Dynamic and Thermodynamic Analysis of Film-Cooling

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International Review of Mechanical Engineering (I.RE.M.E.), Vol. 7, N. 3

ISSN 1970 - 8734 March 2013

Dynamic and Thermodynamic Analysis of Film-Cooling

Nor Azwadi Che Sidik, Ehsan Kianpour, Iman Golshokouh

Abstract – This study was done to extend database knowledge about the film cooling action at the end of combustor and inlet of turbine. The well-known Brayton cycle plays a key role to get higher engine efficiency in gas turbine engines. But the high temperature of the combustor exit flow creates unpleasant environment. These surroundings lead to a reduction in the expected life of critical parts. There are two separate ways for Gas turbine cooling: internal cooling and external cooling. Film cooling is one of the most effective external cooling methods. In this system, a low temperature thin boundary layer such as buffer zone is formed and attached on the protected surface. In this study, a literature survey was done on the limited surveys that considered the effects of flow structure variations on film cooling, particularly from the first of 21th century.

Keywords:Gas Turbine Engine, Combustor, Film Cooling, Cooling Holes, Turbine Vanes

I. Introduction

The propulsion science was grown since 150 B.C.; In that time, mathematicians and philosophers named Hero designed the first steam engine –the aeolipile– (Fig. 1).

In this engine, a simple closed spherical vessel was set up on bearings and this mechanism allowed it to have rotational movement due to the exerted tangential forces which are created by the steam discharge of the nozzles.

Thereafter, in 1930, Frank Whittle received the patent for the first gas turbine engine invention. The Newton`s third law of motion stated that to every action there is always opposed and equal reaction. This is the fact of flight as we know it today.

The famous Brayton cycle is a key to achieve higher turbine gas engine efficiency and power to weight ratio.

To increase the efficiency, this thermodynamic cycle expressed that the outlet temperature of the combustor should increase [1].

However, the operating temperature is such above that all materials cannot resist against this value of temperature [2]. In addition, turbine inlet flow temperature augmentation provides an extremely unpleasant environment at the end of combustor and inlet of turbine. The incidence of such condition can destroy the critical components downstream the combustor (Figure 2). Therefore, a cooling technique must be applied to prevent the thermal degradation of turbine components. While, the early gas turbine engines functioned at temperature range of 1200℃ to 1500℃, the advanced engines operated at the turbine inlet temperature of 1950℃ to 2010°C. However, the turbine inlet temperature increased above 2000°C with new patterns of cooling since the first of 21^th century [3].

Gas turbine cooling classified into two different schemes: internal cooling and external cooling.

In the internal cooling method, coolant provided by the compressor, is forced into the cooling flow circuits inside turbine components.

In the external way, the injected coolant is directly perfused from coolant manifold to save downstream components against hot gases. In the external cooling, coolant is used to quell the heat transfer from hot gas stream to a component.

External cooling contains several ways. Film cooling is the most well-known method of preservation. In this system, a low temperature thin boundary layer such as buffer zone is formed and attached on the protected surface (Fig. 3).

The change of flow field parameters is a topic which motivated researchers among years.

Fig. 1. Schematic of aeolipile engine

Fig. 2. Schematic view of the turbine first vane damage

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Nor Azwadi Che Sidik, Ehsan Kianpour, Iman Golshokouh

Therefore, this investigation increases the basement knowledge about the film cooling and the effects of dynamic and thermodynamic characteristics variations on cooling performance.

Fig. 3. Schematic view of film cooling

II. Effects of the Flow Structure Changes

Zhi et al. [4] measured the film cooling heat transfer coefficient distribution for two different conditions of stationary and rotating cases (Fig. 4). They calculated the experimental uncertainty by using previous studies such as Kline McClintock [5]. The application of air and carbon dioxide as a coolant provide two averaged velocity ratios used in this study. Adjacent the cooling holes, the blowing ratio changes have remarkable effect on heat transfer.

A dominant effect on the stream wise heat transfer coefficient is seen under density ratio variations, particularly for rotating cases. The results indicated that the usage of air and carbon dioxide as a coolant influence the effects of blowing ratio on the heat transfer coefficient.

Fig. 4. Flat blade arrangement

Ou et al. [6], [7] showed that for carbon dioxide, heat transfer coefficient is increased at first and then decreased with blowing ratio enhancement, while for the air injection, heat transfer coefficient has fluctuations. In addition, the density ratio affected the distributions of the stream wise heat transfer coefficient significantly.

