Adiabatic Film-cooling Effectiveness Research Articles

Large-eddy simulation of the interaction of wall jets with external stream

Large eddy simulations are performed for a wall jet with an external stream. The external stream is in the form of a heated boundary layer. This is separated from a cold wall jet by a thin plate. The Reynolds number based on the displacement thickness, for the incoming boundary layer is 2776. A series of jet velocity ratios in the range M=Uj/U∞=0.30–2.30, is considered. The wall jet and outer stream velocities are Uj and U∞, respectively. The jets with M⩽1.0 develop von-Karman type shed vortices in the wake region. The higher velocity ratio jets with M>1.0 undergo Kelvin–Helmholtz instability and develop closely spaced counter-clockwise rolling structures. These structures determine the mean flow field behaviour and near wall heat transfer. At any given streamwise location adiabatic film-cooling effectiveness for M<1.0 increases rapidly with increasing M. For M>1.0 it decays slowly with further increase in M. For M<1.0 heat transfer from the hot outer stream to the wall depends on two factors; mean wall normal velocity and wall normal turbulent heat flux. For M>1.0 only a wall normal turbulent heat flux is responsible for heat transfer to the wall. The scaling behaviour shows that the near wall flow scales with wall parameters for all values of M. However, scaling in the outer region is highly dependent on M. The flow develops towards a boundary layer in the farfield for M<1.0 and towards a wall jet for the highest velocity ratio M=2.30.

Influence of Coolant Density on Turbine Blade Film-Cooling With Axial and Compound Shaped Holes

Adiabatic film-cooling effectiveness is examined systematically on a typical high pressure turbine blade by varying three critical flow parameters: coolant blowing ratio, coolant-to-mainstream density ratio, and freestream turbulence intensity. Three coolant density ratios 1.0, 1.5, and 2.0 are chosen for this study. The average blowing ration and the turbulence intensity are 1.5% and 10.5%, respectively. Conduction-free pressure sensitive paint (PSP) technique is used to measure film-cooling effectiveness. Foreign gases are used to study the effect of coolant density. Two test blades feature axial angle and 45 deg compound-angle shaped holes on the suction side and pressure side. Both designs have 3 rows of 30 deg radial-angle cylindrical holes around the leading edge region. The inlet and the exit Mach number are 0.27 and 0.44, respectively. Reynolds number based on the exit velocity and blade axial chord length is 750,000. Overall, the compound angle design performs better film coverage that axial angle. Greater coolant-to-mainstream density ratio results in lower coolant-to-mainstream momentum and prevents coolant to lift-off.

Enhancement of Adiabatic Film Cooling Effectiveness by Using Conical Shape Hole

Film cooling is one of the methods used to protect the surfaces exposed to hightemperature flows, such as those exist in gas turbines. It involves the injection of coolant fluid (at a lower temperature than that of the main flow) to cover the surface to be protected. This injection is through holes that can have various shapes; simple shapes, such as those with straight cylindrical or shaped holes (included many holes geometry, like conical holes). The computational results show that immediately downstream of the hole exit, a horseshoe vortex structure consisting of a pair of counter-rotating vortices is generated. This vortex generation affected the distribution of film coolant over the surface being protected. The fluid dynamics of these vortices are dependent upon the shape of the film cooling hole, and blowing ratio, therefore the film coolant coverage which determines the film cooling effectiveness distribution and also has an effect on the heat transfer coefficient distribution. Differences in horseshoe vortex structures and in resultant effectiveness distributions are shown for cylindrical and conical hole cases for blowing ratios of 0.5 and 1. The computational film cooling effectiveness values obtained are compared with the existing experimental results. The conical hole provides greater centerline film cooling effectiveness immediately at the hole exit, and better lateral film coolant coverage away of the hole exit. The conical jet hole enhanced the average streamwise adiabatic film cooling effectiveness by 11.11% and 123.2% at BR= 0.5 and 1.0, respectively, while in the averaged lateral adiabatic in the spanwise direction, the film cooling effectiveness enhanced by 61.75% and 192.6% at BR= 0.5 and 1.0, respectively

