Numerical investigations of cooling holes system role in the protection of the walls of a gas turbine combustion chamber

Numerical simulations in a gas turbine Swirl stabilized combustor were conducted to investigate the effectiveness of a cooling system in the protection of combustor walls. The studied combustion chamber has a high degree of geometrical complexity related to the injection system as well as the cooling system based on a big distribution of small holes (about 3,390 holes) bored on the flame tube walls. Two cases were considered respectively the flame tube without and with its cooling system. The calculations were carried out using the industrial CFD code FLUENT 6.2. The various simulations made it possible to highlight the role of cooling holes in the protection of the flame tube walls against the high temperatures of the combustion products. In fact, the comparison between the results of the two studied cases demonstrated that the walls temperature can be reduced by about 800°C by the mean of cooling holes technique.     

Effect of film cooling on the aerodynamic performance of an airfoil

The effect of film cooling on the aerodynamic performance of turbine blades is becoming increasingly important as the gas turbine operating temperature is being increased in order to increase the performance. The current paper investigates the effect of blowing ratio on the aerodynamic losses of a symmetric airfoil by pressure measurements and Particle Image Velocimetry (PIV). The test model features 4 rows of holes located on the suction side at 5%, 10%, 15% and 50% of the chord length. The Reynolds number based on the airfoil chord is 1.2 × 10⁵. Experiments are performed by varying the location of air injection, the angle of attack, and the mainstream velocity. The coolant air is injected at ambient temperature and the blowing ratio is varied from 0 to 1.91. It is observed that the losses due to film cooling increase with blowing ratio of 0 to 0.48, and the wake is shifted towards the suction side. Conversely, the aerodynamic losses decrease when the blowing ratio is increased further from 0.64 to 1.91. This trend has been observed for all the experimental configurations. The effect of blowing ratio on flow separation is investigated with the time-averaged velocity fields obtained from PIV measurements. It is observed that low blowing ratios, the separation point shifts upstream and at high blowing ratios the ejected coolant energizes the flow and delays separation. The pressure field around the airfoil is reconstructed from the integration of the Poisson equation based on the PIV velocity fields. The experimental results can be used for validation of numerical models for predicting losses due to film cooling.     

Berechnung der Störung der Brennkammer-Filmkühlung durch Einblasen von Luftstrahlen mit Hilfe einer Analogiebetrachtung zwischen Strahl und Zylinder

The influence of cold air jet injection upon the film cooling of combustion chamber walls is investigated theoretically and experimentally. Considering the cold air jet acting like a cylinder in the combustion gas and film cooling streams, an equation can be derived theoretically which describes the disturbance of the film cooling effectiveness downstreams of the air jet injection. The experiments, where single cold air jets are injected normal to the film cooled test chamber wall, show considerable decreases of the film cooling effectiveness downstream of the injection points. The same effect is noted in real combustion chambers of aircraft gas turbine engines. Additional experiments with cylinders simulating the disturbance effect of the air jets prove the analogy consideration between jet and cylinder. The theoretically predicted cooling effectiveness is in agreement with own experiments as well as other test results.     

Aerodynamic performance of an annular cascade of film cooled nozzle guide vanes under engine representative conditions

This paper addresses two important issues relevant to efficiency measurements in film-cooled annular cascades: the definition of the ideal flow to be used in loss calculation, and the measurements that are necessary for such loss calculation. The paper also addresses the question of the correct parameterisation of coolant density effects, showing that the momentum flux ratio, rather than the blowing rate, is the most appropriate parameter. Experiments examining the effect of extensive aerofoil surface film cooling on the aerodynamic efficiency of an annular cascade of transonic nozzle guide vanes are reported. A dense foreign gas (SF₆/Ar mixture) is used to simulate engine representative coolant-to-mainstream density ratios, momentum ratios and blowing rates under ambient temperature conditions. Experiments are also conducted with air coolant to study the effect of density ratio on efficiency. The flowfield measurements have been obtained using a four-hole pyramid probe in a short duration blowdown facility which correctly models engine Reynolds and Mach numbers. This work compares the measured aerodynamic efficiencies of uncooled vanes with those which employ an extensive amount of cooling. The engine-representative cooling geometry investigated features 14 rows of cylindrical cooling holes, and a second geometry where 8 of these rows are replaced by holes having a fan-shaped exit. The effects of adding trailing edge slot ejection are also presented. By selectively blocking rows of holes, the cumulative effect on the mid-span efficiency of adding rows of cooling holes has also been determined. Experimental results are presented as area traverse maps (total pressure, isentropic Mach number and flow angles), from which the relative changes in efficiency due to film cooling have been calculated. These calculations reveal that the presence of foreign-gas coolant from the cylindrical-hole geometry increases the aerodynamic loss (relative to the uncooled baseline) by 6.7%; and coolant from the fan-shaped holes increases the loss by 15%. The effects of different assumptions for the coolant total pressure are shown to significantly change the measured loss; if the loss measurements are based on the mainstream total pressure, rather than the total pressure in the coolant cavity, the respective increase in loss (relative to the uncooled baseline) of cylindrical and fan-shaped holes is 4.5% and 12.5%. Experimental data is compared with loss predictions using a Hartsel model. Received: 4 December 1998/Accepted: 1 September 1999     