Ethridge et al. [8] experimentally studied the film cooling effectiveness on the suction surface of first stage turbine vane. The experimental results showed that at low blowing ratios, increasing turbulence level has no significant effect on film cooling, however, at high blowing ratios, the coolant not attached well on the

protected surface and thereby film cooling effectiveness reduction occurred (Fig. 5). These results rejected the Sarkar and Bose [9]. Because in a numerical study, they stated that at low blowing ratios, 25% reduction in adiabatic film cooling effectiveness happened when the free stream turbulent increased about 2 to 12 percent.

Fig. 5. The variation of lateral average effectiveness

Using a two dimensional PIV, Wright et al. [10]

studied the cooling holes structure on film cooling performance for cylindrical holes (Fig. 6). The tests were accomplished under different turbulent intensities and blowing ratios. The results indicated that for both turbulent intensities, flow attached to the surface at lowest blowing ratio of 0.5. In concurred with Bazdidi Tehrani and Mahmoodi [11] it is found that at blowing ratio of M=0.50 and inclination angle of 35deg, the film cooling effectiveness reached to optimum value. At the end of cooling holes, the jet runs especially at higher blowing ratios of M=1.0 and M=1.50. By the injection holes, the horse shoe vortex and counter rotating vortices are found due to the existence of stream wise flow field.

Sarkar and Bose [9] numerically studied the coolant jet aerodynamic in a hot cross flow. They used k−ε turbulent model to solve the Navier Stokes equation. The results showed that adiabatic film cooling performance on the downstream zone of combustor and the aerodynamic specifications of injected coolant are affected by the coolant jet blowing ratio. In addition, blowing ratio enhancement thickened the thermal layer, thereby modified cooling process. This is in contrast with Bernsdorf et al. [12] because he showed that blowing ratio and density ratio increase intensified the entrainment of coolant jets and therefore, coolant boundary layer becomes thinner.

Scrittore et al. [13] measured the flow characteristics inside a combustor simulator (Fig. 7). For uncertainty analysis, like many previous researches, they used Moffat [14] partial derivative and sequential perturbation methods. The experimental results showed that at the investigated range of blowing ratios, the injected cooling flow height is not dependent to the momentum flux ratio.

Furthermore, there is no significant difference between the boundary layer effectiveness augmentations of for all blowing ratios, however just little modification is seen at large blowing ratios in comparison with lower values.

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Fig. 6. Geometry of cylindrical film cooling holes

Fig. 7. Schematic view of the test plate

In order to measure the time averaged U-Velocity and kinetic energy profiles downstream the jet exit, Farhadi et al. [15], studied the influence of flow hydrodynamic and film cooling performance for two small cooling jets exactly before of the main jet.

In this study, they studied the influences of density and velocity variations among the jets and cross flow as well as each of the jets.

The findings showed that when the main flow density ratio is lesser than cooling jets, the adiabatic film cooling effectiveness and span wise averaged is modified further and near the jet.

Maiteh and Jurban [16] experimentally investigated the effect of pressure gradient on film cooling performance. Both pressure gradients of −1 × 10^-6 to 1.11

× 10^-6 used in this study. The results indicated that suitable pressure gradient raise the effect of dilution jets.

Rozati and Danesh Tafti [17] used Large Eddy Simulation to study the effects of the coolant blowing ratios on film cooling performance at the blade leading edge. They considered the variety of blowing ratios of 0.4, 0.8 and 1.2. The results indicated that at blowing ratio of 0.4, three important vortexes of preliminary entrainment vortex, vortex tubes and hairpin vortex are found due to the main stream-coolant turbulence interaction. At this blowing ratio, the entrainment of vortexes is demonstrated due to the contribution of forward and aft vortex tubes. At blowing ratio of 0.8 and 1.2, it does not seem to have longer vortex tubes, while the preliminary vortex is seen at this region. While, the test results illustrated that at blowing ratio of 1.20, the performance is higher than obtained effectiveness at blowing ratio of 0.80, the numerical findings showed opposite of this statement.

Baldauf et al. [18] accomplished an experimental set up to study the heat transfer coefficient distribution on a flat blade after a row of cylindrical cooling holes with d=5mm by using high resolution infrared thermograph technique. In particular, blowing and density ratio variations were investigated. Experimental results showed that further far from the situation without coolant ejection, the heat transfer coefficient is strongly relating to the blowing ratio (Fig. 8).

Chanteloup and Bolcs [19] constructed an experimental set up to find out the behavior of flow at the position with 180deg curve and after it. This test was accomplished whereas the coolant was developing. The turbulence and mean velocity of the flow was determined by measuring the three velocity components with PIV technique.

The results indicated that above the ribbed surfaces, the distribution of heat transfer is improved by stream wise symmetrical profile.

