Film Cooling of Turbine Blade Surface With Extended Exit Holes

2014 ◽  
Author(s):  
Fariborz Forghan ◽  
Omid Askari ◽  
Uichiro Narusawa ◽  
Hameed Metghalchi
Keyword(s):  
High Temperature ◽  
Turbine Blade ◽  
Film Cooling ◽  
Gas Flow ◽  
Three Dimensional ◽  
Suction Side ◽  
Ideal Gas ◽  
Coolant Flow ◽  
Blade Surface ◽  
Flow Through

The main goal of gas turbine design is the effective use of energy. Usually, the efficient high temperature first and second stage turbine blade surface is cooled by jet of coolant flow from extended exit holes (EEH). Against the prevailing hot gas flow, the flow through EEH must be designed to form a film of cool air over the blade. Computational analyses are performed to examine the cooling effectiveness of flow from EEH over the suction side of a blade by solving conservation equations (mass, momentum and energy) and the ideal gas equation of state for the three-dimensional, turbulent, compressible flow. A diverging flow through EEH is typically choked at its throat, resulting in a supersonic flow, a shock and then a subsonic flow downstream. The location of the shock relative to the high-temperature gas flow over the blade determines the temperature distribution along the blade surface; which is analyzed in detail when the coolant flow rate is varied.

Cooling of Turbine Blade Surface With Expanded Exit Holes: Computational Suction-Side Analysis

2016 ◽  
Vol 138 (5) ◽  
Author(s):  
Fariborz Forghan ◽  
Omid Askari ◽  
Uichiro Narusawa ◽  
Hameed Metghalchi
Keyword(s):  
Mass Flow Rate ◽  
Turbine Blade ◽  
Flow Rate ◽  
Temperature Difference ◽  
Film Cooling ◽  
Gas Flow ◽  
Suction Side ◽  
Blade Surface ◽  
Cooling Effectiveness ◽  
Flow Through

Turbine blade surfaces are cooled by jet flow from expanded exit holes (EEHs) against the prevailing hot gas flow. The flow through EEH must be designed to form a film of cool air over the blade. Computational analyses are performed to examine the cooling effectiveness of flow from EEH over the suction side of a blade by solving conservation equations and the ideal gas equation of state for turbulent and compressible flow. For a sufficiently high coolant mass flow rate, the flow through EEH, which acts as a converging–diverging nozzle, is choked at the nozzle throat, resulting in a supersonic flow, a shock, and then a subsonic flow downstream. The location of the shock relative to the high-temperature gas flow determines the temperature distribution along the blade surface; which is analyzed in detail when the following conditions are varied: coolant mass flow rate, the temperature difference between gas-and coolant-flow, EEH location on the blade surface, EEH inclination angle to the blade surface, and exit-to-inlet area ratio (AR) of EEH. The film cooling effectiveness is calculated along the surface of the blade. The results show (1) increasing the coolant flow rate improves the effectiveness, (2) change in temperature difference between the mainstream and the coolant slightly affects the effectiveness, (3) inclination angle of EEH has a pronounced effect on film cooling and the corresponding effectiveness, (4) both the location of the EEH on a blade and the AR of the EEH slightly change the effectiveness.

Calculations of Cooled Turbine Efficiency

2008 ◽  
Vol 130 (1) ◽  
Author(s):  
J. H. Horlock ◽  
Leonardo Torbidoni
Keyword(s):  
Film Cooling ◽  
Gas Flow ◽  
Coolant Flow ◽  
The Other ◽  
Turbine Stage ◽  
Isentropic Expansion ◽  
Turbine Efficiency ◽  
Flow Through ◽  
Definition Of ◽  
Industrial Gas Turbine