Investigation of mixing processes of effusion cooling air and main flow in a single sector model gas turbine combustor at elevated pressure

Mixing processes between main flow and effusion cooling air are investigated in an effusion cooled, swirl-stabilized pressurized single sector gas turbine combustor using advanced laser diagnostics. Quantitative planar laser-induced fluorescence of the hydroxyl radical (OH-PLIF) and planar laser-induced fluorescence of nitric oxide, seeded to the effusion cooling air, (NO-PLIF) are employed in the primary zone and close to the effusion cooled liner. This data is used to identify mixing events at three stages of premixed combustion, i.e. mixing before reaction, mixing during reaction and mixing after reaction. A parametric study of swirl and cooling air mass flow is conducted to investigate the mutual interaction between flame and cooling air. Within the primary zone, a significant radial asymmetry of OH concentration is observed. This asymmetry is partly explained by the presence of effusion cooling air within the unburned fresh gas, leading to lowered OH concentration within local reaction zones and their post-flame equilibrium concentration. Near the effusion cooled liner, adiabatic mixing after reaction is the dominant process across all investigated operating conditions. Notable mixing before reaction is only observed for the first effusion hole on the center line at low swirl conditions.     

汽油机燃烧室内温度场测量的激光剪切干涉法

采用激光剪切干涉法测量了汽油机燃烧室内的温度场,并分析了剪切干涉法测温的基本原理,推导出利用干涉条纹图求解温度分布的关系式。在一台二冲程火花点火发动机上设置石英窗和信号同步系统,建立了适合高速摄影的激光剪切干涉测量装置。通过改变剪切干涉量,测取发动机燃烧室内干涉条纹图,从而获得缸内燃烧的二维温度场。结果表明,激光剪切干涉法抗振性强,光路简单可靠,可以进行高速摄像,是研究内燃机燃烧过程的有效方法。从温度场可以看出,燃烧过程中缸内大致可分为三个区,即已燃区、未燃区和燃烧区,具有不同的温度分布和温度梯度。燃烧区温度最高,温度梯度大;已燃区温度次之,梯度较小;未燃区温度最低,但梯度较大。     

LES Investigation of the Flow Through an Effusion-Cooled Aeronautical Combustor Model

The present study is devoted to the analysis of the behaviour of the flow through an effusion-cooled aeronautical combustor model. High-fidelity calculations are performed on an experimental model of a combustion chamber multi-perforated wall and compared to experimental measurements. The effect of combustion instability on the effusion-cooling system is investigated by studying the interaction of an acoustic wave with the jets-in-crossflow issued from the cooling plate. It is shown that the mass-flow rate through the plate can be drastically reduced by the acoustic wave, which demonstrates the destructive effect that such instability may have on the cooling of an aeronautical combustion chamber.     

A parametric investigation of vane pressure side cutback film cooling by dual luminophor PSP

Because of its geometry, the vane trailing edge is one of the most difficult regions to be cooled. A trailing edge cutback cooling system is one of the most effective solution for cooling the trailing edge of high-pressure gas turbine nozzle vanes. In this study, a parametric analysis of the thermal performance of a nozzle vane cascade with a pressure side cooling system including two rows of cylindrical holes and a trailing edge cutback featuring 8 rectangular slots was carried out by using dual luminophor (binary) PSP technique. Coolant to mainstream mass flow rate (MFR), density ratio (DR), main flow Mach number (Ma_2is) and turbulence intensity level (Tu₁) and the state of the approaching boundary layer were the considered parameters. Binary PSP was able to measure the coolant concentration independently from temperature, thus allowing to compute the true adiabatic film cooling effectiveness in a complex environment. MFR was shown to have a strong impact on both holes and cutback performance. The thermal protection over the cutback region was promoted by high Ma_2is and high DR values, while Tu₁ and the boundary layer state only marginally affected the thermal behavior, especially at high MFR.     

