Film-cooling Holes Research Articles

Stress intensity factors analysis for crack around film cooling holes in Ni-based single crystal with contour integral method

PurposeThis paper aims to provide SIF calculation method for engineering application.Design/methodology/approachIn this paper, the stress intensity factors (SIFs) calculation method is applied to the anisotropic Ni-based single crystal film cooling holes (FCHs) structure.FindingsBased on contour integral, the anisotropic SIFs analysis finite element method (FEM) in Ni-based single crystal is proposed. The applicability and mesh independence of the method is assessed by comparing the calculated SIFs using mode of plate with an edge crack. Anisotropic SIFs can be calculated with excellent accuracy using the finite element contour integral approach. Then, the effect of crystal orientation and FCHs interference on the anisotropic SIFs is clarified. The SIFs of FCH edge crack in the [011] orientated Ni-based single crystal increases faster than the other two orientations. And the SIF of horizontal interference FCHs edge crack is also larger than that of the inclined interference one.Originality/valueThe SIFs of the FCH edge crack in the turbine air-cooled blade are innovatively computed using the sub-model method. Both the Mode I and II SIFs of FCHs edge crack in blade increase with crack growing.

Casting of film-cooling holes for single-crystal blades

The flawless formation of film-cooling holes (FCHs) is critical for manufacturing single-crystal blade. Presently, high-energy beam drilling is mostly used to process gas film holes, which introduces the limitations of complex hole structures that are difficult to form and processing thermal defects that are difficult to eliminate. This study demonstrates an approach to form FCHs using casting, where an indirect additive manufacturing technology using a core/shell integral ceramic mold is applied for casting. Cylindrical 0.4, 0.8, and 4.0 mm-diameters hole are cast. The influence of fine ceramic cores on the microstructure of single-crystal blades during directional solidification is investigated using simulations and experiments. The fine ceramic cores of the FCHs affect the growth morphology of dendrites around the hole, which is determined by the hole diameter, primary dendrite spacing, and local temperature filed. Owing to the thermal conductivity difference between the alloy and ceramic, the local temperature filed around the holes changes abruptly, leading to the formation of shrinkage porosity and cavities. Furthermore, the process window for the casting of FCHs is determined using the longitudinal temperature gradient (GZ) and solidification rate (VZ). The proposed approach is of practical significance for the defect-free casting of tiny structures in single-crystal blades.

Experimental investigations on flow characteristics of impingement/swirl cooling structures inside a blade leading edge

Experimental investigations were conducted to study the flow characteristics of impingement/swirl cooling structures inside a turbine-blade leading edge (LE) using particle image velocimetry. The Reynolds number (Re) was maintained at either 10 000 or 20 000. The effects of Re and the ratio of the impingement-cooling (IC) hole offset distance to the IC hole diameter (e/d) on the flow phenomena and the characteristics of the impingement/swirl cooling structures inside the blade LE were analyzed. The effects of the coolant-outflow mode were also taken into consideration. The results show that when e/d = 0, the diffusion flow of the impingement jet moves more to the side of the LE with a larger rounded corner, and this phenomenon is particularly obvious at Re = 20 000. Regardless of the e/d value, both the impingement-jet velocity and vorticity increase significantly at a higher Re value. The velocity of the impingement jet reaching the internal impingement target surface of the LE decreases significantly as e/d is changed from 1.5 to −1.5 with Re = 20 000. Generally speaking, regardless of the e/d value, the impingement jet can form a swirling vortex on the side with the larger rounded corner. Under the conditions of Re = 20 000 and e/d = 1.5, regardless of the coolant-outflow mode, the flow velocity of the impingement jet decreases monotonically along the flow direction, and this is more obvious inside the LE model without film-cooling holes.

Comparisons of Overall Performance Among Double-Jet Film Cooling Holes, Cylinder Holes, and Fan-Shaped Holes

Abstract In this paper, the film-cooling effectiveness (η) and heat transfer coefficient (h) of different film hole geometries are investigated, including double-jet film cooling (DJFC) holes, streamwise cylindrical holes, and fan-shaped holes, both experimentally and numerically. Results reveal that when the blowing ratio is less than 1.0, the DJFC holes have the highest η and the highest h, as well as the highest net heat flux reduction (NHFR). However, a higher blowing ratio (>1.0) leads to a quickly decreasing NHFR of DJFC holes. The asymmetric antikidney vortex and the high turbulent kinetic energy (TKE) are dominant in the performance of the DJFC holes. Owing to medium effectiveness and the lowest heat transfer coefficient, the fan-shaped holes possess the highest net heat flux reduction at M = 2.0 although the value is negative. The relatively weak kidney vortex and the low TKE can explain the phenomena. The cylindrical holes have the lowest η and the lowest NHFR due to the kidney vortex and relatively higher TKE.

