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.     

Analysis on the mechanism of evolutionary process of counter-rotating vortex pair in film cooling based on hybrid thermal lattice Boltzmann method

Massively parallel simulation applying multiple graphic processing units (multi-GPUs) is carried out to perform a deep going investigation on the counter-rotating vortex pair (CVP) in single-jet film cooling based on hybrid thermal lattice Boltzmann method (HTLBM) and large eddy simulation (LES). The mechanism of evolutionary process of CVP and its influence on the Reynolds shear stress and cooling performance are studied in detail. In the simulation, the blowing ratio of coolant jet is kept to BR = 0.5 and the inclined angle is α = 30^°. The Reynolds number based on the crossflow velocity and diameter of the jet hole is Re = 4000. The secondary anti-kidney vortices, tertiary kidney vortices and quaternary anti-kidney vortices are captured, which increase the spanwise coolant-film coverage on the surface of the bottom wall. The size of the primary CVP significantly influences the distribution of Reynolds shear stress R_uv and R_uw. The vortex strength of the primary CVP and the characteristics of minor counter-rotating vortices mainly impact on the streamwise and spanwise distribution of film cooling effectiveness, respectively.     

Film cooling characteristics of rows of round holes at various streamwise angles in a crossflow: Part II. Heat transfer coefficients

Measurements of heat transfer coefficient (h) are presented for rows of round holes at streamwise angles of 30°, 60° and 90° with a short but engine representative hole length (L/D = 4). The study began with a single row of holes with pitch-to-diameter ratios of 3 and 6, followed by two inline and staggered rows for each hole spacing and streamwise inclination, which amount to 105 different test cases in addition to the 21 test cases presented on the single hole [C.H.N. Yuen, R.F. Martinez-Botas, Film cooling characteristics of a single round hole at various angles in a crossflow: Part I. Effectiveness, Int. J. Heat Mass Transfer, in press; C.H.N. Yuen, R.F. Martinez-Botas, Film cooling characteristics of a single round hole at various angles in a crossflow: Part II. Heat transfer coefficients, Int. J. Heat Mass Transfer, in press]. The present investigation is a continuation of the previous work [Yuen and Martinez-Botas, Parts I and II, in press] with the same test facility, operating conditions (freestream Reynolds number, Re_D of 8563, and blowing ratio, 0.33 ⩽ M ⩽ 2), and measurement technique of liquid crystal thermography and the steady-state heat transfer method, therefore the results presented in the form of h/h₀, which is the ratio of heat transfer coefficient with film cooling to that without, are directly comparable. Both local values and laterally averaged ones are presented, the latter refers to the averaged value across the central hole. The corresponding measurements of effectiveness for the rows of holes are presented in a companion paper [C.H.N. Yuen, R.F. Martinez-Botas, Film cooling characteristics of rows of round holes at various angles in a crossflow: Part I. Effectiveness, Int. J. Heat Mass Transfer, submitted for publication]. The low effectiveness observed with the 90° holes in the companion paper [Yuen and Martinez-Botas, submitted for publication] and the relatively large heat transfer coefficient presented here, suggest that the normal injection should only be used in situations where shallower holes are not feasible. The combined performance of effectiveness and heat transfer coefficient suggests that the two inline rows are likely to be advantageous in the film cooling of turbine blades with good coverage per unit mass flow of cooling air and lower thermal stresses due to the smaller heat load.     

