Flow visualization and film cooling effectiveness measurements around shaped holes with compound angle orientations

Experimental results are presented which describe film cooling performance around shaped holes with compound angle orientations. The shaped hole has a 15° forward expansion with an inclination angle of 35°, but the orientation angles vary from 0° to 30° and 60°. The blowing ratios considered are 0.5, 1.0 and 2.0. Flow visualizations are performed using an aerosol seeding method for single enlarged shaped hole to investigate the interaction between the mainstream and the injectant at the hole exit plane. The adiabatic film cooling effectiveness distributions are measured for a single row of seven shaped holes using the thermochromic liquid crystal technique. Flow visualization reveals the occurrence of hot crossflow ingestion into the film hole at the hole exit plane at a large orientation angle such as 60°. Shaped holes with simple angle injection do not provide substantial improvement in the film cooling performance compared to round holes. However, shaped holes with compound angle injection exhibit improved film cooling effectiveness up to 55% in comparison with round hole data at high blowing ratios.     

Effects of pressure gradient on film cooling effectiveness from two rows of simple and compound angle holes in combination

This paper describes the results of an experimental investigation into the effect of pressure gradient on the film cooling effectiveness from compound angle holes at both injection rows or the combination of one row of simple angle holes and one row of compound angle holes. Two pressure gradients were used in the range from −1.11 × 10⁻⁶ to +1.11 × 10⁻⁶.The presence of a favorable pressure gradient tends to increase the dilution of the injected coolant jets, which results in a reduction of the film cooling protection over the surface. The presence of an adverse pressure gradient at high blowing rate tends to dilute the film coolant even at a higher rate than that when zero or favorable pressure gradients are present.     

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.     

异型孔阵对气膜冷却效果的数值研究

对不同孔型在不同的吹风比下的冷却效率进行数值模拟,计算结果表明:收缩进气可以强化孔内的对流传热;扩张出气可以使冷气出口的速度降低,气膜覆盖面更广,提高气膜的覆盖效果。在小吹风比下,孔内的对流传热在总的冷却效果中占主导地位,此时缩扩孔的冷却效果最好,而随着吹风比的增大,气膜覆盖所起的作用在增大,当吹风比增大到一定程度时,扩张孔的冷却效果将具有更好的表现。     

Effect of rotation on leading edge region film cooling of a gas turbine blade with three rows of film cooling holes

Effect of rotation on detailed film cooling effectiveness distributions in the leading edge region of a gas turbine blade with three showerhead rows of radial-angle holes were measured using the Pressure Sensitive Paint (PSP) technique. Tests were conducted on the first-stage rotor blade of a three-stage axial turbine at three rotational speeds. The effect of the blowing ratio was also studied. The Reynolds number based on the axial chord length and the exit velocity was 200,000 and the total to exit pressure ratio was 1.12 for the first-stage rotor blade. The corresponding rotor blade inlet and exit Mach number was 0.1 and 0.3, respectively. The film cooling effectiveness distributions were presented along with the discussions on the influences of rotational speed, blowing ratio, and vortices around the leading edge region. Results showed that different rotation speeds significantly change the film cooling traces with the average film cooling effectiveness in the leading edge region increasing with blowing ratio.     

Influence of secondary hole injection angle on enhancement of film cooling effectiveness with horn-shaped or cylindrical primary holes

Abstract

Film cooling with primary and secondary hole injection is numerically investigated. Effects of primary hole shape and secondary hole injection angle are documented. Each primary hole, either cylindrical or laterally diffused, has two secondary, cylindrical holes located symmetrically about it. Adding secondary holes improves cooling performance. Five cases of different secondary hole injection configuration are analyzed. With a cylindrical primary hole, increasing secondary hole inclination angle provides better cooling; outwardly inclining the secondary holes shows continued improvement. With horn-shaped primary holes, smaller secondary hole inclination angles provide higher cooling at lower blowing ratios; larger secondary hole inclination angles provide higher cooling at higher blowing ratios, and compound-angle secondary hole injection shows no improvement over parallel hole injection.     

