooling rows located further downstream assured that the stagnation line stayed in-between the two cooling rows even at 5° off-design incidence. The second geometry had thus the advantage to be much less sensitive to incidence angle and, unlike the first geometry, provided very similar film cooling performances at 5° and at 0° incidence. The use of shaped holes was found to greatly improve the film cooling performances of the second geometry due to a better lateral coverage and a reduced tendency of jet lift-off.

EPFL2007

Application of the Transient Heater Foil Technique for Heat Transfer and Film Cooling Effectiveness Measurements on a Turbine Vane Endwall

Dominique Charbonnier, Magnus Jonsson, Peter Ott

The paper presents an application of the transient heater foil measurement technique using thermochromic liquid crystals (TLC) to endwall heat transfer and film cooling investigations in a transonic turbine vane cascade. The film cooling configuration consists of an upstream slot, representing the leakage flow area between the interface of the combustor and the turbine, and several rows of cylindrical and fan-shaped holes within the passage. With the transient method chosen, the heat transfer and adiabatic film cooling effectiveness distributions can be obtained simultaneously together with the local heat release within the heater foil. Therefore, the heat release in the foil is not required to be uniform, and the foil can contain discrete holes in the film cooling configuration. Some new developments are presented, which are directed towards improved application of the transient heater foil method for such a complex configuration. This includes tailoring the foil heat-release distribution towards the expected heat transfer patterns, supported by numerical Finite-Element computations and the use of a double-TLC mixture for improved time-wise TLC indications. Additionally, CFDsimulations were used to evaluate the recovery temperature distribution through the vane cascade without film cooling. The experiments were performed in the linear cascade facility at the EPFL-Lausanne. A compressor provides a continuous air flow at near-ambient temperature regulated with heat exchangers. Carbon dioxide is used as coolant in order to achieve engine-representative density ratio between coolant and main flow. Multiple experiments with the same main and coolant flow settings but varying heat flux levels and coolant injection temperatures have been performed and simultaneously analysed using nonlinear regression analysis. The time required between successive experiments to return to homogenous initial conditions, as required by the transient method, has been analysed using an analytical solution for heat-on-heat-off conditions. This permits the model assumption of onedimensional conduction within a semi-infinite wall with a heat releasing layer on the top. Example results for cases without cooling, with film cooling from rows of discrete holes and the addition of slot film cooling are used to illustrate the benefit of the new approaches for the investigated vane cascade.

American Soc Mechanical Engineers, Three Park Avenue, New York, Ny 10016-5990 Usa2008