Film cooling effectiveness distribution on first-stage vane endwall with and without leading-edge fillets

Considered are the effects of leading edge airfoil geometry on endwall film cooling. Three different fillet profiles, with length to height ratios of 2.8, 1.2 and 0.5, are considered, as each is placed at the junction of the airfoil leading edge and endwall. The results of these arrangements are compared with a baseline configuration with no fillet. The film cooling is produced by four rows of compound-angle, laidback fan-shaped holes, where each row of holes is located in the lateral direction. Each fanshaped hole is inclined at an angle of 30° with respect to the endwall surface. The compound angles of the first, second, third, and fourth rows of holes are 0°, 30°, 45°, and 60°, respectively, relative to the axial direction. Each fanshaped hole has a lateral diffusion angle of 10° from the hole-centerline, and a forward expansion angle of 10° relative to the endwall surface. The Reynolds number based on the axial chord and inlet velocity of the free-stream flow is 3.5×105, and the testing is done in a four-blade cascade with low Mach number condition (0.1 at the inlet), while the overall-average blowing ratio of the coolant through the discrete holes is constant at 0.4. Local, surface adiabatic film-cooling effectiveness values are measured using pressure sensitive paint (PSP). For a blowing ratio of 0.4, the baseline geometry (with no fillet) provides the best film cooling performance near leading edge pressure side. As for the leading edge suction side, the best leading edge configuration is the longfillet arrangement, since it is more effective in controlling the suction side leg of the horseshoe vortex compared to the other configurations which are examined.

Turbulent Heat Transfer and Large Coherent Structures in Trailing-edge Cutback Film Cooling

Film cooling is a key technology for improving the thermal efficiency and power output of gas turbines. The trailing-edge section of high-pressure turbine blades can be efficiently cooled by ejecting a film over a cutback on the pressure side of the blade. In this paper, results of Large–Eddy Simulations (LES) are presented that match an existing experimental setup. Altogether, eight simulations with the blowing ratio M varying as the only parameter were performed over a range from M = 0.35 to 1.4. Reasonably good agreement between LES and experiments were obtained for flow field statistics and adiabatic film-cooling effectiveness ηaw. Within a limited range of blowing ratios, an increase in the blowing ratio results in a counter-intuitive decrease of the cooling effectiveness. The present work suggests a mechanism that can explain this behavior. The visualization and analysis of large coherent structures showed that there exists dominant clockwise-rotating structures that can give rise to a combined upstream- and wall-directed turbulent heat flux. This turbulent heat flux represents the main contribution of the total heat flux and causes a significantly intensified thermal mixing process, which in turn results in the counter-intuitive decrease of the cooling effectiveness.

Impact of Different Film-Cooling Modes at Trailing Edge on the Aerodynamic Performance of Heavy Duty Gas Turbine Cascade

In this paper, the trailing edge film cooling flow field of a heavy duty gas turbine cascade has been studied by central difference scheme and multi-block grid technique. The research is based on the three-dimensional N-S equation solver. By way of analysis of the temperature field, the distribution of profile pressure, and the distribution of film-cooling adiabatic effectiveness in the region of trailing edge with different cool air injection mass and different angles, it is found that the impact on the film-cooling adiabatic effectiveness is slightly by changing the injection mass. The distribution of profile pressure dropped intensely at the pressure side near the injection holes line with the large mass cooling air. The cooling effect is good in the region of trailing edge while the injection air is along the direction of stream.

