Experimental investigation on impingement/effusion cooling with short normal injection holes

An experimental investigation on cooling performances of integrally impingement/effusion cooling configurations with film cooling holes angled normal to the mainstream flow is conducted. The adiabatic film cooling effectiveness and the overall cooling effectiveness are measured on a polycarbonate test plate and a stainless steel plate respectively. Effects of the blowing ratio (ranged from 0.6 to 2.4), multi-hole arrangement (inline and staggered), hole-to-hole pitch ratio (ranged from 3 to 5) and jet-to-target spacing ratio (ranged from 2 to 4) on the cooling performance are examined. In addition, jet impingement heat transfer is measured to evaluate the dense array jet impingement behaviors with local extraction of coolant via effusion holes. A new parameter named corrected blowing ratio is introduced in the present to evaluate the cooling effectiveness for different effusion or impingement–effusion configurations under a given quantity of cooling air. In an integrally impingement–effusion cooling configuration, multiple jet impingement with local extraction of coolant via effusion holes is able to produce higher overall heat transfer under lower jet-to-target spacing and denser jet array. The action of additional jet impingement heat transfer on improving overall cooling performance is highly dependant on the blowing ratio, multi-hole arrangement and jet-to-target spacing, which seem to be behaved superior in the situations where the film cooling effect isolating the wall surface from the hot mainstream is weak. For an integrally impingement–effusion cooling configuration, the densest hole-to-hole array is favorable in the situations where the coolant mass flow rate per unit area of cooled surface is low. As the coolant mass flow rate per unit area of cooled surface increases, the hole-to-hole pitches could be gradually enlarged to make effective utilization of array jet impingement.     

Film cooling performance of converging slot-hole rows on a gas turbine blade

Experimental tests have been performed to investigate the film cooling performance of converging slot-hole (console) rows on the turbine blade. Film cooling effectiveness of each single hole row is measured under three momentum flux ratios based on the wide-band liquid crystal technique. Measurements of the cooling effectiveness with all the hole rows open are also carried out under two coolant–mainstream flux ratios. Film cooling effectiveness of cylindrical hole rows on the same blade model is measured as a comparison. The results reveal that the trace of jets from both consoles and cylindrical holes is converging on the suction surface and expanding on the pressure surface by the influence of the passage vortex, while the influence of passage vortex on the jets from consoles is weaker. The film coverage area and the film cooling effectiveness of single/multiple console row(s) are much larger than those of single/multiple cylindrical hole row(s). When the console row is discrete and the diffusion angle of the console is not very large, the adjacent jets cannot connect immediately after ejecting out of the holes and the cooling effectiveness in the region between adjacent holes is relatively lower. On the pressure surface, the film cooling effectiveness of console rows increases notably with the increasing of momentum flux ratio or coolant–mainstream flux ratio. But on the suction side, the increase in cooling effectiveness is not very notable for console row film cooling as the coolant flux increases. Moreover, for the film cooling of single console row at the gill region of the suction surface, the jets could lift off from the blade surface because of the convex geometry of the suction surface.     

Study on trench film cooling on turbine vane by large-eddy simulation

Abstract

Combined with infrared thermography experiments, large-eddy simulation was used for studying trench film cooling on C3X vane model at the mainstream Reynolds number of 2.5?×?10⁵ based on the chord length, and nominal blowing ratios of 0.5 and 1.5. The instantaneous and time-averaged characteristics for trench film cooling were analzyed in detail. Inside the trench, a pair of recirculation vortices promotes the coolant spreading on spanwise direction, mitigates the jet penetration into mainstream, and improves cooling effectiveness. On pressure surface, hairpin vortices play the dominate role in the unsteady flow fields. Downstream of the trench, a streamwise vortex pair corresponding to anti-CRVP (Counter rotating vortex pair) is generated on both sides of hairpin structures, and causes high turbulent fluctuation. On suction surface, the mainstream boundary layer transits from laminar to turbulent flow in the upstream of the coolant exit, and large numbers of small-scale vortices dominate the flow dynamics. Spectrum analysis of pressure signals shows that, on pressure surface, trench and round-hole film cooling both exhibit strong periodicity. On suction surface, randomness is more pronounced. The statistical characteristics of velocity and temperature fluctuations were also discussed in detail. Overall, significant cooling augmentation by trench hole is seen on both the suction and pressure surfaces, especially at high blowing ratio.     

