Cooling Effectiveness Distributions Research Articles

Film cooling performance evaluation of bi-directional diffusion hole with compound-angle on turbine blade: Study on spanwise width

A bi-directional diffusion hole configuration, which expands along the flow direction and blade tip with a compound-angle, is proposed based on an actual turbine blade in the present study. The RKE turbulence models were adjusted to enhance simulation accuracy by correcting the turbulent viscosity coefficient, significantly improving the prediction of jet diffusion in stagnation areas. The film cooling effectiveness distribution was studied through pressure-sensitive paint (PSP) experimental measurements. The effect of the hole outlet spanwise width on the bi-directional diffusion hole film cooling performance is characterized by a significant influence, non-monotonicity and weak coupling. The optimal hole outlet spanwise width of bi-directional diffusion holes is explored. The dimensionless temperature distribution and vortex structure of the fluid in the hole outlet section reveal the intrinsic mechanism of the effect of the hole outlet spanwise width on the film cooling characteristics. The numerical simulation results and experimental validation results show that the optimal hole outlet spanwise width of the bi-directional diffusion hole is approximately 4.3D with the goal of achieving the optimal level of film cooling effectiveness. Considering the flow blending situation, the optimal hole outlet spanwise width is determined by balancing the increase in film cooling effectiveness caused by the coolant film spreading covering mode and the decrease in film cooling effectiveness caused by mainstream intrusion. Therefore, the engineering applicability of the optimal hole outlet spanwise width as a design criterion for bi-directional diffusion holes of the turbine blade is demonstrated.

Effect of the hole configurations on effusion cooling effectiveness under swirl impact in gas turbine combustor

Effusion cooling characteristics of the cylindrical and fan-shaped hole configurations are studied under realistic swirl flows at blowing ratios ranging from 1.2 to 6.0. RANS computations with the k-ω SST model are used to evaluate the interaction between swirl mainstream and cooling air. The results show that the cooling effectiveness distribution for the cylindrical and fan-shaped hole configurations are similarly controlled by swirl impact. Two high-temperature regions emerge near the impact location of the swirl main flow on the liner wall. The fan-shaped hole configuration has higher cooling effectiveness, and the difference is relative to location. Quantitatively analyzing, the fan-shaped holes are 19.7 %–53.2 % higher than the cylindrical holes in impact zones. In the corner recirculation zone, the difference ranges from 39.1 % to 84.2 %. The computations reflect the interaction between swirl flows and cooling jets is stronger for fan-shaped holes due to lower outlet velocity. Therefore the cooling air is easier to be suppressed by swirl impact under low BR, while the increasing blowing ratio can enhance the resistance of cooling air against swirl flows.

Study on effects of inlet nonuniformities on aerothermal performance of NGV pressure side with different film cooling designs

High pressure (HP) turbine works under nonuniform inlet conditions produced mainly by modern lean burn combustors. This article employed the conjugate heat transfer method to investigate the influences of nonuniform inflow conditions and film hole arrangements on the film coolant coverage and heat transfer characteristics. The researched region is on the pressure side (PS) of the nozzle guiding vane (NGV). A scaled-up model of the GE-E3 engine HP turbine stator vane was used and the film cooling performances of four different film hole arrangement schemes on the PS were discussed. In addition, the results of four nonuniform inflow conditions, including positive/negative swirl aligned with the stator vane leading edge (case PSW-LE and NSW-LE), and positive/negative swirl aligned with the stator vane mid passage (case PSW-PA and NSW-PA), were compared with the uniform inlet condition. The results indicate that the temperature distributions of stator vane PS are dominated by the film hole arrangements and the nonuniform inlet conditions. In the positive swirl (PSW) influenced region, film cooling effectiveness can be improved if the coolant orientation is consistent with the swirl direction. The film cooling effectiveness can be improved by 12%–20% under the Leading Edge (LE) clocking position and 45% under the Passage (PA) clocking position. In contrast, the negative swirl dominates the film cooling effectiveness distribution. Under the PA clocking position, the highest film cooling effectiveness can be enhanced by 45% compared with the LE. In contrast, the design of film holes has a weak contribution to the coolant outflow direction under the negative swirl (NSW) condition. However, the film holes with negative angles can enhance the cooling effectiveness behind the 13th row holes.