Suitability of three different two-equation turbulence models in predicting film cooling performance over a rotating blade

The suitability of three different two-equation turbulence models in predicting film cooling effectiveness on a rotating blade was investigated and they are the commonly used standard k-ε model, the k-ω model and the shear stress transport k-ω model. To fulfill this target, both numerical simulation and the experimental investigation were carried out for a rotating blade having a flat test surface with a 4 mm diameter straight circular cooling hole in 30° inclined injection. The blade rotated at five different speeds of 0, 300, 500, 800 and 1000 rpm. The momentum ratio was set to be 0.285 and the Reynolds (Re_D) number based on the mainstream velocity and hydraulic diameter of the mainstream channel is 1.45 × 10⁵. The averaged density ratio was chosen to be 1.026 with air as both the coolant and the mainstream. Comparison between the numerical work and the experimental results indicated that (1) the rotating speed is the most critical parameter influencing the film cooling effectiveness distributions and the pressure surface could be remarkably different from the suction surface, (2) as for the algebraic averaged film cooling effectiveness, numerical predictions of the three turbulence models all overshoot compared with the experimental results, (3) among the three turbulence models, the standard k-ε model gave the poorest prediction.     

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.     

不同安装角下透平静叶气膜冷却流动特性的试验研究

在直流式低速风洞试验室内,采用粒子图像测速技术（PIV）对带6排气膜冷却孔的透平静叶在不同吹风比M和不同叶片安装角口下的流场结构进行了测量,并对流场速度和湍动能进行了分析．结果表明：吸力面的速度梯度明显大于压力面,并且随着叶片安装角的减小和吹风比的增大,速度梯度逐渐变大;安装角对气膜的贴壁性有影响,当β=70°时,气膜的贴壁性最好,且吸力面的贴壁性强于压力面;湍动能的最大值位于冷气发生弯曲的上部,即主流与射流掺混的边界;随着安装角的减小,在压力面下游远场出现的湍动能集中区的脉动程度逐渐减弱,当β=70°时达到最小,随着安装角的再度减小,该湍动能集中区在吸力面下游出现并逐渐增大．     

Computational study of a new scheme for a film-cooling hole on convex surface of turbine blades

An investigation on a heart-shaped film cooling hole is performed for one row, two staggered rows, and three staggered rows on a convex surface. The results are compared with those obtained from a simple cylinder hole as a baseline. Three-dimensional computational study for two heart-shaped holes of two crown angles, φ = 60° and φ = 90°, is conducted to ascertain adiabatic film cooling effectiveness to validate whether a heart-shaped cooling hole mitigates the vortexes responsible for the lift-off phenomenon in conventional simple cylinder hole. This paper also attempts to show that the staggered arrangement of a heart-shaped hole provides higher cooling performance. The result reveals that a heart-shaped hole highly mitigates the vortexes, thereby providing more coolant-surface attachment. The results reflect the tremendous increment in centerline and lateral adiabatic film cooling effectiveness for both crown angles, φ = 60° and φ = 90°. The heart-shaped hole of crown angle φ = 60° shows higher centerline effectiveness compared with that of the heart-shaped hole crown angle of φ = 90°. The latter provides higher lateral effectiveness. The relatively small volume and high effectiveness of the heart-shaped cooling hole is important and promising for the aero engine industry.     

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.     

Influence of Trenched Shaped Holes on Turbine Blade Leading Edge Film Cooling

Computational results are presented for a row of coolant injection holes on each side of a high-pressure turbine blade near the leading edge. Seven hole configurations have been used to show the effect of various diffusion shaped holes and their trenching on film cooling effectiveness: (1) cylindrical film hole; (2) forward diffused film hole; (3) trenched forward diffused film hole; (4) conically flared film hole; (5) trenched conically flared film hole; (6) laterally diffused film hole; and (7) trenched laterally diffused film hole. Computational solutions of the steady, Reynolds-averaged Navier–Stokes equations are obtained using a finite-volume method. Results show that the main effect of trenching is the reduction of jet lifting off from the blade surface and so the prevention of sudden lowering of cooling effectiveness after the injection location. Moreover, hole trenching has more effect on film cooling flow on the suction side than on the pressure side. Also, the trenched laterally diffused shaped hole has the highest laterally averaged effectiveness on both the suction side and the pressure side of the blade.     