Numerical approach to hole shape effect on film cooling effectiveness over flat plate including internal impingement cooling chamber

Numerical approach have been conducted on a flat, three-dimensional discrete-hole film cooling geometries that included the mainflow, injection tubes, impingement chamber, and supply plenum regions. The effects of blowing ratio and hole’s shape on the distributions of flow field and adiabatic film cooling effectiveness over a flat plate collocated with two rows of injection holes in staggered-hole arrangement were studied. The blowing ratio was varied from 0.3 to 1.5, while the density ratio of the coolant to mainstream is kept at 1.14. The geometrical shapes of the vent of the cooling holes are cylindrical round, simple angle (CYSA), forward-diffused, simple angle (FDSA) and laterally diffused, simple angle (LDSA). Diameter of different shape of cooling holes in entrance surface are 5.0 mm and the injection angle with the main stream in streamwise and spanwise are 35° and 0° respectively. Ratio of the length of the cooling holes and the diameter in the entrance surface is 3.5. The distance between the holes in the same row as well as to the next row is three times the diameter of hole in the entrance surface.The governing equation is the fully elliptic, three-dimensional Reynolds-averaged Navier–Stokes equations. The mesh used in the finite-volume numerical computation is the multi-block and body-fitted grid system. The simulated streamwise distribution of spanwise-averaged film cooling effectiveness exhibited that low Reynolds number k–ε model can give close fit to the experimental data of the previous investigators. Present study reveals that (1) the geometrical shape of the cooling holes has great effect on the adiabatic film cooling efficiency especially in the area near to the cooling holes. (2) The thermal-flow field over the surface of the film-cooled tested plate dominated by strength of the counter-rotating vortex pairs (CRVP) that generated by the interaction of individual cooling jet and the mainstream. For LDSA shape of hole, the CRVP are almost disappeared. The LDSA shape has shown a highest value in distribution of spanwise-averaged film cooling effectiveness when the blowing ratio increased to 1.5. It is due to the structure of the LDSA is capable of reducing the momentum of the cooling flow at the vent of the cooling holes, thus reduced the penetration of the main stream. (3) The structure of the LDSA can also increase the lateral spread of the cooling flow, thus improves the spanwise-averaged film cooled efficiency.     

Endwall heat transfer and film cooling measurements in a turbine cascade with injection upstream of leading edge

Detailed heat transfer measurements were conducted on the endwall surface of a large‐scale low‐speed turbine cascade with single and double row injection on the endwall upstream of leading edge. Local film cooling effectiveness and the heat transfer coefficient with coolant injection were determined at blowing ratios 1.0, 2.0, and 3.0. In conjunction with the previously measured flow field data, the behaviors of endwall film cooling and heat transfer were studied. The results show that endwall film cooling is influenced to a great extent by the secondary flow and the coverage of coolant on the endwall is mainly determined by the blowing ratio. An uncovered triangle‐shaped area with low effectiveness close to pressure side could be observed at a low blowing ratio injection. The averaged effectiveness increases significantly when injecting at medium and high blowing ratios, and uniform coverage of coolant on the endwall could be achieved. The averaged effectiveness could be doubled in the case of double row injection. It was also observed that coolant injection made the overall averaged heat transfer coefficient increase remarkably with blowing ratio. It was proven that film cooling could reduce endwall heat flux markedly. The results illustrate the need to take such facts into account in the design process as the three‐dimensional flow patterns in the vicinity of the endwall, the interactions between the secondary flow and coolant, and the augmentation of heat transfer rate in the case of endwall injection. © 2004 Wiley Periodicals, Inc. Heat Trans Asian Res, 33(3): 141–152, 2004; Published online in Wiley InterScience ( www.interscience.wiley.com ). DOI 10.1002/htj.20007     

Numerical approach to film cooling effectiveness over a plate surface with coolant impingement