Influence of Internal Cyclone Flow on Adiabatic Film Cooling Effectiveness

Influence of Internal Cyclone Flow on Adiabatic Film Cooling Effectiveness

Large-Eddy Simulations of trailing-edge cutback film cooling at low blowing ratio

For high-pressure turbine blades, an efficient cooling of the trailing edge can be achieved by ejecting a film into the flow over a cutback on the pressure-side of the blade. Here, results of well-resolved Large-Eddy Simulations (LES) are reported that match an existing experimental setup. A low blowing ratio M = 0.5 was chosen and compared to results for an engine-typical value of 1.1. LES and experiments agreed reasonably well for mean and r.m.s. velocity profiles and adiabatic film-cooling effectiveness η aw. By imposing different flow conditions in the coolant channel, the LES data show that the flow regime of the coolant at ejection has a significant impact on the performance of the resulting cooling film for both blowing ratios. There are two surprising results: (1) a counter-intuitive behavior causing an also experimentally observed increase in η aw for a reduced blowing ratio. (2) For M = 0.5, a turbulent coolant sustains a higher cooling effectiveness farther downstream compared to a laminar coolant, whereas, for M = 1.1, the opposite is observed. It is shown that both phenomena are related to a change in type and strength of the dominant coherent structures that are formed behind the cutback lip and convected downstream along the trailing edge.

Computational Analysis of Surface Curvature Effect on Mist Film-Cooling Performance

Air-film cooling has been widely employed to cool gas turbine hot components, such as combustor liners, combustor transition pieces, turbine vanes, and blades. Studies with flat surfaces show that significant enhancement of air-film cooling can be achieved by injecting water droplets with diameters of 5–10 μm into the coolant airflow. The mist/air-film cooling on curved surfaces needs to be studied further. Numerical simulation is adopted to investigate the curvature effect on mist/air-film cooling, specifically the film cooling near the leading edge and on the curved surfaces. Water droplets are injected as dispersed phase into the coolant air and thus exchange mass, momentum, and energy with the airflow. Simulations are conducted for both 2D and 3D settings at low laboratory and high operating conditions. With a nominal blowing ratio of 1.33, air-only adiabatic film-cooling effectiveness on the curved surface is lower than on a flat surface. The concave (pressure) surface has a better cooling effectiveness than the convex (suction) surface, and the leading-edge film cooling has the lowest performance due to the main flow impinging against the coolant injection. By adding 2% (weight) mist, film-cooling effectiveness can be enhanced approximately 40% at the leading edge, 60% on the concave surface, and 30% on the convex surface. The leading edge film cooling can be significantly affected by changing of the incident angle due to startup or part-load operation. The film cooling coverage could switch from the suction side to the pressure side and leave the surface of the other part unprotected by the cooling film. Under real gas turbine operating conditions at high temperature, pressure, and velocity, mist-cooling enhancement could reach up to 20% and provide a wall cooling of approximately 180 K.

Increasing Adiabatic Film-Cooling Effectiveness by Using an Upstream Ramp

A new design concept is presented to increase the adiabatic effectiveness of film cooling from a row of film-cooling holes. Instead of shaping the geometry of each hole; placing tabs, struts, or vortex generators in each hole; or creating a trench about a row of holes, this study proposes a geometry modification upstream of the holes to modify the approaching boundary-layer flow and its interaction with the film-cooling jets. Computations, based on the ensemble-averaged Navier–Stokes equations closed by the realizable k‐ε turbulence model, were used to examine the usefulness of making the surface just upstream of a row of film-cooling holes into a ramp with a backward-facing step. The effects of the following parameters were investigated: angle of the ramp (8.5deg, 10deg, 14deg), distance between the backward-facing step and the row of film-cooling holes (0.5D,D), blowing ratio (0.36, 0.49, 0.56, 0.98), and “sharpness” of the ramp at the corners. Results obtained show that an upstream ramp with a backward-facing step can greatly increase surface adiabatic effectiveness. The laterally averaged adiabatic effectiveness with a ramp can be two or more times higher than without the ramp by increasing upstream and lateral spreading of the coolant.