Numerical simulations of heat transfer distribution of a two-pass square channel with V-rib turbulator and bleed holes

The blade tip region in gas turbine encounters high thermal loads due to temperature difference and hence efforts for high durability and safe operations are essential. Improved and robust methods of cooling are required to downgrade heat transfer rate to turbine blades. The blade tip regions, which are exposed to high gas flow, suffers high local thermal load which are due to external tip leakage. Jet impingement, pin cooling etc. are techniques used for cooling blades. A more usual way is to use serpentine passage with 180-degree turn. In this study, numerical simulation of heat transfer distribution of a two-pass square channel with rib turbulators and bleed holes were done. Periodical rib turbulators and bleed holes were used in the channel. The ribs arrangement were 60 degree V rib, 60 degree inverted V ribs, combination of 60 degree V rib at inlet and 60 inverted V rib at outlet section and combination of Inverted V at inlet and V rib at the outlet. The results were numerically computed using Fluent with Reynolds number of 12,500 and 28,500. Turbulence models used for computations were k-ω-SST and RSM. Temperature based and shear stress based techniques were used for heat transfer distribution prediction. The results for 60 degree V rib, 60 degree inverted V ribs were compared with the experimental results for validation of the results obtained. Detailed distribution shows distinctive peaks in heat transfer around bleed holes and rib turbulator. Comparisons of the overall performance of the models with different orientation of rib turbulator are presented. It is found that due to the combination of 60 degree inverted V rib in inlet and 60 V rib in outlet with bleed holes provides better heat treatment. It is suggested that the use of rib turbulator with bleed holes provides suitable for augmenting blade cooling to achieve an optimal balance between thermal and mechanical design requirements.     

Film cooling on a convex surface: influence of external pressure gradient and Mach number on film cooling performance

 The film cooling performance on a convex surface subjected to zero and favourable pressure gradient free-stream flow was investigated. Adiabatic film cooling effectiveness values were obtained for five different injection geometries, three with cylindrical holes and two with shaped holes. Heat transfer coefficients were derived for selected injection configurations. CO₂ was used as coolant to simulate density ratios between coolant and free-stream close to gas turbine engine conditions. The film cooling effectiveness results indicate a strong dependency on the free-stream Mach number level. Results obtained at the higher free-stream Mach number show for cylindrical holes generally and for shaped holes at moderate blowing rates significant higher film cooling effectiveness values compared to the lower free-stream Mach number data. Free-stream acceleration generally reduced adiabatic film cooling effectiveness relative to constant free-stream flow conditions. The different free-stream conditions investigated indicate no significant effects on the corresponding heat transfer increase due to film injection. The determined heat flux ratios or film cooling performance indicated that coolant injection with shaped film cooling holes is much more efficient than with cylindrical holes especially at higher blowing rates. Heat flux penalties can occur at high blowing rates when using cylindrical holes. Received on 29 May 2000     

On the laws of combustion wave suppression by free water in a homogeneous porous layer of organic combustible materials

The conditions of suppression of a combustion wave propagating along a homogeneous porous layer of organic combustible materials are investigated basing upon a physico-mathematical model. The dependence of free water effervescence intensity on the water volume fraction and the medium temperature is presented. A water supply source in motion is employed to suppress the combustion wave. It is shown that an increase in the mass water flow rate results in a considerable increase in the maximum velocity of the source motion, at which the combustion wave can be suppressed. This is due to a reduction in the loss of the water evaporating above the burning zone. The effect of the water supply point displacement on the efficiency of combustion wave suppression is analyzed.     