The aim of the present study is conducting the numerical approach to a combination of internal jet impingement and external film cooling over a flat plate. A multi-block three-dimensional Navier-Stokes code, CFX 4.4, with k-e turbulence model is used to simulate this complicated thermal-flow structure induced by the interaction of coolant jet and hot cross mainstream. By assuming the adiabatic wall boundary condition on the tested film-cooled plate, both the local and the spanwise-averaged adiabatic film cooling effectiveness are evaluated for comparison of the cooling performance at blowing ratios of Br=0.5, 1.0, and 1.5. Film flow data were obtained from a row of five cylindrical film cooling holes, inclined in angle of 35?and 0?in direction of streamwise and spanwise, respectively. The film cooling hole spacing between adjacent holes is 15 mm for all the holes. Before the coolant flow being injected through individual cooling hole then encountered with the mainstream, an impingement chamber containing an impingement plate with 43 holes is located on the path of coolant flow. Present study also focused on the effect of impingement spacing, 10mm, 20mm, and 30mm. Compare the results, we find the impingement jet has a significant effect on the adiabatic film cooling effectiveness. As the coolant impingement spacing is fixed, results indicated that higher blowing ratio would enhance the local and the spanwise-averaged adiabatic film cooling effectiveness. Moreover, neither uniform nor parabolic distribution of pressure distribution are observed within the coolant hole-pipe.     

Investigation of the Influence of Inclination Angle and Diffusion Angle on the Film Cooling Performance of Chevron Shaped Hole

The film cooling performance of chevron holes with different inclination angles and exit lateral diffusion angles has been studied experimentally and numerically. The inclination angles include 35° and 55°. The exit lateral diffusion angles include 20° and 25°. The film cooling effectiveness, heat transfer coefficient and discharge coefficient were measured on a flat plate model by transient liquid crystal measurement technique under four blowing ratios. The results show that the large inclination angle reduces the film cooling effectiveness. The influence of diffusion angle has two aspects: the large diffusion angle leads to mainstream ingestion and decreases film cooling effectiveness at M=1.0 and 1.5; however, the large diffusion angle increases the film cooling effectiveness at high blowing ratio of 2.0, because the larger hole exit area decreases the normal momentum component of the film jet. The large inclination angle decreases the heat transfer coefficient in the right downstream region at M=0.5 and 1.0. The large diffusion angle enhances the heat transfer in the right downstream of the holes in M=0.5～1.5 conditions. The chevron hole with large inclination angle generally has the highest discharge coefficient.     

Numerical investigation of the flow field and heat transfer characteristics for upstream continuous and truncated ribs

To verify the applicability of upstream ribs in film cooling, the present numerical study examines heat transfer characteristics and flow field for ribs located upstream of the film hole. Five ribs including bilaterally truncated ribs, centrally truncated ribs, and continuous ribs are explored with the smooth case at two blowing ratios and fixed crossflow Reynolds number. The results show that the film cooling effectiveness of cases with ribs outperforms the case without rib at a low blowing ratio. Centrally truncated ribs and continuous ribs provide superior cooling effectiveness than bilaterally truncated ribs and smooth cases. The introduction of ribs makes the distribution of the heat transfer coefficient (HTC) uneven after the hole. Among these, centrally truncated ribs increased the HTC, while bilaterally truncated ribs reduce the HTC in the far hole area at a high blowing ratio. It is found that anti-kidney-shaped vortex pairs are generated between two adjacent jets for centrally truncated rib cases, while they are generated in front of the jets for bilaterally truncated rib cases. For continuous rib, the impingement of the mainstream gas on the jet leads to a reduction in strength of the kidney-shaped vortex, which allows the coolant to form a better coverage.     

