Effects of Trenched Film Hole Configurations on the Endwall Film Cooling and Suction Side Phantom Cooling

To maximize the turbine thermal efficiency, modern gas turbine's inlet temperature is significantly augmented within the past few decades. To prolong the lifespan of gas turbines, many efficient cooling techniques have been proposed and applied in the endwall cooling schemes. However, conventional discrete film hole does not take effect at the leading edge nearby region. In this research, how the trenched film hole configurations affects the endwall cooling and phantom cooling characteristics were deeply studied by using a verified approach. Steady 3D Reynolds-averaged Navier-Stokes(RANS) governing equations together with the shear stress transport(SST) k-w turbulence model have been solved. Firstly, results indicate that trenched film holes greatly influence the cooling effectiveness at leading edge nearby region compared to normal case. Nevertheless, suction side phantom cooling is hardly influenced by the trenched film holes. Secondly, the case with a smaller trench width obtains higher endwall cooling effectiveness, particularly at upstream region. More importantly, the cases with W=3D achieve large cooling effectiveness at leading edge nearby region with little influence by trench depth. Additionally, majority of trenched film holes coolant flow is driven towards middle passage. Therefore, the suction side phantom cooling is unaffected by the trenched film holes.     

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.     

进口气流冲角对环形叶栅端壁气膜冷却效率的影响

进口气流冲角变化会改变端壁横向压力梯度致使端壁气膜冷却气体覆盖发生变化。本文在环形叶栅端壁30％、60％、90％轴向弦长处和距前缘-10％轴向弦长端壁处布置单排圆柱形气膜冷却孔,运用CFD方法在吹风比为1．0条件下对来流冲角为10°、0°和-15°时端壁气膜冷却效率进行对比分析。结果显示冲角由正值到负值,压力面和吸力面一侧两支马蹄涡夹裹冷却气膜主体向压力面偏移,致使吸力面一侧端壁未冷却区域扩大,压力面一侧端壁未冷却区域缩小。吸力面一侧端壁气膜冷却效率沿流向逐渐提升,并在流道中、下游完全冷却覆盖端壁,而压力面一侧端壁一直未能获得冷却气膜覆盖。     

Film cooling on a gas turbine blade pressure side or suction side with axial shaped holes

The film cooling effectiveness on the surface of a high pressure turbine blade is measured using the pressure sensitive paint (PSP) technique. Four rows of axial laid-back, fan-shaped cooling holes are distributed on the pressure side while two such rows are provided on the suction side. The coolant is only injected to either the pressure side or suction side of the blade at five average blowing ratios ranging from 0.4 to 1.5. The presence of wakes due to upstream vanes is simulated by placing a periodic set of rods upstream of the test blade. Effect of the upstream wakes is recorded at four different phase locations with equal intervals along the pitch-wise direction. The freestream Mach numbers at cascade inlet and exit are 0.27 and 0.44, respectively. Results reveal that the tip leakage vortices and endwall vortices sweep the coolant film on the suction side to the midspan region. The film cooling effectiveness on the suction side is usually higher than that on the pressure side except the regions affected by the secondary vortices. The presence of upstream wakes results in lower film cooling effectiveness on the blade surface. The moderate blowing ratios (M = 0.6 or M = 0.9) give higher film cooling effectiveness immediately downstream of the film cooling holes. Further downstream of the holes, higher blowing ratios cover wider surface area.     

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.     

Leading edge film cooling enhancement for an inlet guide vane with fan-shaped holes

This paper describes the improvement of leading edge film cooling effectiveness for a turbine inlet guide vane by using fan-shaped film cooling holes. The modification details are presented in comparison with the base-line configuration of cylindrical holes. Numerical simulations were carried out for the base-line and modified configurations by using CFX, in which the κ-ε turbulence model and scalable wall function were chosen. Contours of adiabatic film cooling effectiveness on the blade surfaces and span-wise distributions of film cooling effectiveness downstream the rows of cooling holes interested for the different cooling configurations were compared and discussed. It is showed that with the use of fan-shaped cooling holes around the leading edge, the adiabatic film cooling effectiveness can be enhanced considerably. In comparison with the cylindrical film cooling holes, up to 40% coolant mass flow can be saved by using fan-shaped cooling holes to obtain the comparable film cooling effectiveness for the studied inlet guide vane.     