Experimental and numerical investigations of vane endwall film cooling with different cooling hole configurations

In the present study, measurements are conducted to investigate endwall film cooling and its secondary effect on airfoil surfaces in a linear cascade. Validated numerical simulations are performed to obtain flow fields. Several configurations of upstream double-row cylindrical holes with an inclination angle (α) of 30 deg are studied, including streamwise holes, compound-angled holes, and double-jet film-cooling (DJFC) holes. The effects of blowing ratios ranging from 0.5 to 2.0 are also considered. Results indicate that at lower blowing ratios, strong secondary flows dominate the flow field and lead to weak coolant coverage on the endwall and the airfoil surfaces. The differences amongst the configurations are not obvious due to the insufficient coolant momentum. At higher blowing ratios, the film cooling performance on the endwall and the airfoil surfaces is greatly improved. Moreover, the effect of cooling configuration becomes more apparent as the blowing ratio increases, which determines the distribution of coolant coverage, especially on the airfoil surfaces.

Cooling performance of film-cooling holes fed by channels of various shapes

Internal and external cooling is pertinent to maintaining operable surface temperatures of hot-section gas turbine components. Film cooling on the external surfaces is supplied from cooling passages placed inside of the blades and can affect the external film-cooling performance. Assessing internal cooling designs that were once difficult to investigate because of the complexity and costs of manufacturing cast turbine blades is now possible to more quickly and less costly through additive manufacturing (AM).This paper provides results and analysis of an experimental and computational study of the cooling performance of meter-diffuser shaped film-cooling holes fed by a range of individual channel shapes with different cross-sectional shapes: circle, hexagon, pentagon, ellipse, diamond, square, rectangle, trapezoid, triangle – vertex up, and triangle – vertex down. These geometries were made possible through the use of AM. The perimeter, and consequently the surface area of the supply channels to the cooling holes, was maintained constant between the different channel shapes. The predictive computational results showed that the shape of the channel affects the secondary flows through the hole that in turn affects the overall film cooling as determined experimentally. The triangle – vertex up created flow structures in the diffuser-shaped cooling hole that caused the coolant jet to create a film on the surface with wide lateral spread; whereas the triangle – vertex down created flow through the cooling hole that caused the jet to fully detach at the same blowing ratio. Consequently, the triangle – vertex up had a 14% improvement in overall effectiveness over the triangle – vertex down and a 33% improvement over the circle which had the lowest overall effectiveness of all the tested shapes.

Unveiling particle deposition characteristics on flat plate with a shaped film cooling hole

Particle deposition on turbine blades poses a significant challenge to the safe operation of gas turbines. The blockage of film-cooling holes and alteration in the aerodynamic shape of blades affect the turbine's efficiency, highlighting the importance of reducing particle deposition on turbine blades. To investigate the characteristics and potential mechanisms of deposition, experimental and numerical studies were conducted on two typical film-cooling holes: cylindrical and 777-shaped. A multi-perspective scanning method was used to measure the three-dimensional deposition topography on flat plates under a particle-laden environment. The wax deposition experiments showed that the deposition thickness of both holes increased with the blowing ratio. However, the 777-shaped hole exhibited a 10–40% reduction in deposition compared with the cylindrical hole. The computational predictions of deposition patterns agreed well with the experimental results. The numerical simulations revealed that the presence of film cooling reduced deposition in some areas but increased deposition in the vicinity of the film coverage region owing to the entrainment of vortices. Overall, this study further elucidates particle deposition characteristics and the influencing factors, which can guide the design of blade cooling systems with reduced deposition.

Influence of Spatial Characteristic Parameters of Film Cooling Holes on the Strength of Turbine Blade

In order to improve the fatigue resistance of turbine blades with special-shaped film cooling holes in real working environments, the spatial geometric characteristics of the gas film holes are studied by using the numerical calculation method of gas-thermal-solid depth coupling, which is based on the spatial geometric characteristics of special-shaped film cooling holes. Then, the influence of stress level and durable strength at the hole edge is developed, which includes the parameters of the film jet angle, the flow direction divergence angle, the span-wise divergence angle, and the pore diameter on the film cooling holes. The results show that the flow direction expansion angle and the spanwise expansion angle are the most significant impacts on the hole edge stress at the same pore diameter and film jet angle. Compared with the cylindrical gas film hole, it is decreased by about 9.3% to 25.8%, which is the hole edge equivalent stress of the special-shaped gas film hole. The durable strength of the hole edge increases by about 11.1% to 25.2%.

