A Numerical Investigation of Air/Mist Cooling Through a Conjugate, Rotating 3-D Gas Turbine Blade With Internal, External, and Tip Cooling

2019 ◽  
Author(s):  
Ramy Abdelmaksoud ◽  
Ting Wang
Keyword(s):  
Numerical Investigation ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Outer Surface ◽  
Film Cooling ◽  
Internal Cooling ◽  
Blade Tip ◽  
Mist Cooling ◽  
Film Cooling Holes ◽  
Cooling Air

Abstract This paper describes a numerical investigation to study the effect of injecting mist (tiny water droplets) into the cooling air used to cool down rotating gas turbine blades. In this study, the conjugate heat transfer method is employed which consists of the simulation of the air/mist fluid flow inside and outside the blades as well as the heat conduction through the blade body. The complete 3-D blade with internal cooling passages and external film cooling holes on the surface and blade tip is simulated in a rotating, periodic sector of the blade. The discrete phase model (DPM) is used to simulate and track the evaporation and movement of the tiny water droplets. The rotation effect of the turbine blade is included in the CFD simulation by using the moving reference frame method. The effects of different parameters such as the mist/air ratio (10–20%) and the mist droplets size (20–40μm) on mist cooling enhancement are investigated. The results show that the mist cooling enhancements are about 10% to 25% on the outer surface of the blade and reach 50% in some locations inside the blade on the internal cooling passages walls. Most of the liquid droplets completely evaporate inside the internal cooling passages; only a limited amount of mist is able to escape from the film cooling holes to enhance the blade outer surface and blade tip cooling. The effect of 10% mist on enhanced cooling is also converted to an equivalent of a 30% reduction in cooling air flow.

A Numerical Investigation of Air/Mist Cooling through a Conjugate, Rotating 3-D Gas Turbine Blade with Internal, External, and Tip Cooling

2020 ◽  
pp. 1-30
Author(s):  
Ramy Abdelmaksoud ◽  
Ting Wang
Keyword(s):  
Numerical Investigation ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Outer Surface ◽  
Film Cooling ◽  
Internal Cooling ◽  
Blade Tip ◽  
Mist Cooling ◽  
Film Cooling Holes ◽  
Cooling Air

Abstract This paper describes a numerical investigation to study the effect of injecting mist into the cooling air used to cool down rotating gas turbine blades. The conjugate heat transfer method is employed which consists of the simulation of the air/mist fluid flow inside and outside the blades as well as the heat conduction through the blade body. The complete 3-D blade with internal cooling passages and external film cooling holes on the surface and blade tip is simulated in a rotating, periodic sector of the blade. The discrete phase model (DPM) is used to simulate and track the evaporation and movement of the tiny water droplets. The rotation effect of the turbine blade is included in the simulation by using the moving reference frame method. The effects of different parameters such as the mist/air ratio (10-20%) and the mist droplets size (20-40 µm) on mist cooling enhancement are investigated. The results show that the mist cooling enhancements are about 10% to 25% on the outer surface of the blade and reach 50% in some locations inside the blade on the internal cooling passages walls. Most of the liquid droplets completely evaporate inside the internal cooling passages; only a limited amount of mist is able to escape from the film cooling holes to enhance the blade outer surface and blade tip cooling. The effect of 10% mist on enhanced cooling is also converted to an equivalent of a 30% reduction in cooling air flow.

A Numerical Investigation of Air/Mist Cooling in a Conjugate, 3-D Gas Turbine Vane With Internal Passage and External Film Cooling

2019 ◽  
Author(s):  
Ramy Abdelmaksoud ◽  
Ting Wang
Keyword(s):  
Numerical Investigation ◽  
Size Distribution ◽  
Gas Turbine ◽  
Wall Temperature ◽  
Film Cooling ◽  
Droplet Size Distribution ◽  
Cooling Enhancement ◽  
Mist Cooling ◽  
Film Cooling Holes ◽  
Uniform Droplet

Abstract This paper describes a numerical investigation to study the effect of injecting mist (tiny water droplets, micrometers in size) into the cooling airstream to cool down gas turbine vanes. In this study, the conjugate heat transfer method is employed which consists of the simulation of the air/mist fluid flow inside and outside the vanes as well as the heat conduction through the vane body. The complete 3-D vane with internal cooling passages and external film cooling holes on the surface is simulated in a rotational periodic sector. The discrete phase model (DPM) is used to simulate and track the evaporation and movement of the tiny water droplets. The effects of different parameters such as the mist/air ratio (10–20%) and the mist droplets size (20–50μm) on mist cooling enhancement are investigated. The results show that by using a mist/air ratio of 10%, 15%, and 20% with 20 μm droplets size, on the pressure side, a maximum wall temperature reduction of 250 K, 340 K, and 450 K respectively can be achieved. On the suction side, the corresponding maximum wall temperature reductions are 160 K, 260 K, and 360 K, respectively. Using larger droplets of 50μm did not achieve better cooling enhancement because the droplets were rushed far away from the surface by the acceleration through the film cooling holes. Using the uniform droplet size distribution provides noticeably better cooling enhancement in the first 40% of the vane’s height (from the shroud) than the non-uniform droplet size distribution (Rosin-Rammler Distribution) does.