DNS study of pulsed film cooling for enhanced cooling effectiveness

Direct Numerical Simulation (DNS) of a pulsed film-cooling jet is presented to examine if pulsations of the coolant jet can enhance film-cooling effectiveness. Calculations are performed for a cylindrical jet inclined at 30°. The jet pulsation is defined by the duty cycle (DC) and the Strouhal number (St), both of which are varied in this study. Baseline calculations are done for a steady blowing ratio of 1.5. Both frequency and duty cycle are observed to influence the cooling effectiveness. For a peak blowing ratio (M) of 1.5, pulsing with a St = 0.32 and DC = 0.5 is shown to reduce jet blow-off and improve centerline and spanwise-averaged effectiveness over the steady M = 1.5 case.     

Mist film cooling simulation at gas turbine operating conditions

Air film cooling has been successfully used to cool gas turbine hot sections for the last half century. A promising technology is proposed to enhance air film cooling with water mist injection. Numerical simulations have shown that injecting a small amount of water droplets into the cooling air improves film-cooling performance significantly. However, previous studies were conducted at conditions of low Reynolds number, temperature, and pressure to allow comparisons with experimental data. As a continuous effort to develop a realistic mist film cooling scheme, this paper focuses on simulating mist film cooling under typical gas turbine operating conditions of high temperature and pressure. The mainstream flow is at 15 atm with a temperature of 1561 K. Both 2D and 3D cases are considered with different hole geometries on a flat surface, including a 2D slot, a simple round hole, a compound-angle hole, and fan-shaped holes. The results show that 10–20% mist (based on the coolant mass flow rate) achieves 5–10% cooling enhancement and provides an additional 30–68 K adiabatic wall temperature reduction. Uniform droplets of 5–20 μm are used. The droplet trajectories indicate the droplets tend to move away from the wall, which results in a lower cooling enhancement than under low pressure and temperature conditions. The commercial software Fluent is adopted in this study, and the standard k–ε model with enhanced wall treatment is adopted as the turbulence model.     

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.     

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.     

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.     

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.     

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.     

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.     

Flow-field analysis of anti-kidney vortex film cooling

Film cooling is an important measure to enable an increase of the inlet temperature of a gas turbine and, thereby, to improve its overall efficiency. The coolant is ejected through spanwise rows of holes in the blades or endwalls to build up a film shielding the material. The holes often are inclined in the downstream direction and give rise to a kidney vortex. This is a counter-rotating vortex pair, with an upward flow direction between the two vortices, which tends to lift off the surface and to locally feed hot air towards the blade outside the pair. Reversing the rotational sense of the vortices reverses these two drawbacks into advantages. In the considered case, an anti-kidney vortex is generated using two subsequent rows of holes both inclined downstream and yawed spanwise with alternating angles. In a previous study, we performed large-eddy simulations (which focused on the fully turbulent boundary layer) of this anti-kidney vortex film-cooling and compared them to a corresponding physical experiment. The present work analyzes the simulated flow field in detail, beginning in the plenum (inside the blade or endwall) through the holes up to the mixture with the hot boundary layer. To identify the vortical structures found in the mean flow and in the instantaneous flow, we mostly use the λ 2 criterion and the line integral convolution (LIC) technique indicating sectional streamlines. The flow regions (coolant plenum, holes, and boundary layer) are studied subsequently and linked to each other. To track the anti-kidney vortex throughout the boundary layer, we propose two criteria which are based on vorticity and on LIC results. This enables us to associate the jet vortices with the cooling effectiveness at the wall, which is the key feature of film cooling.     

Experimental investigation on film cooling performance of pressure side in annular cascades

Experimental investigations were conducted to study the film cooling performance in a low speed annular cascades using Thermochromic Liquid Crystal (TLC) technique. The test blade was placed in the second stage, where 18 blades were installed with chord length of 124.3 mm and height of 99 mm. A film hole with diameter of 4 mm, angled 28° to the tangential of the pressure surface in streamwise, was set in the middle span of the blade. The Reynolds number based on the outlet mainstream velocity and the blade ...