Fig. 8. Spatially averaged net heat flux reduction

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The affection on the stream wise velocity profile is much susceptible especially under the movement of secondary flow. Where the extraction happens, the imbalance condition for the stream wise symmetrical profile is seen into the wall.

By using the large-scale, low-speed experiments, Rowbury et al. [20], [21] explain unexpected flow- interaction phenomena witnessed during annular cascade studies into the influence of external cross flow on film cooling hole discharge coefficients. They investigated the effects of hole geometry, pressure ratio across the hole, Reynolds number and Mach number on discharge coefficient of coolant.

The results declared that near the end of injection hole with external cross flow, the static pressure loss relative to the assumed value led to the discharge coefficient enhancement (Fig. 9).

Ames et al. [22] studied the heat transfer distribution around the vane endwall for a low turbulent condition and aero derivative combustion system. A large-scale, low-speed linear cascade was analyzes under a range of turbulence condition. The results indicated that at low turbulence level, the heat transfer contours are affected by secondary flows. In addition, the influence of the secondary flows on surface heat transfer augmentation and the effect of the passage vortex and horseshoe suction surface leg of are distinguished by the leading edge horseshoe vortex and trailing edge wake survey.

Kwak and Je-Chin [23] studied the heat transfer and film cooling effectiveness distribution under a range of blowing ratios on a gas turbine blade tip with different gap clearances by using TLC technique. The results declared that the blockage of the cooling lead to blowing ratio enhancement; thereby the tip heat transfer coefficient decrease and static pressure increase above the shroud is found. Shaohua et al. [24] surveyed the effects of blowing ratio variation on heat transfer by k−ε turbulence model.

The results showed that among blowing range of 0.6 to 2.0, the cooling performance is relating to the blowing ratios. From the current research and Murata et al. [25]

which studied the film cooling performance at the trailing edge of turbine vane experimentally, it is found that the maximum cooling efficiency and the best heat transfer are achieved at higher blowing ratios.

Fig. 9. Comparison between with-crossflow measurements and additive loss coefficient predictions

This statement is rejected the Bunker [26]

experimental data. They studied the effects of different cooling holes configurations at the turbine vane entrance and said “low variability in heat transfer coefficient enhancement with blowing ratio”. In addition Bunker stated that blowing ratio variation from 0.5 to 2.0 has moderate effectiveness on film cooling performance.

Cun-liang et al. [27], studied the influence of waist- shaped slot hole on film cooling performance experimentally and numerically. The results declared that the momentum flux ratio enhancement causes to the normalized heat transfer coefficient increase as well as the optimum laterally averaged film cooling effectiveness is achieved at mass flux ratio of 2.

Aga et al. [28] conducted an experimental analysis to validate and calibrate the 3-dimensional computational data. This study was done on one hole in a row of cooling holes lateral angle of 45deg and inclination angle of 20deg (Fig. 10). They stated that blowing ratio enhancement deviate the coolant from the initial injection axis due to the vortices and entrainment effects.

Furthermore, thinning the boundary layer and as a result increased the wall heat transfer.

Saumweber et al. [29] determined the two dimensional adiabatic film cooling effectiveness and discharge coefficients.

They considered a range of Mach number up to 0.60 and blowing ratio from 0.50 to 1.50. In line with Dittmar et al. [30] and Elnady et al. [31], it is claimed that film cooling effectiveness is enhanced at moderate and elevated blowing ratios downstream the fan-shaped holes and at the leading edge of gas turbine vane. Also, coolant cross flow is deviated vertically to the hole axis. In addition, for fan-shaped hole film cooling performance is influenced due to the vertical movement of coolant cross flow.

The time-averaged performance study shows the advantages of cooling configurations. Fawcett et al. [32]

focused on the types of unsteadiness and its effect on the flow interaction.

Fig. 10. Definition of geometrical Parameters

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In this test, they measured the effects of blowing ratios by PIV and high speed photography technique for both cylindrical and fan-shaped holes. Within the studied limit of blowing ratios, at the end of cooling holes, a laminar flow is seen for cylindrical jet. For the fan shaped hole and the blowing ratio more than unity, the flow is turbulent. However, for both types of cooling holes, no lateral unsteadiness was seen in the jet.

Saumweber et al. [33] presented the experimental results of free stream turbulence with different cooling holes. Three different free stream turbulence intensities of 3.5, 7.5 and 11 were considered. They found that within a range of blowing rates from low to high values, free stream turbulence increase reduced film cooling performance.