孔型对水雾/空气气膜冷却特性影响的数值模拟研究

在平板表面分别开设了圆柱孔、展向扩张孔和收缩扩张孔。对比研究了3种孔型的纯空气气膜冷却和水雾/空气气膜冷却特性。在3种吹风比：0.66、1.04、1.44下展开研究。将圆柱孔的数值计算结果与文献中的实验结果进行对照以验证水雾/空气冷却数值计算方法的正确性。对3种孔型下冷却气体混合物的无量纲速度矢量图和部分水雾颗粒的运动轨迹进行了比较和分析。对3种孔型中心线和展向平均气膜冷却效率进行了比较和分析。结果表明：圆柱孔和展向扩张孔射流形成的肾形涡将水雾颗粒抬离平板表面。收缩扩张孔射流形成的肾形涡增强了水雾颗粒的展向扩散并将靠近孔口两侧区域的水雾颗粒逐渐抬离平板表面。对于圆柱孔和展向扩张孔,其射流形成的肾形涡削弱了水雾颗粒对于展向平均气膜冷却效率的提高作用,收缩扩张孔水雾/空气冷却的展向平均气膜冷却效率在3种吹风比下均大于0.6,当吹风比为1.44时,收缩扩张孔的展向平均气膜冷却效率约为展向扩张孔的2倍,圆柱孔的4倍。2种冷却方式下,在吹风比从1.04增大到1.44时,展向扩张孔中心线气膜冷却效率降低0.3左右,而收缩扩张孔中心线冷却效率的降幅小于0.1。     

不同叶顶结构对燃气透平动叶顶部气膜冷却性能的影响

对不同叶顶结构的GE-E3叶片的气膜冷却现象进行了数值研究,比较了三种不同的叶顶结构：平顶、凹槽顶和平顶开槽孔结构在叶顶部的流动和冷却现象,并分析了吹风比对这三种结构的冷却性能的影响。发现凹槽顶和平顶开槽孔在结构上具有相似性;在叶顶开槽后,既降低了射流动量,又降低了顶端泄漏流速,有助于提高冷却效果,同时由于凹槽顶的槽比开槽孔的槽大,冷却气体和燃气在槽内充分混合,使得凹槽顶结构具有最高的冷却效率值和最低的换热系数值,平顶开槽孔结构次之。     

Experimental investigations on overall cooling effect of ribbed channel with air bleeds

An experimental investigation on overall heat transfer performance of a rectangular channel, in which one wall has periodically placed oblique ribs to enhance heat exchange and cylindrical film holes to bleed cooling air, has been carried out in a hot wind tunnel at different mainstream temperatures, hot mainstream Reynolds numbers, coolant Reynolds numbers and blowing ratios. To describe the cooling effect of combined external coolant film with the internal heat convection enhanced by the ribs, the overall cooling effectiveness at the surface exposed in the mainstream with high temperature was calculated by the surface temperatures measured with an infrared thermal imaging system. The total mass flow rate of cooling air through the coolant channel was regulated by a digital mass flow rate controller, and the blowing ratio passing through the total film holes was calculated based on the measurements of another digital-type mass flow meter. The detailed distributions of overall cooling effectiveness show distinctive peaks in heat transfer levels near the film holes, remarkable inner convective heat transfer effect over entire channel surface, and visible conductive heat transfer effect through the channel wall; but only when the coolant Reynolds number is large enough, the oblique rib effect can be detected from the overall cooling effectiveness; and the oblique bleeding hole effect shows the more obvious trend with increasing blowing ratios. Based on the experimental data, the overall cooling effectiveness is correlated as the functions of Re_m (Reynolds number of hot mainstream) and Re_c (Reynolds number of internal coolant flow at entrance) for the parametric conditions examined.     

Turbine endwall film cooling with combustor-turbine interface gap leakage flow: Effect of incidence angle