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.     

Influence of shaped injection holes on turbine blade leading edge film cooling

To improve the film cooling performance by shaped injection holes for the turbine blade leading edge region, we have investigated the flow characteristics of the turbine blade leading edge film cooling using five different cylindrical body models with various injection holes, which are a baseline cylindrical hole, two laidback (spanwise-diffused) holes, and two tear-drop shaped (spanwise- and streamwise-diffused) holes, respectively. Mainstream Reynolds number based on the cylinder diameter was 7.1 × 10⁴ and the mainstream turbulence intensities were about 0.2%. The effect of injectant flow rates was studied for various blowing ratios of 0.7, 1.0, 1.3 and 1.7, respectively. The density ratio in the present study is nominally equal to one. Detailed temperature distributions of the cylindrical body surfaces are visualized by means of an infrared thermography (IRT). Results show that the conventional cylindrical holes have poor film cooling performance compared to the shaped holes. Particularly, it can be concluded that the laidback hole (Shape D) provides better film cooling performance than the other holes and the broader region of high effectiveness is formed with fairly uniform distribution.     

Improvement of film cooling performance by trenched holes on turbine leading-edge models

The present study aims to improve cooling performance over the leading edge surface with the high temperature and high thermal stress by the introduction of trenched holes. Three staggered rows of leading-edge film cooling holes with different trench arrangements and hole orientations are included under blowing ratios of 0.5, 1.0, 1.5, and 2.0, compared with round-hole cases. Under the conditions of leading-edge flow patterns and convex curvature, the trenched hole with 2D width plays a role of “protection” of coolant against the impinging hot gas at a large range of blowing ratios. This contributes to the better lateral spread of coolant and cooling performance. Besides, the trenched holes narrow the regions with a high heat transfer coefficient and reduce the detrimental heating on the surface. Compared with round holes, the trenched holes guarantee the downstream coolant coverage and higher cooling performance at a larger inclined angle, in spite of the changed compound angles.     

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.     

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.     

发散孔纵向波纹隔热屏气膜冷却特性研究

对燃烧室内开有发散孔的纵向波纹隔热屏进行了数值模拟。研究隔热屏的四种结构参数开孔率、波纹板高度、气膜孔直径和冷却通道高度的改变对隔热屏冷却效果的影响。研究表明：在气膜孔出流总量相同的情况下,3％开孔率比6％开孔率的隔热屏平均冷却效率较高;汉纹板高度对隔热屏冷却效果影响较大,波纹板无量纲高度B=1％时的隔热屏平均冷却效率最高;冷却通道高度和气膜孔直径对隔热屏冷却效果影响较小,冷却通道高度只影响隔热屏前段的冷却效率,发散孔气膜孔直径的大小则对隔热屏冷却效率几乎没有影响。     

矩形肋片导流角度对内冷通道流动与换热特性的影响

利用数值模拟方法分析了矩形仿螺旋肋片内冷通道中肋片导流角度对内冷通道三维流场特性、换热特性以及流动阻力特性的影响。数值计算结果表明,肋片导流角度对内冷通道的流动与换热特性具有较大的影响。流场中冷却介质螺旋流动的强度随着肋片导流角增大而增强,肋片导流角度越大则内冷通道的换热强度越强,同时通道中流动阻力也明显增大。从内冷通道的综合换热效果来看,当肋片导流角度为7。时,矩形仿螺旋肋片内冷通道的综合换热效果最好。     

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.     

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.     