For instance, when turbulence intensity increased from 3.6 to 11 percent, the local and laterally averaged performance decreased up to 40% and 25% respectively (Fig. 11). While for the cylindrical holes, the heat transfer varying was not dependent to the blowing ratios at higher free stream turbulence, this is reversed for the shaped cases. Later, Bunker [26] in the Electronic Global Research Centre tested the effects of blowing ratios, compound angle, coolant flow characters and main stream turbulence intensities. Heat transfer coefficient augmentation has moderate sensitive with blowing ratio as well as turbulence intensity.

Colban et al. [34], [35] took measurements on a two- passage cascades to investigate the hole injection film cooling effectiveness. According to Peng and Jiang [36]

findings, it is debated that the fan-shaped holes raised the film cooling effectiveness. Furthermore, Lutum et al.

[37] and Gao et al. [38] stated that by 75 percent film cooling performance is found for the shaped holes at higher blowing ratios compared to cylindrical cases (Figure 12). While for the shaped holes, elevated turbulence intensity has a slight influence on the cooling effectiveness, using cylindrical holes, changes the dependency between cooling performance and mass flow variation. Porter et al. [39] measured the velocity and turbulence values for the round and fan-shaped cooling holes. The survey of the velocity illustrated that intense shear and turbulence is created around the round jet. In addition, an entrainment main flow is encouraged by slight pressure area.

Fig. 11. Laterally averaged effectiveness

Fig. 12. Area-averaged film-cooling effectiveness for all cases

Using a large-scale low-speed wind tunnel, Zhu et al.

[40] studied the effects of cooling holes shapes, secondary injection Reynolds number and blowing ratios on heat transfer. The test section included five different cooling holes as dustpan-shaped holes, fan-shaped holes and round holes. The results declared that the critical blowing ratio was 1.30 for the fan and dust pan shaped holes, while for the round holes this ratio was 0.75.

Mhetras et al. [41] and Gao et al. [42], [43] studied the film cooling performance along the high pressure turbine blade by pressure sensitive paint (PSP) technique.

While, Mhetras showed that the better film cooling effectiveness is achieved on the pressure surface at higher blowing ratios, Gao et al. [42], [43] and Colban et al. [35] expanded this statement for both suction and pressure surfaces. Furthermore, they said “Secondary flow vortices such as the passage and the tip vortex significantly impact the suction side film-cooling effectiveness distribution with little or no coolant coverage in the impacted region“. In concurred with and Hong-Wook Lee et al. [44], the findings declared that higher effectiveness is achieved with blowing ratio increase on the pressure side. Vakil and Thole [1] and Barringer et al. [45] presented experimental results of the combustor simulator. In this study, a real large scale of combustor was simulated and the coolant flow and high momentum dilution jets were spread into the main flow (Fig. 13). The results indicated that high temperature gradient was developed upstream of the dilution holes.

The injection of the flow from the first row of dilution holes lead to the combustor temperature decrease by 25%. The results indicated that while, the dilution jets declined the total pressure and velocity fields, the turbulence level at the end of combustor reached to 24%.

This quantity is under predicted compared to Colban et al. [46] findings which defined the turbulence level between 25 to 30 percent. Kianpour et al in 2012 [47], simulated this model again. They used k-ε and RNG k-ε turbulent models to solve the Navier Stokes equation.

The results declared that to detect the variation of temperature, RNG k-ε turbulent model was a good choice. And this is confirmed the Colban et al. [46] and Patil et al. [48]. They mentioned that best consistency is seen between experimental and numerical findings by using the RNG k- turbulent model.

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Fig. 13. Schematic view of the combustor

The study was a computational and experimental one aiming at investigation the effects of blowing ratios on heat transfer coefficient and the film cooling performance while two rows of rectangular cooling holes arrangements were used. Koc et al. [49] showed that the film cooling effectiveness is influenced by the blowing ratio and the temperature of the penetrating flow. The best film cooling effectiveness is achieved at the mainstream and lateral direction blowing ratio of M=0.50. The temperature difference between the coolant and main stream flow increased the cooling for both lateral direction and main flow.

Finally, it is shown that heat transfer coefficient is affected at low blowing ratios.

These findings are confirmed by the results of other former studies as Koc [50].

Kassab et al. [51] simulated the heat transfer increase within one three dimensional film cooled turbine blade.

In this computational study, they used a couple’s boundary FVM method. An investigation of metal temperature differences of two distinct models declared that these differences have intense effect on metal design.

In order to specify the net heat flux reduction, Harrison et al. [52] studied the effects of film cooling and heat transfer coefficient around the trenched axial hole over the suction surface of turbine vane. The results showed that at low blowing ratios both cylindrical and trenched holes had similar behavior, however, increasing blowing ratio led to cooling performance dramatic reduction for cylindrical holes and cooling effectiveness enhancement for trenched holes and it is concurred with Shupping [53] and Baheri Islami and Jurban [54]

findings. Baheri Islami et al. studied the film cooling performance over the symmetrical turbine vane.