Leading Edge Shielding Concept in Gas Turbines With Can Combustors

2012 ◽  
Author(s):  
Ioanna Aslanidou ◽  
Budimir Rosic ◽  
Vasudevan Kanjirakkad ◽  
Sumiu Uchida
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Gas Turbines ◽  
Leading Edge ◽  
High Heat ◽  
Gas Temperature ◽  
Cooling Strategies ◽  
Film Cooling Holes ◽  
Cooling Air ◽  
Heat Loads

The remarkable developments in gas turbine materials and cooling technologies have allowed a steady increase in combustor outlet temperature and hence in gas turbine efficiency over the last half century. However, the efficiency benefits of higher gas temperature, even at the current levels, are significantly offset by the increased losses associated with the required cooling. Additionally, the advancements in gas turbine cooling technology have introduced considerable complexities into turbine design and manufacture. Therefore, a reduction in coolant requirements for the current gas temperature levels is one possible way for gas turbine designers to achieve even higher efficiency levels. The leading edges of the first turbine vane row are exposed to high heat loads. The high coolant requirements and geometry constraints limit the possible arrangement of the multiple rows of film cooling holes in the so called showerhead region. In the past, investigators have tested many different showerhead configurations, varying the number of rows, inclination angle and shape of the cooling holes. However the current leading edge cooling strategies using showerheads have not been shown to allow further increase in turbine temperature without excessive use of coolant air. Therefore new cooling strategies for the first vane have to be explored. In gas turbines with multiple combustor chambers around the annulus, the transition duct walls can be used to shield, i.e. to protect the first vane leading edges from the high heat loads. In this way the stagnation region at the leading edge and the shower-head of film cooling holes can be completely removed, resulting in a significant reduction in the total amount of cooling air that is otherwise required. By eliminating the showerhead the shielding concept significantly simplifies the design and lowers the manufacturing costs. This paper numerically analyses the potential of the leading edge shielding concept for cooling air reduction. The vane shape was modified to allow for the implementation of the concept and non-restrictive relative movement between the combustor and the vane. It has been demonstrated that the coolant flow that was originally used for cooling the combustor wall trailing edge and a fraction of the coolant air used for the vane showerhead cooling can be used to effectively cool both the suction and the pressure surfaces of the vane.

scholarly journals Cooling Process of Gas Turbine Blade: A Comparison Study

2020 ◽  
Vol 13 (3) ◽  
pp. 215-222
Author(s):  
Akram Luaibi Ballaoot ◽  
Naseer Hamza
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Power Output ◽  
Film Cooling ◽  
Aviation Industry ◽  
Inlet Temperature ◽  
Internal Cooling ◽  
Gas Turbine Blade ◽  
Turbine Inlet Temperature ◽  
Turbine Inlet

The gas turbine engines are occupied an important sector in the energy production and aviation industry and this important increase day after day for their features. One of the most important parameters that limit the gas turbine engine power output is the turbine inlet temperature. The higher is the turbine inlet temperature, the higher is the power output or thrust but this increases of risks of blade thermal failure due to metallurgical limits. Thus the need for a good and efficient process of blade cooling can lead to the best compromise between a powerful engine and safe operation. There are two major methods: film or external cooling and internal cooling inside the blade itself. . In the past number of years there has been considerable progress in turbine cooling research and this paper is limited to review a few selected publications to reflect recent development in turbine blade film cooling. The maximum drop in the surface temperature of the gas turbine blade and associated thermal stress – due to incorporating cooling systems- were 735 ˚C, 1217 N/mm2 respectively.