点火过程和初始条件对燃烧轻气炮内弹道性能的影响

采用计算流体力学方法对燃烧轻气炮膛内燃烧过程进行数值模拟,分析不同的点火点数目和点火能量以及初始温度和压力对燃烧轻气炮内弹道特性的影响。结果表明,采用合理的点火点数目、初始温度和压力条件可以有效控制氢气的燃烧过程,减弱燃烧室内的压力波动。模拟结果对燃烧轻气炮膛内燃烧过程控制具有重要参考价值。     

Film cooling and heat transfer from a combination of two rows of simple and/or compound angle holes in inline and/or staggered configuration

This paper describes the results of an experimental investigation into the film cooling effectiveness and heat transfer characteristics of two staggered injection rows of either a combination of one row of simple angle holes with another row of compound angle holes or with both rows of compound angle holes. The effect of using various injections holes arrangements as well as the relative location of the compound angle holes row to the simple angle holes row have been investigated for different blowing rates. Using combination of one row of downstream compound angle holes with another upstream simple injection holes row provides a significant increase in the film cooling protection over a flat plate surface, over that obtained from either two rows of only simple injections holes or compound angle holes. Received on 17 July 1998     

当量比对预混燃烧影响的大涡模拟研究

本文对在突扩燃烧室内甲烷和空气的预混燃烧进行了大涡模拟(LES)研究,考虑预混燃料的当量比对燃烧室提供的动力及产生的污染物的影响．利用LES计算了不同当量比条件下燃烧室内湍流预混燃烧反应流场的温度、浓度、涡量和压力分布,最后对当量比0.5时B点和C点的温度和速度进行EMD分解,得到了温度场和速度场的各阶模态的平均周期．结果表明：随着当量比从0.5增加至0.7,燃烧反应趋于剧烈,燃烧室的最高温度提高了350K,平均压力从32.876 Pa增大到34.833Pa,燃烧产生的瞬态径向最高浓度从0.5%增加到0.95%．     

Large area impingement spray cooling from multiple normal and inclined spray nozzles

An inclined spray chamber with four multiple nozzles to cool a 1 kW 6U electronic test card has been designed and tested in this study. The multiple inclined sprays can cover the same heated surface area as that with the multiple normal sprays but halve the volume of the spray chamber. The spray cooling system used R134a as a working fluid in a modified refrigeration cycle. It is observed that increasing mass flow rate and pressure drop across the nozzles improved the heat transfer coefficient with a maximum enhancement of 117 %, and reduced the maximum temperature difference at the heated surface from 13.8 to 8.4 °C in the inclined spray chamber with a heat flux of 5.25 W/cm², while the heat transfer coefficient of the normal spray increased with a maximum enhancement of 215 % and the maximum temperature difference decreased from 10.8 to 5.4 °C under similar operating conditions. We conclude that the multiple inclined sprays could produce a higher heat transfer coefficient but with an increase in non-uniformity of the surface temperature compared with the multiple normal sprays.     

Considerations for the design of swirl chambers for the cyclone cooling of turbine blades and for other applications with high swirl intensity

The focus of this study lies on turbulent incompressible swirling flows with high swirl intensity. A systematic parameter study is conducted to examine the sensitivity of the mean velocity field in a swirl chamber to changes in the Reynolds number, swirl intensity and channel outlet geometry. The investigated parameter range reflects the typical kinematic flow conditions found in heat transfer applications, such as the cooling of the turbine blade known as cyclone cooling. These applications require a swirl intensity, which is typically much higher than necessary for vortex breakdown. The resulting flows are known as flow regime II and III. In comparison to flow regime I, which denotes a swirling flow without vortex breakdown, these flow regimes are characterized by a subcritical behavior. In this context, subcritical means that the flow is affected by the downstream channel section. Based on mean velocity field measurements in various swirl chamber configurations, it is shown that flow regime III is particularly sensitive to these effects. The channel outlet geometry becomes a determining parameter and, therefore, small changes at the outlet can produce entirely different flow patterns in the swirl chamber. In contrast, flow regime II, as well as flow regime I and axial channel flows, are much less sensitive to changes at the channel outlet. The knowledge about the sensitivity of the flow in different flow regimes is highly relevant for the design of a cyclone cooling system. Cooling systems employing flow regime III can result in a weakly robust flow system that may change completely over the operating range. As a remedy, the swirl intensity needs to be decreased so that flow regime III cannot be reached, which, however, reduces the maximum achievable heat transfer in the cooling system. Alternatively, the flow has to transition back from flow regime III to flow regime II or I before the flow leaves the swirl chamber. Two practical methods are presented. These findings can be directly applied in the design processes of future cyclone cooling systems, and other applications of swirling flow.     