Ligrani et al. [55], prepared a computational and experimental study of film cooling performance of a flat plate under different blowing ratios. They said “blowing ratio has a significant effect on effectiveness results, with local and spatially-averaged effectiveness often increasing as the blowing ratio becomes larger”. On the other hand, Harrington et al. [56] showed that the blowing ratio variation affected the spatially averaged effectiveness only about 14% and it has no intense influence.

Gratton [57] conducted an experimental and computational study to investigate the influence of a contoured end wall on heat transfer over a real size turbine stator vane. Noticeable heat transfer increase is seen at great turbulence levels on the both spans. The influence of turbulence intensity was more than contoured end wall effects.

With turbulence intensity enhancement, the boundary layer transferred toward the upstream position. In 2010, Barigozzi et al. [58] studied the effects of contoured end wall for the shaped holes. By 40% modified adiabatic film cooling performance increase is illustrated by shaping the cooling holes. A protection modification is yielded at the end wall surfaces for the extended area holes whereas thermodynamic secondary loss is in the minimal value.

III. Conclusion

Rising the turbine inlet temperature is the key to higher engine efficiency in aero-engine turbine. But such hot flows cause non-uniformities at the end of the combustor and the inlet of the turbine and damage the critical parts.

Film cooling is the most well-known method of preservation. However, many researchers concentrated on the effects of dynamic and thermodynamic parameters variations on the film cooling performance and heat transfer as well as the interaction of coolant and mainstream flow at the exit of combustor and inlet of turbine among 21^th century. Sarkar and Bose [9] showed that blowing ratio increase thinned the thermal boundary layer and as a result raised the cooling performance. In line with this research, Aga, Rose and Abhari [28]

indicated that with thermal boundary layer reduction, heat transfer increase was yielded. Baheri Islami and Jurban [54] showed this for the shaped holes as well, Although, it was denied the Bernsdorf, Rose and Abhari [12] which stated that blowing ratio changes has no effect on boundary layer expect a little variation which is seen at high blowing ratios.

Of course, Rozati and Danesh Tafti [17] studied this issue experimentally and numerically for the cylindrical cooling holes. While, the test results was in concurred with Sarkar and Bose [9] and Baheri Islami and Jurban [54] findings, the numerical results approved the Harrison, Dorrington, Bees, Bogard and Bunker [52]

who declared that blowing ration increase decline the cooling effectiveness for the cylindrical cooling holes.

Furthermore, Sarkar and Bose [9] and Saumweber, Schulz and Wittig [29] showed that for cylindrical cooling holes, cooling effectiveness reduction happened as a result of turbulent reduction. However, Saumweber, Schulz and Wittig [39] declared that at high turbulence level, blowing ratio affected the cooling performance for the shaped holes not cylindrical holes. In line with this study, Harrison, Dorrington, Bees, Bogard and Bunker [52] and Shupping [53] stated this statement for the trenched cooling holes.

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However, the findings of Baldauf, Schulz and Wittig [18] showed that blowing ration augmentation led to heat transfer reduction and rejected the Aga, Rose and Abhari [28] descriptions. At last, Koc, Islamoglu and Akdag [49]

declared that temperature difference between coolant and mainstream led to cooling performance enhancement.

Acknowledgements

This work was supported by Universiti Teknologi Malaysia and Ministry of Higher Education of Malaysia.

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Authors’ information

Nor Azwadi C. Sidik was born in Kelantan on 23^rd September 1977. He received Ph.D degree (2010) from Keio University, Japan His current interest includes computational fluid dynamics, numerical methods and fluid structure interaction. Dr Nor Azwadi is a senior lecturer at Department of Thermofluid, Universiti Teknologi Malaysia, Malaysia.

Ehsan Kianpour was born in Iran on 22nd November 1980. He received Master degree (2008) from Malek Ashtar University of Technology, Iran. His current interest includes numerical heat transfer, combustion and thermodynamics. Mr Ehsan is a PhD candidate at the Department of Thermofluid, Universiti Teknologi Malaysia, Malaysia.

Iman Golshokouh was born in Iran on 4^th June 1981. He received his M.Sc. degree from Department of Mechanical Engineering from Khuzestan Azad University, Iran. Mr. Iman is a Ph.D. candidate in the Faculty of Mechanical Engineering, Universiti Teknologi Malaysia (UTM), Malaysia.