TPIV Experimental Investigation of Film Coolant-to-Mainstream Interaction From Shaped Cooling Holes With Various Inlet Geometries

Abstract In this investigation, multiple sets of time-averaged tomographic particle imaging velocimetry measurements are completed for laid-back, fan-shaped film cooling holes with “racetrack” shaped inlets. Traditional 10-10-10, laid-back, fan-shaped holes are inclined 30 deg to the mainstream flow on a flat plate. The inlet cross section varies from round to two elongated racetrack shapes. The cross-sectional area and the outlet-to-inlet area ratio for all the geometries are held constant. The flat plate is installed in a low-speed wind tunnel with a mainstream turbulence intensity of 8% and an average velocity of 21.6 m/s. The blowing ratios of the film jets range from 0.6 to 1.5 and the density ratio is 1. The Reynolds number of the cooling jet varies from 2600 to 8400. The characteristics of the resulting flowfield are coupled with the detailed film cooling effectiveness distributions. It can be noted from the results that the counter-rotating vortex pair generated by the 2:1 inlet is the closest to the surface and weakest in strength, likely caused by the minimum peak jet momentum of the three. The Reynolds stresses downstream of the 2:1 and 4:1 inlets are significantly lower than those downstream of the shaped hole with a round inlet. An inverse relation between volumetric turbulence accumulation (TA), and surface effectiveness (η), can be correlated for the blowing ratios considered. The turbulence accumulation term can thus be used to evaluate the performance of a film cooling hole design with flowfield data.

Improved Spatiotemporal Distributions of Sweeping Jet Film Cooling With a Shaped Exit Design

Abstract Sweeping jet (SJ) is a self-excited flow control device that can generate continuous and periodic flow, which makes it a promising geometry for film cooling. However, the cooling performance of current SJ configurations is not significantly better than that of 777-shaped holes. This paper presents two modified SJ configurations, named shaped SJ1 and shaped SJ2, which possess distinct shaped exit geometries. The spatiotemporal coolant distributions were comprehensively quantified for the 777-shaped hole, the compact SJ (with compact geometry), the shaped SJ1, and the shaped SJ2. The instantaneous cooling effectiveness distributions of the four configurations at blowing ratios (M) ranging from 0.5 to 3.0 were measured using the fast pressure-sensitive paint technique. The shaped SJ2 integrated the cooling characteristics of the 777-shaped hole and SJ. It maintained stability at a relatively low M, while began to oscillate when M was increased above 1.5. The shaped SJ design exhibited superior cooling effectiveness at both low and high M of 0.5 to 3.0. A proper orthogonal decomposition analysis and simulations illustrated the oscillation processes and flow structures within the SJ configuration. The shaped SJ2 showed a reduced outlet momentum and a more even distribution of film coverage than the 777-shaped hole, resulting in considerably improved cooling performance.

Effects of blowing ratio on surface heat transfer characteristics of a transonic turbine squealer blade tip with a double plenum film cooling supply arrangement