Three-Dimensional Flow Prediction and Improvement of Holes Arrangement of a Film-Cooled Turbine Blade Using a Feature-Based Jet Model

A feature-based jet model has been proposed for use in three-dimensional (3D) computational fluid dynamics (CFD) prediction of turbine blade film cooling. The goal of the model is to be able to perform computationally efficient flow prediction and optimization of film-cooled turbine blades. The model reproduces in the near-hole region the macroflow features of a coolant jet within a Reynolds-averaged Navier-Stokes framework. Numerical predictions of the 3D flow through a linear transonic film-cooled turbine cascade are carried out with the model, with a low computational overhead. Different cooling holes arrangements are computed, and the prediction accuracy is evaluated versus experimental data. It is shown that the present model provides a reasonably good prediction of the adiabatic film-cooling effectiveness and Nusselt number around the blade. A numerical analysis of the interaction of coolant jets issuing from different rows of holes on the blade pressure side is carried out. It is shown that the upward radial migration of the flow due to the passage secondary flow structure has an impact on the spreading of the coolant and the film-cooling effectiveness on the blade surface. Based on this result, a new arrangement of the cooling holes for the present case is proposed that leads to a better spanwise covering of the coolant on the blade pressure side surface.

Experimental and Computational Comparisons of Fan-Shaped Film Cooling on a Turbine Vane Surface

The flow exiting the combustor in a gas turbine engine is considerably hotter than the melting temperature of the turbine section components, of which the turbine nozzle guide vanes see the hottest gas temperatures. One method used to cool the vanes is to use rows of film-cooling holes to inject bleed air that is lower in temperature through an array of discrete holes onto the vane surface. The purpose of this study was to evaluate the row-by-row interaction of fan-shaped holes as compared to the performance of a single row of fan-shaped holes in the same locations. This study presents adiabatic film-cooling effectiveness measurements from a scaled-up, two-passage vane cascade. High-resolution film-cooling measurements were made with an infrared camera at a number of engine representative flow conditions. Computational fluid dynamics predictions were also made to evaluate the performance of some of the current turbulence models in predicting a complex flow such as turbine film-cooling. The renormalization group (RNG) k‐ε turbulence model gave a closer prediction of the overall level of film effectiveness, while the v2‐f turbulence model gave a more accurate representation of the flow physics seen in the experiments.

Numerical simulation of film-cooling concave plate as coolant jet passes through two rows of holes with various orientations of coolant flow

Computations are performed to predict the three-dimensional flow and heat transfer of concave plate that is cooled by two staggered rows of film-cooling jets. This investigation considers two coolant flow orientations: (1) the coolant jets were ejected from a straight-blow supply plenum, so the coolant supply plane is parallel to the entrance plane of the coolant jets; (2) the coolant jets were ejected from the cross-blow supply plenum so the coolant supply plane was normal to the entrance plane of the coolant jets. The effects of numerous film-cooling parameters were investigated, including the mainstream Reynolds number, the angular locations of the two-row injections and the blowing ratio. The mainstream Reynolds number, determined by the diameter of the injection hole as the characteristic length, varied from 3440 to 13,760. The blowing ratio ranged from 0.5 to 2.0 with a fixed density ratio of 1.14. Additionally, two angles of injections, 40° and 42°, from the exit plane of the entrance duct are considered. Results are presented as the surface adiabatic film-cooling effectiveness, the temperature distribution and the velocity vector profile. The formation and trace of counter-rotating vortex pairs that result from the interaction between the mainstream hot gas and the cooling jets was clearly exhibited. The laterally averaged film-cooling effectiveness over the concave surface with a straight-blow plenum is slightly higher than that of a cross-blow plenum at all test blowing ratios. Results of this study demonstrate that the blowing ratio is one of the most significant film-cooling parameters over a concave surface.

Heat Transfer and Film-Cooling Measurements on a Stator Vane With Fan-Shaped Cooling Holes