Numerical evaluation on single-row trenched-hole film cooling performances on turbine guide vane under engine-representative conditions

A numerical investigation is performed to study the effects of transverse trenches on the film cooling performances of single row film cooling holes on the turbine guide vane, under the engine-representative conditions. Two trenches are investigated in the current study, including a conventional straight trench and a novel serrated trench, both having the same slot width of 2.5d and depth of 0.75d. With the presence of trenches, the kidney vortices are effectively destroyed at the suction side, replaced by a pair of secondary vortices originating from the edge of trench slot, with a sense of rotation opposite to the kidney-vortex pair. However, at the pressure side, the kidney vortices remain clearly in the trenched-hole cases. The roles of trenched-hole on the film cooling enhancement behave more pronounced on the pressure surface when compared to the respective one on the suction surface. On the pressure surface, the serrated trench displays a more obvious role in the film cooling enhancement when compared to the straight trench. Combined with the effects of trenched-hole on the enthalpy loss coefficient, it is suggested that the trenches are more reasonably used on the pressure surface.     

Experimental Investigation on the Effect of Mainstream Turbulence on Full Coverage Film Cooling Effectiveness for a Turbine Guide Vane

Experiments have been performed to investigate the effect of mainstream turbulence on the three-dimensional distribution of the full coverage film cooling effectiveness for two enlarged actual twisted vanes with cylindrical or shaped holes. The film cooling effectiveness was measured by transient liquid crystal technique at mainstream turbulence intensities of 2%, 9% and 15%. The mass flow rate ratios range from 5.5% to 12.5%. There are 3, 8 and 7 rows of film holes on the suction side, leading edge and pressure side, respectively. Results show that for the cylindrical hole vane the high mainstream turbulence intensity decreases the film cooling effectiveness in the top region and down region of pressure side in the low mass flow rate ratio of 5.5%, while the effect is opposite in the high mass flow rate ratio of 12.5%. The film cooling effectiveness in the middle region of pressure side decreases obviously with the increase of the turbulence at the low mass flow rate ratio of 5.5%, while the influence of increasing turbulence weakens gradually with the increase of mass flow rate ratio. Moreover, the high mainstream turbulence improves the film cooling effectiveness in the further downstream of the holes on suction side at the high mass flow rate ratio of 12.5%. For the shaped hole vane, the increase of mainstream turbulence decreases the film cooling effectiveness at all mass flow rate ratios. This study reveals the influence rule of the mainstream turbulence on the film cooling effectiveness in the different regions of the three-dimensional vane surface. The results would guide the designs of engineering heat transfer with application in gas turbine blade/vane cooling.     

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.     

冲角和孔排间距对端壁气膜冷却效率的影响

燃气轮机在变工况运转时透平叶栅和级的特性对燃机总体性能影响极大,而叶栅端壁气膜冷却效率是关键因素。为了提高端壁气膜冷却效率,通过优化气膜孔间距排列的方法,在叶栅端壁20%、50%、90%轴向弦长处和距前缘-10%轴向弦长端壁处布置单排带复合角度的圆柱形气膜冷却孔,运用CFD(计算流体动力学)方法对冲角(10°、0°、-10°)在不同吹风比(1、1.5、2)条件下端壁气膜冷却效率进行对比分析。结果表明:采用气膜孔非等距排列方式能有效缓解因横向压力梯度变化引起的马蹄涡在压力侧的阻隔作用,压力侧冷却效率较高;高吹风比的冷却射流会出现抛射冷却,能有效抑制冷却射流脱离壁面,壁面平均冷却效率提高;主流正冲角有利于提高端壁吸力侧气膜冷却效率,压力侧变化不大。     

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.     

Research on cooling effectiveness in stepped slot film cooling vane

As one of the most important developments in air cooling technology for hot parts of the aero-engine,film cooling technology has been widely used.Film cooling hole structure exists mainly in areas that have high temperature,uneven cooling effectiveness issues when in actual use.The first stage turbine vanes of the aero-engine consume the largest portion of cooling air,thereby the research on reducing the amount of cooling air has the greatest potential.A new stepped slot film cooling vane with a high cooling effectiveness and a high cooling uniformity was researched initially.Through numerical methods,the affecting factors of the cooling effectiveness of a vane with the stepped slot film cooling structure were researched.This paper focuses on the cooling effectiveness and the pressure loss in different blowing ratio conditions,then the most reasonable and scientific structure parameter can be obtained by analyzing the results.The results show that 1.0 mm is the optimum slot width and 10.0 is the most reasonable blowing ratio.Under this condition,the vane achieved the best cooling result and the highest cooling effectiveness,and also retained a low pressure loss.     