The efficiency of a cooled turbine stage has been discussed in the literature. All proposed definitions compare the actual power output with an ideal output, which has to be determined; but usually, one of two definitions has been used by turbine designers. In the first, the so-called Hartsel efficiency, the mainstream gas flow, and the various coolant flows to rotor and stator are assumed to expand separately and isentropically to the backpressure. In the second, it is assumed that these flows mix at constant (mainstream) gas pressure before expanding isentropically (sometimes, the rotor coolant flow is ignored in this definition). More recently, it has been suggested that a thermodynamically sounder definition is one in which the gas and coolant flows mix reversibly and adiabatically before isentropic expansion to the backpressure. In the current paper, these three efficiencies are compared, for a typical stage—the first cooled stage of a multistage industrial gas turbine. It is shown that all the efficiencies fall more or less linearly with increase of the fractional (total) coolant flow. It is also shown that the new definition of efficiency gives values considerably lower than the other two efficiencies, which are more widely used at present. Finally, the various irreversibilities associated with the flow through a cooled turbine are calculated. Although all these irreversibilities increase with the fractional coolant flow, it is shown that the “thermal” irreversibility associated with film cooling is higher than the other irreversibilities at large fractional coolant flow.

Film-Cooling on a Gas Turbine Blade Pressure Side or Suction Side With Compound Angle Shaped Holes

2008 ◽  
Vol 131 (1) ◽  
Author(s):  
Zhihong Gao ◽  
Diganta P. Narzary ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Pressure Ratio ◽  
Suction Side ◽  
Pressure Side ◽  
Blade Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Surface ◽  
Compound Angle

The film-cooling effectiveness on the surface of a high pressure turbine blade is measured using the pressure sensitive paint technique. Compound angle laidback fan-shaped holes are used to cool the blade surface with four rows on the pressure side and two rows on the suction side. The coolant injects to one side of the blade, either pressure side or suction side. The presence of wake due to the upstream vanes is simulated by placing a periodic set of rods upstream of the test blade. The wake rods can be clocked by changing their stationary positions to simulate progressing wakes. The effect of wakes is recorded at four phase locations along the pitchwise direction. The freestream Reynolds number, based on the axial chord length and the exit velocity, is 750,000. The inlet and exit Mach numbers are 0.27 and 0.44, respectively, resulting in a pressure ratio of 1.14. Five average blowing ratios ranging from 0.4 to 1.5 are tested. Results reveal that the tip-leakage vortices and endwall vortices sweep the coolant on the suction side to the midspan region. The compound angle laidback fan-shaped holes produce a good film coverage on the suction side except for the regions affected by the secondary vortices. Due to the concave surface, the coolant trace is short and the effectiveness level is low on the pressure surface. However, the pressure side acquires a relatively uniform film coverage with the multiple rows of cooling holes. The film-cooling effectiveness increases with the increasing average blowing ratio for either side of coolant ejection. The presence of stationary upstream wake results in lower film-cooling effectiveness on the blade surface. The compound angle shaped holes outperform the compound angle cylindrical holes by the elevated film-cooling effectiveness, particularly at higher blowing ratios.

Predictions of Cooling From Dirt Purge Holes Along the Tip of a Turbine Blade

2003 ◽  
Author(s):  
E. M. Hohlfeld ◽  
J. R. Christophel ◽  
E. L. Couch ◽  
K. A. Thole
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Film Cooling ◽  
Suction Side ◽  
High Heat ◽  
Blowing Ratio ◽  
High Heat Transfer ◽  
Blade Tip ◽  
Flow Through ◽  
Film Cooling Holes

The clearance gap between the tip of a turbine blade and its associated shroud provides a flow path for leakage from the pressure side of the blade to the suction side. The tip region is one area that experiences high heat transfer and, as such, can be the determining factor for blade life. One method for reducing blade tip heat transfer is to use cooler fluid from the compressor, that exits from relatively large dirt purge holes placed in the tip, for cooling purposes. Dirt purge holes are typically manufactured in the blade tip to extract dirt from the coolant flow through centrifugal forces such that these dirt particles do not block smaller diameter film-cooling holes. This paper discusses the results of numerous computational simulations of cooling injection from dirt purge holes along the tip of a turbine blade. Some comparisons are also made to experimental results in which a properly scaled-up blade geometry (12X) was used to form a two-passage linear cascade. Computational results indicate that the cooling achieved through the dirt purge injection from the blade tip is dependent on the gap size as well as the blowing ratio. For a small tip gap (0.54% of the span) the flow exiting the dirt purge holes act as a blockage for the leakage flow across the gap. As the blowing ratio is increased for a large tip gap (1.63% of the span), the tip cooling increases only slightly while the cooling to the shroud increases significantly.