Failure mechanism of single-crystal superalloy NiAlReRu components with film cooling holes under the coupling high-speed rotating and high temperature condition

In order to make up for the deficiency that the laboratory testing of standard samples cannot approximately simulate the turbine blade geometry and service environment, the creep experiment of the nickel-based single crystal superalloy NiAlReRu components with film cooling holes (FCHs) is carried out under the coupling high-speed rotating and high temperature condition. The component with the FCHs exposed to the coupling high-speed rotating of 20000 rpm and high temperature of 1000 ℃ continuously lasts 255.92 h until it breaks. The microstructural evolution and fracture process are analyzed. The results indicate that the existence of the FCHs evidently changes the stress distribution of the region with the FCHs, which makes the microstructural evolution of γ′ phase significantly different. The component is fractured along the second row of the FCHs and the normal direction of the fracture surface is basically parallel to the centrifugal force direction. The fracture process can be divided into three typical stages. The coupling effect of temperature and stress on the failure mechanism is discussed.

Experimental and Computational Investigation of Shaped Film Cooling Holes Designed to Minimize Inlet Separation

Abstract Film cooling is used to protect turbine components from the extreme temperatures by ejecting coolant through arrays of holes to create an air buffer from the hot combustion gases. Limitations in traditional machining meant film cooling holes universally have sharp inlets, which create separation regions at the hole entrance. The present study uses experimental and computational data to show that these inlet separation are a major cause of performance variation in crossflow fed film cooling holes. Three-hole designs were experimentally tested by independently varying the coolant velocity ratio (VR) and the coolant channel velocity ratio (VRc) to isolate the effects of crossflow on hole performance. Leveraging additive manufacturing (AM) technologies, the addition of a 0.25D radius fillet to the inlet of a 7-7-7 shaped hole is shown to significantly improve diffuser usage and significantly reduce variation in performance with VRc. A second AM design used a very large radius of curvature inlet to reduce biasing caused by the inlet crossflow. Experiments showed that this “swept” hole design did minimize biasing of the coolant flow to one side of the shaped hole, and it significantly reduced variations due to varying VRc. RANS simulations at six VR and three VRc conditions were made for each geometry to better understand how the new geometries changed the velocity field within the hole. The sharp and rounded inlets were seen to have very similar tangential velocity fields and jet biasing. Both AM inlets created more uniform, slower velocity fields entering the diffuser. The results of this article indicate that large improvements in film cooling performance can be found by leveraging AM technology.

Stress evaluation of TBCs with inclined film-cooling hole considering CMAS penetration-induced change of thermo-mechanical properties

CaO-MgO-Al2O3-SiO2 (CMAS) penetration is an essential factor leading to the failure of thermal barrier coatings (TBCs), especially around the film-cooling hole structure of high-temperature gas turbine components. The complex temperature distribution and the free-edge effect around the cooling holes aggravate the influence of CMAS penetration on the failure of the TBCs. In this study, a three-dimensional (3D) model of a TBC-film cooling system considering CMAS attack is presented based on the observation of a retired turbine blade. CMAS penetration-induced change of thermo-mechanical properties of the ceramic top coat (TC) was considered. The thermal gradient around a cooling hole and the uneven growth of the thermally grown oxide (TGO) on the stress distribution in TBCs were investigated. The results showed that the thermal insulation performance of TBCs declined considerably under CMAS penetration, and the induced maximum increase in the temperature at the TC/bond coat (BC)interface reached 60 K. CMAS penetration also led to a considerable stress increase in the TBC-film system. We therefore conclude that CMAS penetration accelerates the crack initiation at the TC/TGO and TGO/BC interfaces.

Experimental and numerical investigations of vane endwall film cooling with different density ratios

In the current paper, endwall film cooling and its secondary effect on vane surfaces in a linear cascade are measured using pressure sensitive paint (PSP) at different density ratios. Flow fields are obtained by a validated numerical method. Upstream double-row film cooling holes with an inclination angle (α) of 30 deg. are studied, including streamwise cylindrical holes, compound-angled cylindrical holes and double-jet film-cooling (DJFC) holes. The density ratio (DR) varies as 1.0, 1.5 and 2.5. Low- and high-blowing ratio conditions (M = 0.5 and 2.0) are also considered. Results reveal that, at a low blowing ratio, higher endwall cooling effectiveness is obtained as the DR increases, while the variation of density ratio shows a limited impact on the secondary cooling effectiveness of vane surfaces. At a high blowing ratio, with the increase of DR, the coolant coverage gradually shrinks, especially on the endwall and the vane pressure surface. Moreover, obvious differences in coolant coverage are observed among the investigated cooling configurations.