High Coverage Blade Tip Film Cooling

2004 ◽  
Author(s):  
M. Kuwabara ◽  
Keizo Tsukagoshi ◽  
T. Arts
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Gas Turbines ◽  
Separation Bubble ◽  
Density Ratio ◽  
Tip Clearance ◽  
Experimental Program ◽  
Film Cooling Effectiveness ◽  
Blade Tip

More sophisticated cooling schemes are required for the turbine blade due to the demand of increased turbine temperature for improved performance. Although the tip portion of a turbine blade is one of the most critical portions in a gas turbine, there are few studies on cooling this portion compared to those for airfoil, especially film cooling strategies. Industrial gas turbines have a more uniform gas temperature profile than aero engines. For these applications, it is more important to understand the characteristics of tip film cooling to improve the blade durability and gas turbine performance by reducing cooling air. A numerical and experimental program was initiated to study film cooling effectiveness on a flat blade tip as a function of tip gap and mass flux ratios. Flow visualization tests were conducted with and without film cooling to verify the numerical CFD findings. The predictions and visualization results showed that a separation bubble forms at the pressure side edge that increases with tip gap. Film effectiveness measurements were carried out on a 1.3X scale blade model in a low speed test while simulating the normalized pressure distribution typical of an engine design. The engine density ratio of the coolant to mainstream was replicated in the film cooling tests to provide the best simulation of the engine. Two rows of holes were placed near the tip of the blade to provide high film coverage prior to the flowing over the tip. The data shows that film effectiveness increases with decreasing tip clearance. Blowing ratio provides an improvement due to the added mass flow, which was shown by a non-dimensionalized correlation.

Characteristics of Tip Leakage Flow of the Turbine Blade With Cutback Squealer and Coolant Injection

2010 ◽  
Author(s):  
Jianhua Wang ◽  
Yalin Liu ◽  
Xiaochun Wang ◽  
Zhineng Du ◽  
Shijie Yang
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Flow Characteristics ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
Coolant Injection ◽  
Blade Tip ◽  
Film Cooling Holes ◽  
Numerical Program

Experimental and numerical investigations of the tip leakage flow characteristics between turbine blade tip and stator wall (shroud) were conducted by a particle image velocimetry (PIV) system and the commercially available software CFX 11.0. A three-time scaled profile of the GE-E3 blade was used as specimen. Two rows of cylindrical film-cooling holes with 1.5mm diameter were arranged in the blade tip. One row with 5 holes was placed in pressure side just below the groove floor, and the other with 11 holes was equidistantly arranged on the tip along the mid camber line. To exhibit the generation and movement of leakage vortex, and to compare the coolant injection effects from different rows, several typical velocity profiles were captured by the PIV system. The experimental results were used as a data source to validate the turbulence model and numerical program. To better understand the mixing characteristics of the coolant injected from different rows with the leakage flow, the fluid fields of the leakage vortex and coolant flow were simulated, and the leakage mass rates from the blade tip in different coolant injection cases and different gaps were quantitatively estimated by the validated numerical program.

Influence of Coolant on Cooling Performance Sensitivity of Internally Convective Turbine Vane

2021 ◽  
Author(s):  
Thanapat Chotroongruang ◽  
Prasert Prapamonthon ◽  
Rungsimun Thongdee ◽  
Thanapat Thongmuenwaiyathon ◽  
Zhenxu Sun ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Internal Cooling ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Turbine Engine ◽  
Performance Sensitivity ◽  
Turbine Vane ◽  
Cooling Air

Abstract Based on the Brayton cycle for gas-turbine engines, the high thermal efficiency and power output of a gas-turbine engine can be obtainable when the gas-turbine engine operates at high turbine inlet temperatures. However, turbine components e.g., inlet guide vane, rotor blade, and stator vane request high cooling performance. Typically, internal cooling and film cooling are two effective techniques that are widely used to protect high thermal loads for the turbine components in a state-of-the-art gas turbine. Consequently, the high thermal efficiency and power output can be obtained, and the turbine lifespan can be prolonged, also. On top of that, a comprehensive understanding of flow and heat transfer phenomena in the turbine components is very important. As a result, both experiments and simulations have been used to improve the cooling performance of the turbine components. In fact, the cooling air used in the internal cooling and film cooling is partially extracted from the compressor. Therefore, variations in the cooling air affect the cooling performance of the turbine components directly. This paper presents a numerical study on the influence of the cooling air on cooling-performance sensitivity of an internally convective turbine vane, MARK II using the computational fluid dynamics (CFD)/conjugate heat transfer (CHT) with the SST k-ω turbulence model. Result comparisons are conducted in terms of pressure, temperature, and cooling effectiveness under the effects of the inlet temperature, mass flow rate, turbulence intensity, and flow direction of the cooling air. The cooling-performance sensitivity to the coolant parameters is shown through variations of local cooling effectiveness, and area and volume-weighted average cooling effectiveness.

