Browsing by Author "Jesuraj, F."

Now showing 1 - 5 of 5

Effect of RANS-Type Turbulence Models on Adiabatic Film Cooling Effectiveness over a Scaled Up Gas Turbine Blade Leading Edge Surface
(Springer, 2018) Yepuri, G.B.; Talanki Puttarangasetty, A.B.; Kolke, D.K.; Jesuraj, F.
Increasing the gas turbine inlet temperature is one of the key technologies in raising gas turbine engine power output. Film cooling is one of the efficient cooling techniques to cool the hot section components of a gas turbine engines in turn the turbine inlet temperature can be increased. This study aims at investigating the effect of RANS-type turbulence models on adiabatic film cooling effectiveness over a scaled up gas turbine blade leading edge surfaces. For the evaluation, five different two equation RANS-type turbulent models have been taken in consideration, which are available in the ANSYS-Fluent. For this analysis, the gas turbine blade leading edge configuration is generated using Solid Works. The meshing is done using ANSYS-Workbench Mesh and ANSYS-Fluent is used as a solver to solve the flow field. The considered gas turbine blade leading edge model is having five rows of film cooling circular holes, one at stagnation line and the two each on either side of stagnation line at 30° and 60° respectively. Each row has the five holes with the hole diameter of 4 mm, pitch of 21 mm arranged in staggered manner and has the hole injection angle of 30° in span wise direction. The experiments are carried in a subsonic cascade tunnel facility at heat transfer lab of CSIR-National Aerospace Laboratory with a Reynolds number of 1,00,000 based on leading edge diameter. From the Computational Fluid Dynamics (CFD) evaluation it is found that K–? Realizable model gives more acceptable results with the experimental values, compared to the other considered turbulence models for this type of geometries. Further the CFD evaluated results, using K–? Realizable model at different blowing ratios are compared with the experimental results. © 2016, The Institution of Engineers (India).
Experimental and numerical investigation of effect of blowing ratio on film cooling effectiveness and heat transfer coefficient over a gas turbine blade leading edge film cooling configurations
(2013) Yepuri, G.B.; Lalgi, G.; Puttarangasetty, A.B.T.; Jesuraj, F.; Kenkere, S.R.V.; Nanjundaiah, V.K.
Film cooling is one of the cooling techniques to cool the hot section components of a gas turbine engines. The gas turbine blade leading edges are the vital parts in the turbines as they are directly hit by the hot gases, hence the optimized cooling of gas turbine blade surfaces is essential. This study aims at investigating the film cooling effectiveness and heat transfer coefficient experimentally and numerically for the three different gas turbine blade leading edge models each having the one row of film cooling holes at 15, 30 and 45 degrees hole orientation angle respectively from stagnation line. Each row has the five holes with the hole diameter of 3mm, pitch of 20mm and has the hole inclination angle of 20deg. in spanwise direction. Experiments are carried out using the subsonic cascade tunnel facility of National Aerospace Laboratories, Bangalore at a nominal flow Reynolds number of 1,00,000 based on the leading edge diameter, varying the blowing ratios of 1.2, 1.50, 1.75 and 2.0. In addition, an attempt has been made for the film cooling effectiveness using CFD simulation, using k-? realizable turbulence model to solve the flow field. Among the considered 15, 30 and 45 deg. models, both the cooling effectiveness and heat transfer coefficient shown the increase with the increase in hole orientation angle from stagnation line. The film cooling effectiveness increases with the increase in blowing ratio upto 1.5 for the 15 and 30 deg. models, whereas on the 45 deg. model the increase in effectiveness shown upto the blowing ratio of 1.75. The heat transfer coefficient values showed the increase with the increase in blowing ratio for all the considered three models. The CFD results in the form of temperature, velocity contours and film cooling effectiveness values have shown the meaningful results with the experimental values. Copyright � 2013 by ASME.
Experimental and numerical investigation of effect of blowing ratio on film cooling effectiveness and heat transfer coefficient over a gas turbine blade leading edge film cooling configurations
(American Society of Mechanical Engineers infocentral@asme.org, 2013) Yepuri, G.B.; Lalgi, G.; Puttarangasetty, A.B.T.; Jesuraj, F.; Kenkere, S.R.V.; Vinod Kumar, V.K.