Investigated are the effects of blowing ratio on the surface heat transfer characteristics of a film cooled transonic turbine squealer blade tip with a unique double plenum supply arrangement for the coolant. The present configuration is unique because the forward film cooling holes are supplied by an upstream plenum, while the aft film cooling holes are supplied separately by a second downstream plenum. Considered are varying forward blowing ratios as the aft blowing ratio is maintained constant. Nine total blowing ratio combinations are investigated, such that the forward blowing ratio is varied from 1.89 to 3.07, whereas the aft blowing ratio is varied from 0.96 to 1.45. Surface data are provided along the upper pressure side surface and the squealer tip surface of the turbine blade. Included are spatially-resolved and line-averaged distributions of heat transfer coefficient ratio and adiabatic film cooling effectiveness, which vary in a complex manner as upstream and downstream blowing ratios are altered. Local adiabatic film cooling effectiveness distributions evidence good film coverage along the squealer tip surface for all considered blowing ratios, especially for portions of the trailing edge, recess, and pressure side rim regions. Results indicate that the optimal blowing ratio associated with holes supplied by the upstream plenum depends upon the blowing ratio value associated with holes supplied by the downstream plenum.

Experimental investigation of full-coverage film cooling characteristic of single/double-wall cooling vane

This paper employed the pressure sensitive paint (PSP) technique to measure the adiabatic film cooling effectiveness distributions of both the single-wall cooling vane (SCV) and the double-wall cooling vane (DCV). The experiments were implemented in an annual cascade under a Reynolds number of 3.0 × 105 condition. Furthermore, the effects of mass flow ratio (MFRs include 4.99 %, 6.65 %, 8.32 %, 9.98 %, and 11.65 %) and density ratio (DRs include 1.0, 1.5, and 2.0) on the adiabatic effectiveness distribution were also investigated. The results reveal that the DCV improves dramatically the film cooling effectiveness of the pressure side and the suction side compared with the SCV, and the effectiveness is elevated by 12 % and 25 %, respectively. As the MFR increases, the lateral average film cooling effectiveness of the SCV initially increases and then decreases, while that of the DCV enhances monotonously. Additionally, when the DR is raised from 1.0 to 1.5, the lateral average film cooling effectiveness of the pressure side of the SCV is increased by 50 % at the MFR = 6.65 %. Also, the DCV presents both high and even film cooling effectiveness contour distribution at high density ratio conditions, and the surface average film cooling effectiveness of the DCV is greater than 0.4 when the MFR ≥ 8.32 %.

Assessment of Binary Pressure-Sensitive Paint for Temperature and Heat Transfer Coefficient Measurement of Leading-Edge Film Cooling

Abstract Film cooling is a common technique for protecting gas turbine components from the hot combustor exhaust. Highly resolved film cooling effectiveness distributions are often obtained by measuring the mass transfer of a foreign gas coolant in mainstream air using pressure-sensitive paint (PSP). However, PSP is not able to measure the heat transfer coefficient, which is necessary to fully quantify the impact of film cooling. Instead, binary pressure-sensitive paint (BPSP) has an additional luminophore that is sensitive to temperature and can be used to measure the heat transfer coefficient. In this experiment, the film cooling effectiveness and heat transfer coefficient were measured using BPSP on the leading edge of a cylinder. The cylinder had a 7.62-cm diameter with two rows of cooling holes at ±15 deg from the leading edge. Each row contained ten holes with a 0.475-cm diameter, spaced 4 diameters apart in the spanwise direction, and angled 30 deg from the cylinder axis. The mainstream Reynolds number was 100,000 based on cylinder diameter with a turbulence intensity of 7.1%. The coolant-to-mainstream density ratio was 1.0, and the blowing ratio was 0.8. The heat transfer coefficient was measured in a transient heat transfer experiment using the reference signal from the BPSP. Despite the high uncertainty of the measurement, ranging from 24.0% to 71.1%, the results demonstrate the feasibility of the method and identify the best test methodology to minimize conduction errors.