Revealing the Formation of Recast Layer around the Film Cooling Hole in Superalloys Fabricated Using Electrical Discharge Machining

A film cooling hole is an efficient and reliable cooling method, which is widely used in aeroengine turbine blades to effectively improve the thrust–weight ratio of the engine. Electrical discharge machining is the most common manufacturing process for film cooling holes. Due to the rapid quenching after high-temperature melting, a certain thickness of the recast layer will be formed in the vicinity of the hole wall. The microstructure of the recast layer is considered to be an important factor affecting the performance of single-crystal blades. Generally, the recast layer has been thought of as one of the main reasons for the failure of turbine blades. Accordingly, the formation of the recast layer is an important and interesting issue to be revealed. In this work, the recast layer formed using electrical discharge machining on a single-crystal superalloy is studied with TEM. It is found that the recast layer is in the state of supersaturated solution with a single-crystal structure epitaxially grown from the matrix, and many dislocations were observed therein.

Optimization of the Double-Expansion Film-Cooling Hole Using CFD

Film cooling is a major cooling technique used in modern gas turbines and air engines. The geometry of film-cooling holes is the fundamental aspect affecting the cooling performance. In this paper, a new cooling configuration called the double-expansion film-cooling hole has been put forward, which yields better performance than the widely used shaped holes and is easy to manufacture. The double-expansion holes at inclination angles of α=30∘, 45∘, and 60∘ are optimized using the genetic algorithm and the Kriging surrogate model, which is trained by CFD data randomly sampled using the Latin hypercube method. The numerically optimized double-expansion holes at different inclination angles were experimentally evaluated and compared with the optimized single-expansion laid-back fan-shaped holes, and the optimized double-expansion hole at α=30∘ was manually modified based on experiment results. Compared with the optimal single-expansion holes, the area-averaged cooling effectiveness of the double-expansion holes was increased by 34.5% at α=30∘, by 27.8% at α=45∘, and basically the same at α=60∘, showing the benefit of the double-expansion concept. The loss mechanism of film cooling was also analyzed in the perspective of the entropy generation rate, showing the optimal double-expansion holes have 21% less loss compared to a baseline narrow single-expansion hole. It was also found that CFD sometimes predicts a different trend from the experiment in optimization, and the experimental validation is necessary.

Film cooling improvement analysis in gas turbine blades by swirling coolant flow using a numerical study and an RBF artificial neural network

BackgroundA numerical simulation is carried out to assess the film cooling improvement of the gas turbine blades using proposed hole shapes. The manipulation of the hole shapes is only conducted on the suction and pressure side. The study follows having converging round-to-slot holes and making the coolant flow to be swirled by considering the screwed holes’ wall or inserting a twisted tape, with different rotation angles or the tape's widths, in the film-cooling holes. MethodsUsing the numerical simulation results, total and individual coolant mass flow rates, bulk-averaged blowing ratio, energy loss, and averaged adiabatic coolant effectiveness are computed. A post-processing method is used for calculating the parameters. The numerical results are also fed to an artificial neural network to evaluate its ability to predict. Significant findingsConsidering an overall comprehensive evaluation, the swirling patterns can be beneficial, as they lead to 11% higher adiabatic cooling effectiveness and 4% lower energy loss, compared to the traditional cylindrical holes. Further, inserting twisted tape in the holes requires a 33% lower coolant mass flow rate compared to cylindrical holes. RBF neural network shows a powerful tool to predict mass flow rates.

Film-cooling hole optimization and experimental validation considering the lateral pressure gradient

The flow in the turbine endwall region consists of the complicated secondary flow structures driven by the lateral pressure gradient, which heavily affects the performance of film cooling. In this work, the film-cooling hole design optimization is performed considering the existence of the lateral pressure gradient in the real flow environment. Results have shown that the optimal film-cooling hole design is heavily influenced by the lateral pressure gradient in the endwall region, especially the compound angle design is clearly different from the flat plate flow environment. The optimization results are further validated with experiments using the pressure-sensitive paint (PSP) technique, and the film cooling performance is shown to be improved by 42.9%. This work demonstrates the importance of considering the real flow environment in the film-cooling hole design and also can provide guidance to the film-cooling hole design in the endwall region.