In a typical gas turbine engine, the gas exiting the combustor is significantly hotter than the melting temperature of the turbine components. The highest temperatures in an engine are typically seen by the turbine inlet guide vanes. One method used to cool the inlet guide vanes is film cooling, which involves bleeding comparatively low-temperature, high-pressure air from the compressor and injecting it through an array of discrete holes on the vane surface. To predict the vane surface temperatures in the engine, it is necessary to measure the heat transfer coefficient and adiabatic film-cooling effectiveness on the vane surface. This study presents heat transfer coefficients and adiabatic effectiveness levels measured in a scaled-up, two-passage cascade with a contoured endwall. Heat transfer measurements indicated that the behavior of the boundary layer transition along the suction side of the vane showed sensitivity to the location of film-cooling injection, which was simulated through the use of a trip wire placed on the vane surface. Single-row adiabatic effectiveness measurements without any upstream blowing showed jet lift-off was prevalent along the suction side of the airfoil. Single-row adiabatic effectiveness measurements on the pressure side, also without upstream showerhead blowing, indicated jet lifted-off and then reattached to the surface in the concave region of the vane. In the presence of upstream showerhead blowing, the jet lift-off for the first pressure side row was reduced, increasing adiabatic effectiveness levels.

Fan-Shaped Hole Effects on the Aero-Thermal Performance of a Film-Cooled Endwall

The present paper investigates the effects of a fan-shaped hole endwall cooling geometry on the aero-thermal performance of a nozzle vane cascade. Two endwall cooling geometries with four rows of holes were tested, for different mass flow rate ratios: the first configuration is made of cylindrical holes, whereas the second one features conical expanded exits and a reduced number of holes. The experimental analysis is mainly focused on the variations of secondary flow phenomena related to different injection rates, as they have a strong relationship with the film cooling effectiveness. Secondary flow assessment was performed through downstream 3D aerodynamic measurements, by means of a miniaturized 5-hole probe. The results show that at high injection rates, the passage vortex and the 3D effects tend to become weaker, leading to a strong reduction of the endwall cross flow and to a more uniform flow in spanwise direction. This is of course obtained at the expense of a significant increase of losses. The thermal behavior was then investigated through the analysis of adiabatic effectiveness distributions on the two endwall configurations. The wide-banded thermochromic liquid crystals (TLC) technique was used to determine the adiabatic wall temperature. Using the measured distributions of film-cooling adiabatic effectiveness, the interaction between the secondary flow vortices and the cooling jets can be followed in good detail all over the endwall surface. Fan-shaped holes have been shown to perform better than cylindrical ones: at low injection rates, the cooling performance is increased only in the front part of the vane passage. A larger improvement of cooling coverage all over the endwall is attained with a larger mass flow rate, about 1.5% of core flow, without a substantial increase of the aerodynamic losses.

Systematic Investigation on Conjugate Heat Transfer Rates of Film Cooling Configurations

For the determination of the film‐cooling heat transfer, the design of a turbine blade relies on the conventional determination of the adiabatic film‐cooling effectiveness and heat transfer conditions for test configurations. Thus, additional influences by the interaction of fluid flow and heat transfer and influences by additional convective heat transfer cannot be taken into account with sufficient accuracy. Within this paper, calculations of a film‐cooled duct wall and a film‐cooled real blade with application of the adiabatic and a conjugate heat transfer condition have been performed for different configurations. It can be shown that the application of the conjugate calculation method comprises the influence of heat transfer within the cooling film. The local heat transfer rate varies significantly depending on the local position.

Assessment of Various Film-Cooling Configurations Including Shaped and Compound Angle Holes Based on Large-Scale Experiments

The demand of improved thermal efficiency and high power output of modern gas turbine engines leads to extremely high turbine inlet temperatures and pressure ratios. Sophisticated cooling schemes including film cooling are widely used to protect vanes and blades from failure and to achieve high component lifetimes. Besides standard cylindrical cooling hole geometry, shaped injection holes are used in modern film cooling applications in order to improve cooling performance and to reduce the necessary cooling air flow. However, complex hole shapes may lead to manufacturing constraints and high costs. This paper evaluates some film-cooling injection geometry with different complexity. The comparison is based on measurements of the adiabatic film-cooling effectiveness and the heat transfer coefficient downstream of the injection location. In total, four different film-cooling hole configurations are investigated: a single row of fanshaped holes with and without a compound injection angle, a double row of cylindrical holes and a double row of discrete slots both in staggered arrangement. All holes are inclined 45 deg with respect to the model’s surface. During the measurements, the influence of coolant blowing ratio is determined. Additionally, the influence of cooling air feeding direction into the fanshaped holes with the compound injection angle is investigated. An infrared thermography measurement system is used for highly resolved mappings of the model’s surface temperature. Accurate local temperature data is achieved by an in-situ calibration procedure with the help of single thermocouples embedded in the test plate. A subsequent finite elements heat conduction analysis takes three-dimensional heat fluxes inside the test plate into account.