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     

冷却流量对涂层叶片冷却性能影响的数值研究

采用气热耦合数值方法研究冷却流量对热障涂层气冷涡轮叶片冷却性能的影响。分析表明：叶片的冷却效率随冷却流量的增加而增大,但增幅则逐渐下降。吸力面上,附加热障涂层的效果更好。基准工况下,温度的降幅为72.6K,冷却效率的增幅为6.5%。尾部部分区域内部冷却不足,热障涂层阻碍热量从金属表面向流体传递,金属表面温度升高,综合冷却性能下降,因此,只有配合高效的内冷技术,才能达到理想的冷却效果。     

Computational prediction into staggered film cooling holes on convex surface of turbine blade

The purpose of this paper is to predict the film cooling performance of inline configuration of cooling holes in comparison to the staggered arrangement on convex surface. Three‐dimensional computational study for 10° diffused hole (β = 10°, γ = 0°) and compound hole of 10° diffused and 45° with the downstream direction (β = 10°, γ = 45°) film cooling holes were investigated for adiabatic film cooling effectiveness and have been compared with that of simple hole (β = 0°, γ = 0°) film cooling on convex surface. Both the diffused and compound holes showed better film cooling effectiveness than the simple hole in all models. In one row film cooling investigation, the centerline adiabatic film cooling effectiveness of diffused hole is slightly higher than that of the compound hole near the hole trailing edge. In the staggered case, the centerline effectiveness of the compound hole was higher for both two staggered rows and three staggered rows. For the lateral effectiveness investigations of one row, diffused hole showed higher effectiveness compared with the simple hole and the right side of the compound hole while the left side of the compound hole dominated the lateral investigations in all the models. Staggered distribution of diffused and compound holes showed higher protection for the convex surface. The present results are important dissemination in many practical applications of aero engine industry.     

动叶端壁离散孔气膜冷却的数值研究

基于某F级燃气轮机第一级动叶栅的数值模拟,以实现动叶端壁气膜冷却全覆盖为目标,分析定常下动叶端壁的流动与传热特征,拟综合考虑端壁二次流结构特征与级间封严冷气泄漏流的影响,将端壁划分为四个具有不同流动传热特征的区域,并据此设计了叶根端壁仅13孔数的离散气膜孔布置方案各区域采取不同的冷却方式根据不同吹风比下的研究结果发现:吹风比为0.75时端壁冷却有效度均值在0.2以上,实现了全端壁冷却的目标;前缘附近端壁冷却效果受吹风比影响显著,吹风比在0.75以上时冷却有效度达到0.5以上;除近压力面区域,气膜冷却效果随吹风比的增大而提高。     

Measurement of Film Effectiveness for Cylindrical and Row Trenched Cooling Holes at Different Blowing Ratios

This study was performed to investigate the effects of cylindrical and row trenched cooling holes with alignment angle of 0° and 90° at different blowing ratios on the film-cooling performance adjacent the endwall surface of a combustor simulator. The film-cooling blowing ratios varied from 1.25 to 3.18. In this study, a three-dimensional representation of a Pratt and Whitney gas turbine engine was simulated and analyzed with a commercial finite volume package FLUENT 6.2.26. The analysis has been carried out with Reynolds-averaged Navier–Stokes turbulence model on internal cooling passages. This combustor simulator was combined with the interaction of two rows of dilution jets, which were staggered in the streamwise direction and aligned in the spanwise direction. Film cooling was placed along the combustor liner walls. In comparison with the baseline case of cooling holes, the application of row trenched hole near the endwall surface doubled the performance of film-cooling effectiveness.     

Effect of hole configurations on film cooling performance

This study aims to investigate the cooling performance of various film cooling holes, including combined hole, cylinder hole, conical hole, and fan-shaped hole. For film cooling technology, a novel combined hole configuration is first proposed to improve the cooling protection for gas turbine engines. This combined hole consists of a central cylinder hole (an inclination angle of 35°) and two additional side holes (a lateral diffusion angle of 30°). Film holes for four-hole configurations have the same inlet diameter of 8?mm. The adiabatic film cooling effectiveness for each hole configuration is analyzed for varying blowing ratios (M?=?0.25, 0.5, 0.75, and 1.0). Results show that the best cooling performance for the conical and fan-shaped holes is obtained at the blowing ratio of 0.75. In addition, the combined hole configuration provides a more uniform cooling protection and a better cooling performance than the other hole configurations.     

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.