Investigation on film cooling with swirling coolant flow by optimizing the inflow chamber

Three different kinds of coolant chamber configuration for film cooling are proposed to develop the swirling coolant flow at blowing ratios ranging from 0.5 to 2.0. The results show that the difference of film cooling effectiveness for three kinds of coolant chamber configuration is little at low blowing ratio, but the advantage of swirling film cooling becomes obviously with the increase of blowing ratio. When the blowing ratio is 2.0, the jet momentum of original coolant chamber configuration is large and uniform, which leads to the lowest cooling effectiveness due to the formation of a strong kidney vortex. The first coolant chamber configuration has a low jet momentum region at upstream of the film hole, the coolant in this region interacts with high temperature mainstream and bypasses the large jet momentum coolant to attach cooling surface at downstream. The second coolant chamber configuration is sprayed with the structure of unidirectional vortex, which forms a vortex pressing on other vortex, making the coolant in pressed vortex attach surface better, producing the best coverage and the higher film cooling effectiveness.     

Effect of shock waves on supersonic film cooling with a slotted wall

A slotted wall with a cavity which reduces the effect of the shock wave on the film cooling was developed through understanding of the mechanism by which the shock wave affects the supersonic film cooling. Numerical results show that the supersonic film cooling effectiveness with the slotted wall is improved after the shock wave incidence, even better than that without the shock wave effect. The cooling stream flows into the cavity upstream of the slotted wall and flows out downstream, which bypasses more cooling gas to protect the surface downstream after the shock wave incidence, which weakens the effect of the shock wave on the film cooling. Upstream of the shock wave incidence, the slotted wall reduces the mixing between the mainstream and the cooling stream and the coolant boundary layer thickness, which reduces the film cooling effectiveness for both structures than without the slotted wall, with an effectiveness which nearly the same as or even a little better than without the slotted wall for another structure.     

Two-phase flow simulation of mist film cooling with deposition for various boundary conditions

Air film cooling is a conventional cooling technique that has been successfully used for gas turbine hot-section components, such as combustor liners, combustor transition pieces, and turbine vanes and blades. However, the increased benefit seems to approach a limit. This paper investigates the film cooling effectiveness considering mist injection. All the studies for various boundary conditions are conducted numerically, including the effects of droplet size, the flow rates of droplet injection, and the coolant air. Film cooling is also affected by the interaction between deposition and mist injection. A deposition configuration is located near the film hole with an inclination angle of 35°. Results show that the combined effect of injection and deposition is to weaken the film cooling effectiveness, especially upstream of x/d?=?19. For the coolant air at a low speed, the mist injection cannot provide better cooling protection than without the mist injection.     

An experimental study of film cooling in a negatively bowed cascade with a large turning angle

Transverse measurements are carried out at the outlet of the straight and negatively bowed turbine cascades with a 106° turning angle respectively with and without cooling air injection. The air injection is from single or multiple rows of ten holes (26 schemes in all). Experimental results show that the negatively bowed cascade produces more energy loss than the straight one without cooling air injection. Cooling air injection from the holes on both the pressure and the suction surface near the trailing edge can reduce energy losses in both the straight and the negatively bowed cascades. The increase in loss is less in the negatively bowed cascade than that in the straight one with cooling air injection from holes of multiple rows at the leading edge, whereas cooling air injection from holes of multiple rows at the trailing edge reduces the energy loss in the negatively bowed cascade. © 2004 Wiley Periodicals, Inc. Heat Trans Asian Res, 34(1): 1–8, 2005; Published online in Wiley InterScience ( www.interscience.wiley.com ). DOI 10.1002/htj.20042     

Investigation of film cooling and superposition method for double row dustpan‐shaped holes

Film cooling effectiveness for injection from single row and double row dustpan‐shaped holes was studied experimentally. The difference in cooling performance between single row and double row holes are described. The results show that the optimal blowing ratio of injection from double row holes is higher than that of injection from a single row hole. In the case of higher blowing ratios, the traditional superposition calculation method underpredicts the film cooling effectiveness of the double row injection based on that of the single row hole injection. The modified superposition calculation method is given in this paper. © 2008 Wiley Periodicals, Inc. Heat Trans Asian Res, 37(4): 208–217, 2008; Published online in Wiley InterScience ( www.interscience.wiley.com ). DOI 10.1002/htj.20204