Effect of Upstream Leakage Flow on Film Cooling Characteristic of a Turbine Convex Endwall

Abstract This paper investigates the effect of upstream leakage coolant flow on the film cooling performance of the convex endwall. The experiment is conducted in an annular passage with four vanes, and the endwall film cooling effectiveness distribution is measured by the pressure-sensitive paint technique. The near-wall flow field distribution predicted by computational fluid dynamics is employed to gain a deeper understanding of the influence of the passage secondary flow on the film cooling characteristic of the endwall. Furthermore, the effect of mass flow ratios (MFR = 0.75%, 1.0%, 1.25%, and 1.5%), leakage slot inclination angles (α = 30 deg, 45 deg, and 60 deg), density ratios (DR = 1.0, 1.5, and 2.0), as well as Reynolds number (Re = 2.0 × 105, 3.0 × 105, and 4.0 × 105) on the endwall film cooling is also investigated. Results indicate that the mass flow ratio has a great influence on the endwall film cooling effectiveness distribution, and a larger uncooled region can be observed when MFR is less than 1.25% due to the cooling air more easily being entrained by the secondary flow near the endwall. Reducing the inclination angle of the leakage slot enhances the axial velocity component of the coolant flow, thereby weakening the secondary flow near the endwall and improving the film cooling effectiveness distribution of the endwall. As the density ratio increases, the jet momentum of the leakage flow is reduced, particularly for higher MFR cases, which results in a significant reduction in the film cooling performance of the endwall. Besides, thinning the incoming boundary layer of the passage reduces the strength and size of the secondary flow in the passage and improves the endwall adiabatic effectiveness distribution when the Reynolds number is increased.

Effects of dusting film cooling placement and configuration on surface heat transfer characteristics of a transonic turbine squealer blade tip

Investigated are the effects of dusting hole film cooling placement and configuration on the surface heat transfer characteristics of a transonic turbine squealer blade tip. Three different blade tip arrangements are investigated, which are denoted C1, C2, and C3. Data are provided which include spatially-resolved and line-averaged distributions of heat transfer coefficient ratio and adiabatic film cooling effectiveness. The University of Alabama in Huntsville SS/TS/WT (supersonic/transonic/wind tunnel) blow down facility is employed with a transonic test section containing a five blade linear cascade for the investigation. Spatially-resolved surface heat transfer characteristics are measured using time- and spatially-resolved infrared thermography, in conjunction with an impulse response transient measurement procedure. Film coolant is supplied with a carbon dioxide injection system. Results indicate that heat transfer coefficient and adiabatic film cooling effectiveness distributions vary significantly as dusting hole placement and configuration are altered. For example, heat transfer coefficients for the C1 and C2 blade configurations are substantially lower, compared to baseline values with no film cooling, within the squealer recess region downstream of the dusting holes. Smaller heat transfer coefficient deviations, relative to baseline values, are present when the C3 configuration is employed. In general, distributions and magnitudes of local adiabatic film cooling effectiveness indicate that all three dusting hole configurations provide reasonably good film coverage within portions of squealer recess regions. For downstream portions of the blades, the C1 and C3 configurations both provide good surface protection which are indicated by film cooling effectiveness values, whereas coverage provided by the C2 arrangement ends abruptly at the location of a structural rib which is located within the squealer recess region.

Coupling of Volumetric Flowfield and Surface Effectiveness Measurements for Flat-Plate Film Cooling With Cylindrical Holes Using Tomographic Particle Image Velocimetry and Pressure-Sensitive Paint

AbstractIn this study, a series of time-averaged tomographic particle image velocimetry (TPIV) measurements are completed for simple angle, cylindrical film cooling holes on a flat plate with an inclination angle of 30 deg and a diameter of 4 mm. The flat plate is installed in a low-speed wind tunnel with a mainstream turbulence intensity of Tu=8% and an average velocity of 21.6 m/s. The blowing ratios of the jet range from M = 0.3 to M = 1.5, and the density ratio of the jet is fixed at DR = 1. The Reynolds number of the cooling jet varies from 1700 to 8400. The repeatability and accuracy of the tomographic particle imaging velocimetry (TPIV) measurements are compared against the results obtained by other flowfield measurement techniques. The results of the TPIV measurements are presented in velocity and vorticity iso-surface distributions in 3D, as well as 2D planar slices of velocity, vorticity, turbulence intensity, and Reynolds Stress distributions within the measurement volume. The characteristics of the flowfield are coupled with the detailed film cooling effectiveness distribution obtained using the pressure-sensitive paint (PSP) technique. An inverse relation among volumetric turbulence accumulation (TA) and surface film effectiveness, η, can be correlated.