This paper is focused on the film cooling performance of combustor-turbine leakage flow at off-design condition. The influence of incidence angle on film cooling effectiveness on first-stage vane endwall with combustor-turbine interface slot is studied. A baseline slot configuration is tested in a low speed four-blade cascade comprising a large-scale model of the GE-E³Nozzle Guide Vane (NGV). The slot has a forward expansion angle of 30 deg. 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). The blowing ratio of the coolant through the interface gap varies from M = 0.1 to M = 0.3, while the blowing ratio varies from M = 0.7 to M = 1.3 for the endwall film cooling holes. The film-cooling effectiveness distributions are obtained using the pressure sensitive paint (PSP) technique. The results show that with an increasing blowing ratio the film-cooling effectiveness increases on the endwall. As the incidence angle varies from i = +10 deg. to i = ?10 deg., at low blowing ratio, the averaged film-cooling effectiveness changes slightly near the leading edge suction side area. The case of i = +10 deg. has better film-cooling performance at the downstream part of this region where the axial chord is between 0.15 and 0.25. However, the disadvantage of positive incidence appears when the blowing ratio increases, especially at the upstream part of near suction side region where the axial chord is between 0 and 0.15. On the main passage endwall surface, as the incidence angle changes from i = +10 deg. to i = ?10 deg., the averaged film-cooling effectiveness changes slightly and the negative incidence appears to be more effective for the downstream part film cooling of the endwall surface where the axial chord is between 0.6 and 0.8.     

Investigation on cooling effectiveness and aerodynamic loss of a turbine cascade with film cooling

This paper describes the numerical study on film cooling effectiveness and aerodynamic loss due to coolant and main stream mixing for a turbine guide vane. The effects of blowing ratio, mainstream Mach number, surface curvature on the cooling effectiveness and mixing loss were studied and discussed. The numerical results show that the distributions of film cooling effectiveness on the suction surface and pressure surface at the same blowing ratio (BR) are different due to local surface curvature and pressure gradient. The aerodynamic loss features for film holes on the pressure surface are also different from film holes on the suction surface.     

Film cooling performance of a row of dual-fanned holes at various injection angles

Film cooling performance about a row of dual-fanned holes with injection angles of 30°, 60 ° and 90° were experimentally investigated at blowing ratios of 1.0 and 2.0. Dual-fanned hole is a novel shaped hole which has both inlet expansion and outlet expansion. A transient thermochromic liquid crystal technique was used to reveal the local values of film cooling effectiveness and heat transfer coefficient. The results show that injection angles have strong influence on the two dimensional distributions of film cooling effectiveness and heat transfer coefficient. For the small injection angle of 30 degree and small blowing ratio of 1.0, there is only a narrow spanwise region covered with film. The increase of injection angle and blowing ratio both leads to the enhanced spanwise film diffusion, but reduced local cooling ability far away from the hole. Injection angles have comprehensive influence on the averaged film cooling effectiveness for various x/d locations. As injection angles are 30 and 60 degree, two bands of high heat transfer coefficients are found in mixing region of the gas and coolant. As injection angle increases to 90 degree, the mixing leads to the enhanced heat transfer region near the film hole. The averaged heat transfer coefficient increases with the increase of injection angle.     

Experimental investigation on the leading edge film cooling of cylindrical and laid-back holes with different hole pitches

Experimental investigation has been performed to study the film cooling performances of cylindrical holes and laid-back holes on the turbine blade leading edge. Four test models are measured for four blowing ratios to investigate the influences of film hole shape and hole pitch on the film cooling performances Film cooling effectiveness and heat transfer coefficient have been obtained using a transient heat transfer measurement technique with double thermochromic liquid crystals. As the blowing ratio increases, the trajectory of jets deviates to the spanwise direction and lifts off gradually. However, more area can benefit from the film protection under large blowing ratio, while the is also higher. The basic distribution features of heat transfer coefficients are similar for all the four models. Heat transfer coefficient in the region where the jet core flows through is relatively lower, while in the jet edge region is relatively higher. For the models with small hole pitch, the laid-back holes only give better film coverage performance than the cylindrical holes under large blowing ratio. For the models with large hole pitch, the advantage of laid-back holes in film cooling effectiveness is more obvious in the upstream region relative to the cylindrical holes. For the cylindrical hole model and the laid-back hole model with the same hole pitch, heat transfer coefficients are nearly the same with each other under the same blowing ratios. Compared with the models with large hole pitch, the laterally averaged film cooling effectiveness and heat transfer coefficient are larger for the models with small hole pitch because of larger proportion of film covering area and strong heat transfer region.     