Film cooling is one of the cooling techniques to cool the hot section components of a gas turbine engines. The gas turbine blade leading edges are the vital parts in the turbines as they are directly hit by the hot gases, hence the optimized cooling of gas turbine blade surfaces is essential. This study aims at investigating the film cooling effectiveness and heat transfer coefficient experimentally and numerically for the three different gas turbine blade leading edge models each having the one row of film cooling holes at 15, 30 and 45 degrees hole orientation angle respectively from stagnation line. Each row has the five holes with the hole diameter of 3mm, pitch of 20mm and has the hole inclination angle of 20deg. in spanwise direction. Experiments are carried out using the subsonic cascade tunnel facility of National Aerospace Laboratories, Bangalore at a nominal flow Reynolds number of 1,00,000 based on the leading edge diameter, varying the blowing ratios of 1.2, 1.50, 1.75 and 2.0. In addition, an attempt has been made for the film cooling effectiveness using CFD simulation, using k-Îµ realizable turbulence model to solve the flow field. Among the considered 15, 30 and 45 deg. models, both the cooling effectiveness and heat transfer coefficient shown the increase with the increase in hole orientation angle from stagnation line. The film cooling effectiveness increases with the increase in blowing ratio upto 1.5 for the 15 and 30 deg. models, whereas on the 45 deg. model the increase in effectiveness shown upto the blowing ratio of 1.75. The heat transfer coefficient values showed the increase with the increase in blowing ratio for all the considered three models. The CFD results in the form of temperature, velocity contours and film cooling effectiveness values have shown the meaningful results with the experimental values. Copyright Â© 2013 by ASME.
Experimental investigation of overall cooling effectiveness on combustion chamber liner with and without impingement holes
(2015) Jesuraj, F.; Rajendran, R.; Narayanappa, K.G.; Yepuri, G.B.; Sasikumar, V.; Poozhiyil, S.
The gas turbine combustor liner which is subjected to high temperature requires efficient cooling. In earlier days concept of slot film cooling is utilized in the combustion liners and in modern combustors multiple row film cooling (effusion cooling) is mainly used. This study aims at the experimental investigation of overall film cooling effectiveness of an effusion plate with and without impingement holes at the backside. The experiments are done at different blowing ratios and the surface temperature measurements are taken using infrared thermography. The effusion and impingement holes are arranged in staggered manner on two parallel plates and each effusion hole is surrounded by four impingement holes. Effusion holes are drilled at an angle of 27� and the impingement plate is kept at a distance of 6D away from the effusion plate. The experiments are done on the effusion plate with and without impingement plate at the backside. The results show, increase in cooling effectiveness as the blowing ratio increases. The comparative results shows that at a particular blowing ratio the overall cooling effectiveness is higher for effusion plate with impingement holes at the backside due to the higher convective heat transfer coefficients produced by the impinging jets at the cold side of the effusion plate. � Copyright 2015 by ASME.
Experimental investigation of overall cooling effectiveness on combustion chamber liner with and without impingement holes
(American Society of Mechanical Engineers, 2015) Jesuraj, F.; Rajendran, R.; Gottekere Narayanappa, K.G.; Yepuri, G.B.; Sasikumar, V.; Poozhiyil, S.
The gas turbine combustor liner which is subjected to high temperature requires efficient cooling. In earlier days concept of slot film cooling is utilized in the combustion liners and in modern combustors multiple row film cooling (effusion cooling) is mainly used. This study aims at the experimental investigation of overall film cooling effectiveness of an effusion plate with and without impingement holes at the backside. The experiments are done at different blowing ratios and the surface temperature measurements are taken using infrared thermography. The effusion and impingement holes are arranged in staggered manner on two parallel plates and each effusion hole is surrounded by four impingement holes. Effusion holes are drilled at an angle of 27Â° and the impingement plate is kept at a distance of 6D away from the effusion plate. The experiments are done on the effusion plate with and without impingement plate at the backside. The results show, increase in cooling effectiveness as the blowing ratio increases. The comparative results shows that at a particular blowing ratio the overall cooling effectiveness is higher for effusion plate with impingement holes at the backside due to the higher convective heat transfer coefficients produced by the impinging jets at the cold side of the effusion plate. Â© Copyright 2015 by ASME.