Effect of blowing ratio, rotation, and film hole row location on film cooling on the suction surface of a rotating turbine blade

Experiments and simulations were performed to study the single-row film cooling performance on a turbine blade suction surface. Effects of blowing ratio, rotation, film hole row location, and the combine effects of the above three factors on single-row film cooling performance were studied. Totally three injection row locations of 12.4% S, 23.2% S, and 34% S were taken into consideration, where S refers to the suction side streamwise length. Experiments were carried out under turbine speeds of 400rpm, 600rpm, and 800rpm respectively with average blowing ratios change in the region of 0.5 to 1.5. Rotational Reynolds number corresponding to the turbine speeds was 3.57 × 105, 5.36 × 105, and 7.15 × 105. Pressure sensitive paint (PSP) technique was employed in obtaining film cooling effectiveness distribution. To assist explain the experimental results, numerical simulations were performed to acquire the mainstream and coolant jet flow field and local blowing ratio of each film hole. Results showed that film traces on the suction surface are bent toward the blade mid-span, and the downstream jet deflects more. Increasing blowing ratio and rotational Reynolds number will slightly reduce the inward film deflection. The local blowing ratios at hole exits are not evenly distributed along streamwise and spanwise direction under the combine effects of the three factors, which result in different film cooling performance at different streamwise locations and different blade heights.

Research on performance predictions using single-hole film cooling based on PointNet

A PointNet-based data-driven neural network model is proposed, which takes the film hole geometry variables and flow conditions as inputs to reconstruct the adiabatic cooling effectiveness distribution. The model aims to realize rapid reconstruction of the film cooling effectiveness field under complex and variable working conditions with a more flexible data organizational form. The dataset is derived from numerical simulations of the jet under crossflow. Select unstructured grid nodes are used to form point clouds for network training. The PointNet architecture includes two modules to extract the global features of the input point cloud and calculate the adiabatic efficiency. The responsiveness of the model to different variables is evaluated from the effectiveness contours, centerline, and laterally averaged effectiveness plots. Furthermore, correlation analysis is used to evaluate the accuracy of model predictions. Over the entire dataset, the mean correlation coefficient is 0.99, indicating that the model has a satisfactory ability to reconstruct and predict the effectiveness field. The main contribution from the area around the film holes to the cooling effectiveness distribution is further confirmed via critical point analysis.

Experimental and numerical investigations on the unsteady flow and film cooling characteristics of the trailing edge cutback

The film cooling characteristics of the trailing edge cutback slot of a high pressure turbine were experimentally and numerically investigated. The unsteady flows have an important effect on the mixing process of the mainstream and coolant flow, resulting in deteriorated film performance. To illustrate the interaction between the unsteady flows and coolant injection, eight blowing ratios ranging from 0.3 to 1.2 were considered for the baseline cases. Pressure sensitive paint measurements were conducted to acquire the adiabatic film cooling effectiveness distribution. The counter-intuitive decrease of the film cooling effectiveness was confirmed and further revealed. Particle image velocimetry showed that the unsteady flow patterns downstream of the trailing edge slot varied with increasing blowing ratio. Furthermore, the high-fidelity RANS/LES hybrid method was employed to resolve the detailed flow fields. The numerical method was carefully validated, and it successfully predicted the unsteady flows and mixing process. The spanwise vorticity showed that the two branches of the shedding vortices alternatively become strong. At the blowing ratio of 0.75, the strength of the two branches was comparative, yielding the worst cooling performance. Additionally, the density ratio and lip thickness to slot height ratio were investigated. Thus, the density ratio affects the coolant velocity ratio, causing the forward movement of the inflection point of the film cooling effectiveness map. The decrease of the lip thickness to slot height ratio leads to the reduction of the strength of the shedding vortices and therefore increases the film cooling effectiveness.