Numerical investigation of the influence of incidence angle on asymmetrical turbine blade model showerhead film cooling effectiveness

The influence of various incidence angles on film cooling effectiveness of an axial turbine blade cascade with leading edge ejection from two rows of cooling holes is numerically investigated. The rows are located in the vicinity of the stagnation line. One row is located on the suction side and the other one is on the pressure side. The predicted pressure field for various blowing ratios (M = 0.7, 1.1 and 1.5) is compared with available experimental results at the design condition. Moreover, the effect of various incidence angles (?10°, ?5°, 0°, 5° and 10°) at three blowing rates is investigated by analyzing the results of both laterally averaged and area averaged values of adiabatic film cooling effectiveness. Numerical results indicate that the incidence angle can strongly affect the thermal protection of the blade at low blowing ratio but becomes less dominant at high blowing ratio. In fact, for the low blowing ratio, a small change in the incidence angle that relates to the design condition can deeply affect the thermal protection of the blade, which is evident from the laterally and area averaged film cooling effectiveness distributions.     

Experimental study of biogas combustion using a gas turbine configuration

The aim of the present work is to compare stability combustion domains, flame structures and dynamics between CH₄/air flames and a biogas/air flames (issued from waste methanisation) in a lean gas turbine premixed combustion conditions. Velocity profiles are obtained by Laser Doppler Anemometry measurements. CH* chemiluminescence measurements and temporal acquisition of chamber pressure are performed in order to describe flame structure and instabilities. Changes in flame structure and dynamics when fuel composition is varying are found to strongly depend on laminar flame speed. No clear correlation between the unstable flame and the reaction zone penetration in the corner recirculation can be found.     

Detached eddy simulation of blade trailing-edge cutback cooling performance at various ejection slot angles

Detached eddy simulation (DES) has been carried out to study a three-dimensional trailing-edge (TE) cutback turbine blade model with five rows of staggered circular pin-fin arrays inside the cooling passage, in order to evaluate the cooling performance in relation to coolant ejection slot angle. Simulations were performed by adopting a shear-stress transport k-ω turbulence model, and the effects of three different ejection slot angles 5°, 10° and 15° were investigated in terms of the characteristics of adiabatic film-cooling effectiveness, coefficient of discharge, and vortex shedding frequencies, respectively. The results obtained have shown that the TE cutback blade cooling with a 5° coolant ejection slot angle produced a better heat transfer coefficient than the other two ejection slot angles tested. The distributions of adiabatic film-cooling effectiveness along the cutback walls were found to be sensitive to the coolant ejection slot angle, e.g. the increase of ejection slot angle to 15° yielded near unity of cooling effectiveness along the entire breakout walls, whereas the decrease of ejection slot angle caused a drastic decay of cooling effectiveness after the maximum effectiveness has been reached. Of the three angles studied, a TE cutback blade model with a 15° ejection slot angle produced an optimum film-cooling effectiveness. In the breakout region, vortex shedding was observed along the shear layer between the hot gas and the coolant airflow. The shedding frequencies were evaluated to be 2.93, 2.21, and 2.18 kHz for the ejection slot angles of 5°, 10° and 15°, respectively. The findings from this study could be useful to improve existing TE cutback turbine blade design to achieve optimum film-cooling performance.     

Spreadsheet calculations of jets in crossflow: opposed rows of inline and staggered round holes

The objective of this study was to demonstrate and analyze empirical model results for jet-in-crossflow configurations which are typical in gas turbine combustors. Calculations in this paper, for opposed rows of round holes in both inline and staggered arrangements, were made with an Excel^® spreadsheet implementation of a NASA-developed empirical model for the mean conserved scalar field. Results for cases of opposed rows of jets with the orifices on one side shifted by half the orifice spacing shows that staggering can improve the mixing, particularly for cases that would overpenetrate if the orifices were in an aligned configuration. For all cases investigated, the dimensionless variance of the mixture fraction decreased significantly with increasing downstream distance. The variation between cases at a given downstream location was smaller, but the “best” mixers for opposed rows of jets were found to be inline and staggered arrangements at an orifice spacing that is optimum for inline jets.