A Novel Transient Heater-Foil Technique for Liquid Crystal Experiments on Film-Cooled Surfaces

A novel transient measurement technique has been developed for determining the heat transfer characteristics in the presence of film cooling (heat transfer coefficient and adiabatic film-cooling effectiveness). The method is based on a transient heater foil technique, where a non-homogeneous surface heat flux is applied to the test surface. A regression analysis of multiple transient liquid crystal experiments is used to obtain the heat transfer characteristics. The method introduced here has the advantage that the (often not known) heat flux distribution at the surface is not needed for the analysis of the measured data. The method is used to study the influence of several heater foil configurations on a flat plate with film cooling, elucidating the effect of different thermal boundary conditions on film-cooling performance. The obtained data is also compared to results presented in literature and good agreement is found.

Correlation of Film-Cooling Effectiveness From Thermographic Measurements at Enginelike Conditions

Adiabatic film-cooling effectiveness on a flat plate surface downstream of a row of cylindrical holes is investigated. Highly resolved two-dimensional surface data were measured by means of infrared thermography and carefully corrected for local conduction and radiation effects. These locally acquired data are laterally averaged to give the streamwise distributions of the effectiveness. An independent variation of the flow parameters blowing rate, density ratio, and turbulence intensity as well as the geometrical parameters streamwise ejection angle and hole spacing is examined. The influences of these parameters on the lateral effectiveness is discussed and interpreted with the help of surface distributions of effectiveness and heat transfer coefficients presented in earlier publications. Besides the known jet in cross-flow behavior of coolant ejected from discrete holes, these data demonstrate the effect of adjacent jet interaction and its impact on jet lift-off and adiabatic effectiveness. In utilizing this large matrix of measurements the effect of single parameters and their interactions are correlated. The important scaling parameters of the effectiveness are shaped out during the correlation process and are discussed. The resulting new correlation is designed to yield the quantitatively correct effectiveness as a result of the interplay of the jet in crossflow behavior and the adjacent jet interaction. It is built modularly to allow for future inclusion of additional parameters. The new correlation is valid without any exception within the full region of interest, reaching from the point of the ejection to far downstream, for all combinations of flow and geometry parameters.

Film-Cooled Turbine Endwall in a Transonic Flow Field: Part II—Heat Transfer and Film-Cooling Effectiveness

Thermodynamic and aerodynamic measurements were carried out in a linear turbine cascade with transonic flow field. Heat transfer and adiabatic film-cooling effectiveness resulting from the interaction of the flow field and the ejected coolant at the endwall were measured and will be discussed in two parts. The investigations were performed in the Windtunnel for Straight Cascades (EGG) at the DLR, Goettingen. The film-cooled NGV endwall was operated at representative dimensionless engine conditions of Mach and Reynolds number Ma2is=1.0 and Re2=850,000 respectively. Part II of the paper discusses the thermodynamic measurements. Detailed temperature measurements were carried out on a turbine stator endwall. In order to determine the surface heat transfer and adiabatic film-cooling effectiveness, an infrared camera system in a rectilinear wind tunnel measured the endwall temperatures in the transonic flow field. The adiabatic film-cooling effectiveness was determined using the superposition method. Measurements were carried out to first, validate the assumptions of theory and second, analyze the error associated with the measurements. Effects of coolant ejection from a slot and three rows of holes were investigated and compared with each other. The influence of Mach number and blowing ratio on the heat transfer were further aspects of the investigation. Strong variations in heat transfer and film-cooling effectiveness due to the interaction of the coolant air and the secondary flow field were found. Based on the results of this investigation an improved cooling design will be investigated in a follow-up project.