Numerical investigation on transpiration cooling performance with different porosities and mainstream pressure gradients

Improving transpiration cooling performance is a challenge to meet the requirements for the cooling performance and thermal stress restrict. In the current study, numerical simulations are carried out to investigate the transpiration cooling performance with different porosity configurations and mainstream pressure gradients. Eight porosity configurations, three mainstream pressure gradients, and three injection ratios are considered. Cooling effectiveness distributions and cooling effectiveness uniformities are evaluated. Velocity distributions and coolant allocations are also studied to reveal the cooling mechanism. Results show that all the transpiration cooling cases gain higher cooling effectiveness than the film cooling cases, and the maximum enhancement reaches up to 100%. However, the transpiration cooling cases have non-uniformity distributions of cooling effectiveness due to the uneven coolant allocations caused by the superposition effects. For uniform porosity configuration cases, the cooling effectiveness increases monotonically with the increasing porosity due to the enhanced external coolant coverage, whereas the uniformity of cooling effectiveness decreases due to the more uneven coolant allocation distribution along the mainstream direction, especially for the configurations with large porosity. In addition, when streamwise pressure gradients exist in the mainstream, only the TC-3 case with a small uniform porosity and the TC-8 case with a streamwise decreasing porosity perform better robustness. Among them, the TC-8 case gains obvious advantages in engineering applications due to the highest cooling effectiveness and the smallest non-uniformity, which is caused by the better coolant coverage and more uniform coolant allocation.

Effects of pressure side film cooling hole placement and condition on adiabatic film cooling effectiveness characteristics of a transonic turbine blade tip

The effects of film cooling hole placement location along the upper pressure side of a transonic squealer are considered. The thermal performance of four different film cooling configurations; B1, B2, B3 and B4, are considered using the University of Alabama in Huntsville's SS/TS/WT (supersonic/transonic/wind tunnel) experimental facility and a simulated turbine blade row using a linear cascade. Surface-varying results are provided for both the squealer blade tip surface, and for the upper pressure side of the squealer blade. These results are given for blowing ratios ranging from 0.42 to 3.20 in the form of spatially-resolved and spatially-averaged adiabatic film cooling effectiveness distributions. Because local static pressure variations and gradients vary in a significant manner at different blade locations, coolant accumulations near surfaces (and the associated surface thermal protection and film effectiveness distributions) produced by the different film cooling configurations are vastly different along the squealer blade tip surface. The B1 and B3 film cooling configurations show that substantial magnitudes of adiabatic film cooling effectiveness are present along the pressure side rim and the suction side rim. Also present is a considerable region of locally increased film cooling effectiveness and a wider region of surface protection within the squealer recess region for the B2 film cooling configuration. The contributions are unique because of a deficit of experimental data for film cooled squealer blades which operate with transonic flow, and because the present investigation is the first to consider the effects of hole placement location for upper pressure side film cooling arrangements with transonic flow conditions.

Effects of plasma actuation and hole configuration on film cooling performance

In this paper, plasma actuators are arranged asymmetrically downstream the wall to improve film cooling performance. Effects of blowing ratio, hole configuration and applied voltage on flow characteristics and film cooling effectiveness were investigated numerically on a flat plate. Results show that highest film cooling effectiveness distribution is obtained both in the spanwise and streamwise directions under blowing ratio of 0.5. Average wall film cooling effectiveness of cylindrical hole increases by 251.9% under blowing ratio of 0.5 compared to that under blowing ratio of 1.5. The scale of the counter rotating vortex pairs (CRVP) from fan shaped hole and sister hole are significantly reduced compared to that from cylindrical hole. The console hole has an anti-counter rotating vortex pair (Anti-CRVP), which weakens the entrainment of the CRVP to the coolant air near the wall. Compared with the cylindrical hole, average wall film cooling effectivenesses for fan shaped hole, sister hole and console hole increase by 73.1%, 97.5% and 119.9%. The adherent performance of the coolant air is enhanced after applying plasma actuator. The aerodynamic actuation of the plasma results in the rebound of the fluid close to the wall at 24 kV applied voltage. Average wall film cooling effectiveness of the console hole at 12 kV applied voltage is 10.6% higher than that without plasma.