Calculations of Cooled Turbine Efficiency

2006 ◽  
Author(s):  
J. H. Horlock ◽  
Leonardo Torbidoni
Keyword(s):  
Film Cooling ◽  
Gas Flow ◽  
Back Pressure ◽  
Coolant Flow ◽  
The Other ◽  
Turbine Stage ◽  
Turbine Efficiency ◽  
Flow Through ◽  
Definition Of ◽  
Industrial Gas Turbine

The efficiency of a cooled turbine stage has been discussed in the literature. All proposed definitions compare the actual power output with an ideal output, which has to be determined; but usually one of two definitions has been used by turbine designers. In the first, the so-called Hartsel efficiency, the mainstream gas flow and the various coolant flows to rotor and stator are assumed to expand separately and isentropically to the back pressure. In the second it is assumed that these flows mix at constant (mainstream) gas pressure before expanding isentropically (sometimes the rotor coolant flow is ignored in this definition). More recently it has been suggested that a thermodynamically sounder definition is one in which the gas and coolant flows mix reversibly and adiabatically before isentropic expansion to the back pressure. In the current paper these three efficiencies are compared, for a typical stage — the first cooled stage of a multistage industrial gas turbine. It is shown that all the efficiencies fall more or less linearly with increase of the fractional (total) coolant flow. It is also shown that the new definition of efficiency gives values considerably lower than the other two efficiencies, which are more widely used at present. Finally, the various irreversibilities associated with the flow through a cooled turbine are calculated. Although all these irreversibilities increase with the fractional coolant flow, it is shown that the “thermal” irreversibility associated with film cooling is higher than the other irreversibilities at large fractional coolant flow.

Clocking Effects of Inlet Non-Uniformities in a Fully Cooled High-Pressure Vane: A Conjugate Heat Transfer Analysis

2015 ◽  
Author(s):  
Duccio Griffini ◽  
Massimiliano Insinna ◽  
Simone Salvadori ◽  
Francesco Martelli
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Cooling System ◽  
Hot Spot ◽  
Three Dimensional ◽  
Leading Edge ◽  
Suction Side ◽  
Heat Transfer Analysis

A high-pressure vane equipped with a realistic film-cooling configuration has been studied. The vane is characterized by the presence of multiple rows of fan-shaped holes along pressure and suction side while the leading edge is protected by a showerhead system of cylindrical holes. Steady three-dimensional Reynolds-Averaged Navier-Stokes (RANS) simulations have been performed. A preliminary grid sensitivity analysis with uniform inlet flow has been used to quantify the effect of spatial discretization. Turbulence model has been assessed in comparison with available experimental data. The effects of the relative alignment between combustion chamber and high-pressure vanes are then investigated considering realistic inflow conditions in terms of hot spot and swirl. The inlet profiles used are derived from the EU-funded project TATEF2. Two different clocking positions are considered: the first one where hot spot and swirl core are aligned with passage and the second one where they are aligned with the leading edge. Comparisons between metal temperature distributions obtained from conjugate heat transfer simulations are performed evidencing the role of swirl in determining both the hot streak trajectory within the passage and the coolant redistribution. The leading edge aligned configuration is resulted to be the most problematic in terms of thermal load, leading to increased average and local vane temperature peaks on both suction side and pressure side with respect to the passage aligned case. A strong sensitivity of both injected coolant mass flow and heat removed by heat sink effect has also been highlighted for the showerhead cooling system.

Turbine Blade Platform Film Cooling With Simulated Swirl Purge Flow and Slashface Leakage Conditions

2016 ◽  
Author(s):  
Andrew F. Chen ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
Leakage Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Purge Flow ◽  
Endwall Film Cooling ◽  
Film Cooling Holes