Numerical Study on Flow and Cooling Characteristics for Supersonic Film Cooling

With the booming performances of the gas turbine engine, the turbine vane of the gas turbine engine experiences more extreme thermal environment with supersonic flows. The film cooling applied in the supersonic flow condition has essential difference from that used in the subsonic flow condition in the flow characteristics and cooling effectiveness. This article focused on the film cooling of two parallel flows (primary flow and coolant flow) with supersonic or subsonic velocity, respectively. The results show that: on the condition of supersonic primary flow and subsonic coolant flow, the coolant flow with lower momentum is sheared and dragged by the higher momentum primary flow because of the viscous property of fluid． At the meantime, the thermal and momentum of the primary flow transfers into the coolant flow rapidly． It causes the great damage of the film coverage, and the decrease of the cooling effectiveness dramatically． In contrast, on the condition of supersonic primary flow and supersonic coolant flow, the film coverage of the supersonic coolant flow can last further far than that of the subsonic coolant flow on the same blowing ratio． The turbulence kinetic energy seems to be depressed by the shorten of velocity difference of two supersonic flow. Therefore, the cooling effectiveness is enhanced by 45% for the supersonic primary and coolant flow.     

Influence on film cooling effectiveness of novel holes based on cylindrical configurations

In this study, four novel film cooling hole designs, all based on cylindrical holes, are numerically evaluated, and compared with those of a simple cylindrical hole and a laterally-diffused shaped hole. Film cooling effectiveness and surrounding thermal and flow fields are documented for operation with various blowing ratios. It is shown that the two-stage cylindrical hole can improve film cooling effectiveness at higher blowing ratios. The primary hole with two secondary holes can enhance film cooling performance by creating anti-kidney vortex pairs that will weaken jet liftoff caused by the kidney vortex pair that is created by the primary hole. The tri-circular shaped hole provides better film cooling effectiveness values only near the hole, but worse at downstream positions. The two-stage structure for the tri-circular shaped hole provides better film coverage because it changes the flow structure inside the delivery channel and decreases jet penetration into the passage flow.     

Effectiveness measurements for a cooling film disrupted by a row of jets

The influence of cold air jet injection upon filmcooling of combustion chamber walls has been investigated by simulation. The experiments, in which a row of cold air jets is injected perpendicularly to the film cooled test chamber wall, show considerable decreases of the film cooling effectiveness downstream of the jet exit. The same behavior occurs in real combustion chambers of aircraft gas turbine engines. The tests are selected to represent combinations of geometries and flow variables (except the temperature ratios) which are pertinent to gas turbine combustor design. The present investigations cover effectiveness measurements for various combinations of the parameters: a) velocity ratio coolant film to mainstream b) velocity ratio jet to mainstream c) dimensionless distance between coolant film and jet exit d) spacing ratio e) ratio jet diameter to slot height f) ratio lip thickness to slot height.     

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.     

Experimental Investigation of Louver Cooling Scheme on Gas Turbine Stator

An experimental investigation has been conducted to investigate the film cooling performance of the louver scheme over the surface of a gas turbine stator using a transient thermochromic liquid crystal technique. A two-dimensional airfoil cascade has been employed during the study. The exit Reynolds number based on the true chord length is 1.5E5 and the exit Mach number is 0.23. Two rows of an axially oriented louver scheme are distributed on both suction and pressure sides in a staggered arrangement. The effect of hole location on the cooling performance is investigated for each row individually; then row interaction is investigated at four different blowing ratios ranging from 1 to 2 and a 0.9 density ratio. The detailed local performance distribution and the lateral-averaged normalized performance are presented over both sides of the vane in terms of heat transfer coefficient and cooling effectiveness. The louver scheme provides a better cooling performance compared with the similar cylindrical scheme of the same base diameter at the same cooling amount. The blowing ratio does not influence significantly the performance for the louver scheme due to the considerable decrease in the jet momentum that impedes the jet lift-off at exit. The location of the scheme exit has a high impact on the cooling performance as it affects the development of the boundary layer. The double injection on the pressure side provides a superior effectiveness due to the blockage of the mainstream by the coolant injected from the first row. The louver scheme provides higher net heat flux reduction, which suits the cooling capacity needed for the next generation of gas turbines.