Effect of Rotation on a Gas Turbine Blade Internal Cooling System: Numerical Investigation

2016 ◽  
Author(s):  
E. Burberi ◽  
D. Massini ◽  
L. Cocchi ◽  
L. Mazzei ◽  
A. Andreini ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Flow Field ◽  
Transfer Coefficient ◽  
Numerical Investigation ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Cooling System ◽  
Leading Edge ◽  
Internal Cooling

Increasing turbine inlet temperature is one of the main strategies used to accomplish the demands of increased performance of modern gas turbines. As a consequence, optimization of the cooling system is of paramount importance in gas turbine development. Leading edge represents a critical part of cooled nozzles and blades, given the presence of the hot gases stagnation point and the unfavourable geometry for cooling. This paper reports the results of a numerical investigation aimed at assessing the rotation effects on the heat transfer distribution in a realistic leading edge internal cooling system of a high pressure gas turbine blade. The numerical investigation was carried out in order to support and to allow an in-depth understanding of the results obtained in a parallel experimental campaign. The model is composed of a trapezoidal feeding channel which provides air to the cold bridge system by means of three large racetrack-shaped holes, generating coolant impingement on the internal concave leading edge surface, whereas four big fins assure the jets confinement. Air is then extracted through 4 rows of 6 holes reproducing the external cooling system composed of shower-head and film cooling holes. Experiments were performed in static and rotating conditions replicating the typical range of jet Reynolds number (Rej) from 10000 to 40000 and Rotation number (Roj) up to 0.05, for three crossflow cases representative of the working condition that can be found at blade tip, midspan and hub, respectively. Experimental results in terms of flow field measurements on several internal planes and heat transfer coefficient on the LE internal surface have been performed on two analogous experimental campaigns at University of Udine and University of Florence respectively. Hybrid RANS-LES models were used for the simulations, such as Scale Adaptive Simulation (SAS) and Detached Eddy Simulation (DES), given their ability to resolve the complex flow field associated with jet impingement. Numerical flow field results are reported in terms of both jet velocity profiles and 2D vector plots on symmetry and transversal internal planes, while the heat transfer coefficient distributions are presented as detailed 2D maps together with radial and tangential averaged Nusselt number profiles. A fairly good agreement with experimental measurements is observed, which represent a validation of the adopted computational model. As a consequence, the computed aerodynamic and thermal fields also allow an in-depth interpretation of the experimental results.

Leading Edge Shielding Concept in Gas Turbines With Can Combustors

2012 ◽  
Vol 135 (2) ◽  
Author(s):  
Ioanna Aslanidou ◽  
Budimir Rosic ◽  
Vasudevan Kanjirakkad ◽  
Sumiu Uchida
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Gas Turbines ◽  
Leading Edge ◽  
High Heat ◽  
Gas Temperature ◽  
Cooling Strategies ◽  
Film Cooling Holes ◽  
Cooling Air ◽  
Heat Loads

The remarkable developments in gas turbine materials and cooling technologies have allowed a steady increase in combustor outlet temperature and, hence, in gas turbine efficiency over the last half century. However, the efficiency benefits of higher gas temperature, even at the current levels, are significantly offset by the increased losses associated with the required cooling. Additionally, the advancements in gas turbine cooling technology have introduced considerable complexities into turbine design and manufacture. Therefore, a reduction in coolant requirements for the current gas temperature levels is one possible way for gas turbine designers to achieve even higher efficiency levels. The leading edges of the first turbine vane row are exposed to high heat loads. The high coolant requirements and geometry constraints limit the possible arrangement of the multiple rows of film cooling holes in the so-called showerhead region. In the past, investigators have tested many different showerhead configurations by varying the number of rows, inclination angle, and shape of the cooling holes. However, the current leading edge cooling strategies using showerheads have not been shown to allow a further increase in turbine temperature without the excessive use of coolant air. Therefore, new cooling strategies for the first vane have to be explored. In gas turbines with multiple combustor chambers around the annulus, the transition duct walls can be used to shield, i.e., to protect, the first vane leading edges from the high heat loads. In this way, the stagnation region at the leading edge and the showerhead of film cooling holes can be completely removed, resulting in a significant reduction in the total amount of cooling air that is otherwise required. By eliminating the showerhead the shielding concept significantly simplifies the design and lowers the manufacturing costs. This paper numerically analyzes the potential of the leading edge shielding concept for cooling air reduction. The vane shape was modified to allow for the implementation of the concept and nonrestrictive relative movement between the combustor and the vane. It has been demonstrated that the coolant flow that was originally used for cooling the combustor wall trailing edge and a fraction of the coolant air used for the vane showerhead cooling can be used to effectively cool both the suction and the pressure surfaces of the vane.

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