The combined effects of inlet purge flow and the slashface leakage flow on the film cooling effectiveness of a turbine blade platform were studied using the pressure sensitive paint (PSP) technique. Detailed film cooling effectiveness distributions on the endwall were obtained and analyzed. The inlet purge flow was generated by a row of equally-spaced cylindrical injection holes inside a single-tooth generic stator-rotor seal. In addition to the traditional 90 degree (radial outward) injection for the inlet purge flow, injection at a 45 degree angle was adopted to create a circumferential/azimuthal velocity component toward the suction side of the blades, which created a swirl ratio (SR) of 0.6. Discrete cylindrical film cooling holes were arranged to achieve an improved coverage on the endwall. Backward injection was attempted by placing backward injection holes near the pressure side leading edge portion. Slashface leakage flow was simulated by equally-spaced cylindrical injection holes inside a slot. Experiments were done in a five-blade linear cascade with an average turbulence intensity of 10.5%. The inlet and exit Mach numbers were 0.26 and 0.43, respectively. The inlet and exit mainstream Reynolds numbers based on the axial chord length of the blade were 475,000 and 720,000, respectively. The coolant-to-mainstream mass flow ratios (MFR) were varied from 0.5%, 0.75%, to 1% for the inlet purge flow. For the endwall film cooling holes and slashface leakage flow, blowing ratios (M) of 0.5, 1.0, and 1.5 were examined. Coolant-to-mainstream density ratios (DR) that range from 1.0 (close to low temperature experiments) to 1.5 (intermediate DR) and 2.0 (close to engine conditions) were also examined. The results provide the gas turbine engine designers a better insight into improved film cooling hole configurations as well as various parametric effects on endwall film cooling when the inlet (swirl) purge flow and slashface leakage flow were incorporated.

Cooling Effectiveness on Film Cooled Turbine Blades

2001 ◽  
Author(s):  
M. Derrar ◽  
J. Nagler ◽  
W. W. Koschel
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Separation Bubble ◽  
Turbine Blades ◽  
Suction Side ◽  
Experimental Simulation ◽  
Flow Conditions ◽  
Cooling Effectiveness ◽  
Boundary Interaction ◽  
Simultaneous Modeling

Abstract This paper presents experiments on the cooling effectiveness obtained for two different injection locations on the suction side of a turbine blade at transonic flow conditions. Previous results of a computational analysis and flow visualization indicated that a separation bubble is present on the suction side at a location x/L = 0.43 and the location x/L = 0.575 corresponds to a shock-boundary interaction zone [9]. The scientific interest is primarily focused on the realization of high film cooling efficiencies and its relevant parameters under these flow conditions. Streamwise aligned as well as inclined angled film coolant hole configurations have been investigated for each location. Due to the high number of interacting parameters the experimental simulation of turbine blade film cooling is extremely complex, which can only be solved by a simultaneous modeling using the experimentally measured results. Test rig, instrumentation and data analysis are described in detail. The goal of the investigations is to determine the optimum location of the film coolant injection.

Numerical Simulation of a Particle in Air Flow Around a Turbine Blade

2020 ◽  
Author(s):  
Ippei Oshima ◽  
Mikito Furuichi
Keyword(s):  
Numerical Simulation ◽  
Theoretical Model ◽  
Turbine Blade ◽  
Gas Flow ◽  
Three Dimensional ◽  
Prediction Method ◽  
Stokes Number ◽  
Lagrangian Simulation ◽  
Background Gas ◽  
Liquid Mass Flow Rate

Abstract The Steam turbine is widely used for generating electricity, in the thermal, nuclear and geothermal power generation systems. A wet loss is known as one of the degrading factors of the performance. To reduce the amount of liquid phase generated by condensation and atomization from nozzles, the prediction of the distribution of liquid mass flow rate inside the turbine is important. However, the quantitative understanding and the prediction method of the liquid flow inside the turbine remain unclear because physics inside a turbine is consisting of complex multiscale and multiphase events. In the present study, we proposed a theoretical model predicting the motion of droplet particles in gas flow based on Stokes number whose model does not require numerical simulation. We also conducted the numerical validation test using three-dimensional Eulerian-Lagrangian simulation for the problem with turbine blade T106. The numerical simulation shows that the particle motion is characterized by the Stokes number, that is consistent with the assumption of the theoretical model and previous studies. When Stokes number is smaller than one, the particle trajectory just follows the gas flow streamline and avoids the impacts on the surface of T106. With increasing Stokes number, the particles begin to deviate from the gas flow. As a result, many particles collide with the surface of T106 when the Stokes number is approximately one. When the Stokes number is extremely larger than one, particles move straight regardless of the background gas flow. The good agreements between the theoretical predictions and numerical experiment results justify the use of our proposed theoretical model for the prediction of the particle flow around the turbine blade.

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