CFD analysis of combustion liner with different geometry holes for effusion film cooling

The advanced gas turbine is designed to operate at higher temperature to increase thermal efficiency. Since the gas temperature exceed the allowable material temperature, so cooling techniques of turbine components are more important. Effusion film cooling is an external and highly sophisticated cooling technology, which protects the liner by injecting the secondary cooler fluid at particular locations over the liner surface exposed to the higher gas temperatures. The injected fluid forms as a thin layer (film) on the liner surface and protects the liner surface from higher gas temperatures. In elliptical geometry of holes in plate for effusion cooling is studied for the samples made of adiabatic and conductivity material. To increase the life of combustion chamber it is essential to know the amount of cooling air to be used for cooling of the liner. For effusion film cooling, geometry of the hole is one of the key parameters which influences the coolant mass flow rate and it spread over the liner surface thus affecting the cooling effectiveness. The commercially available FloEFD software is used for calculating the surface temperature and effusion cooling effectiveness distribution throughout the sample plate and result values are in good agreement and this has given confidence to usage of the commercial software FloEFD for predicting the film effectiveness. After result validation finally concluded that span wise elliptical hole (Config-2) have more effectiveness than circular holes.

Experimental and numerical investigation of micro-scale effusion and transpiration air cooling on cascaded turbine blades

Micro cooling utilizing the micro-sized cooling hole or a porous structure on the surface is a promising technology in cooling applications of gas turbine blades. Although there have been numerous parametric studies on basic geometries to enhance the performance of micro cooling, little has been done for experimental study on blade geometry. This study comprehensively investigated micro cooling performance by applying effusion and transpiration cooling to C3X blades arranged in a cascade. The overall cooling effectiveness distribution on the blade surface was estimated by using infrared thermometry. In addition, the velocity and thermal boundary layer formation by cooling air were qualitatively investigated by a flow visualization using the smoke-laser sheet technique and numerical simulation using the shear stress transport k-ω turbulence model. Micro cooling performs effectively because of the convective heat transfer through the microstructure of the blade wall but also the reduction of heat transfer from the hot mainstream due to the formation of a uniform coolant layer on the blade surface. Especially at the mass flow ratio of 5.3% used for typical gas turbine cooling, effusion cooling and transpiration cooling achieve the overall cooling effectiveness of 0.4 and 0.6, respectively.

Deep learning method for fast prediction of film cooling performance

This study examines the predictive capability of deep learning method for adiabatic film cooling effectiveness distribution with variable operating conditions and geometric layouts. A conditional generative adversarial network is trained to establish nonlinear mapping from input to output. We embed the boundary condition information directly into the input tensor, thereby imparting the capability to address variable operating conditions. The processed input tensors include different blowing ratios M, incoming turbulence intensity Tu, geometry profile controlled by the inclination angle ϕ, compound angle θ, number of hole rows n, pitch between hole rows Px, and span-wise hole pitch Py. The output data are adiabatic film cooling effectiveness fields generated by Reynolds-averaged Navier–Stokes simulations. The prediction results are in good agreement with the computational fluid dynamics results in terms of various statistical assessments. Furthermore, compared with conventional methods of solving Navier–Stokes equations, predictions based on deep learning result in better response times. Therefore, the method proposed in this study is of high significance in the early design of cooling structures for turbine blades.