Printability and Overall Cooling Performance of Additively Manufactured Holes With Inlet and Exit Rounding

Abstract To improve cooling effectiveness of gas turbine hardware, various film cooling hole shapes have previously been researched. Unique design modifications have recently been made possible through the design freedom allotted by additive manufacturing (AM). As one example, creating a rounded inlet for a film-cooling hole can mitigate separation at the inlet. This study explores various geometric features by exploiting the uses of additive manufacturing for shaped film cooling holes at engine scale. Both printability and cooling performance were evaluated. Resulting from this study, additively manufactured holes with hole inlet and exit rounding were printed with some variations from the design intent (DI). The largest deviations from the design intent occurred from dross roughness features located on the leeward side of the hole inlet. The measured overall effectiveness indicated that an as-built inlet fillet decreased in-hole convection as well as decreased jet mixing compared to the as-built sharp inlet. Including an exit fillet, which prevented an overbuilt diffuser exit, was also found to decrease jet mixing. A particular insight gained from this study is the importance of the convective cooling within the hole to the overall cooling performance. In-hole roughness, which is a result of additive manufacturing, increased convective cooling within the holes but also increased jet mixing as the coolant exited the hole. The increased jet mixing caused low overall effectiveness downstream of injection.

Effect of protrusion shape on film cooling performance for the cylindrical hole embedded in a contoured crater

The cratered film-cooling hole is regarded as one of the potential applications with high cooling performance and low cost. This study focuses on the influence of the protrusion shape for the contoured crater embedded in the cylindrical hole. Four protrusion shapes, i.e., arc, rectangle, trapezoid, and triangle, are considered. The cooling effectiveness, flow structure, and aerodynamic loss for the cratered holes at blowing ratios of 0.5-2.5 are obtained using the numerical method with the shear stress turbulence model. The numerical results indicate that the arc and triangle protrusion models provide better lateral coolant coverage and higher area-averaged cooling effectiveness at higher blowing ratios, attributed to the ascendant anti-kidney-shaped vortex pair. The rectangle protrusion model provides the lowest area-averaged cooling effectiveness because the kidney-shaped vortex pair dominates the downstream flow field. For the aerodynamic loss, the largest total pressure loss coefficient occurs for the rectangle protrusion model and nearly equivalent values for the other three protrusion models. The contoured cratered holes with arc and triangle protrusions are generally recommended.

Conjugated investigation on the effect of rotation on the leading-edge impingement structures with and without film-cooling holes

A thermal-fluid-structure conjugate analysis is carried out to analyze the temperature and Von Mises stress distribution of impingement structures with and without film-cooling holes. The flow structures under stationary and rotating conditions are compared and the effect of the introduced Coriolis force in rotating channels is analyzed in detail. Turbulence model verification is accomplished by comparing the critical parameters from the experiments that involve mainstream flow and impingement cooling. Several cases of different Reynolds numbers were stimulated and the results indicate that the impingement cooling performance under rotating conditions is worse than that under stationary conditions, due to the centrifugal force enhancing the cross-flow effect in the impingement chamber. The film-cooling hole arrangement can reduce the temperature and achieve a more uniform temperature distribution at the leading edge of the blade. Compared to rotating cases, the stationary cases produce higher stress for all Reynolds numbers. For the pure impingement structure, the maximum stress region appears at the root of the blade while the maximum stress of the impingement structure with film-cooling holes locates at the exit of the film-cooling hole. Though the impingement structure with film-cooling holes achieves lower average stress of the leading edge, the maximum value of the stress is higher than the pure impingement structure.

Experimental Optimization of the Compound Angled Asymmetric Laidback Fan Shaped Film Cooling Hole

In this study, the effect of shape variables on the film cooling effectiveness of the compound angled asymmetric laidback fan shaped hole was experimentally investigated, and the optimum values of select design variables were presented. Among the shape variables of the compound angled asymmetric laidback fan shaped hole, the windward and leeward lateral expansion angles and the compound angle were selected as design variables. Test points were chosen using the central composite design method, and the selected design variables were optimized using the Kriging model. The film cooling effectiveness was measured using the PSP technique, and the experiment was conducted under the two density ratios of 1.5 and 2.0 and four blowing ratios of 1.0, 1.5, 2.0, and 2.5. Experimental results showed that the film cooling performance was improved for higher density ratios than lower density ratios. The main effects analysis indicated that larger windward and leeward lateral expansion angles induced higher film cooling effectiveness; however, the compound angle did not show consistent results. For the optimized hole at the density ratio 2.0, the results indicated that the overall averaged film cooling effectiveness of the optimized compound angled asymmetric laidback fan shaped hole was higher than that of the optimized fan shape holes of previous literature.