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Experimental and Numerical Investigation of Adiabatic Film Cooling Effectiveness Over the Compound Angled

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Experimental and Numerical Investigation of Adiabatic Film Cooling Effectiveness Over the Compound Angled
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  Proceedings of the 2 nd  International Conference on Current Trends in Engineering and Management ICCTEM -2014 17 – 19, July 2014, Mysore, Karnataka, India   91 EXPERIMENTAL AND NUMERICAL INVESTIGATION OF ADIABATIC FILM COOLING EFFECTIVENESS OVER THE COMPOUND ANGLED GAS TURBINE BLADE LEADING EDGE MODEL Giridhara Babu Y. 1* , Ashok Babu T.P. 2 , Anbalagan M. 3 , Meena R. 4   1 Propulsion Division, CSIR-National Aerospace Laboratories, Bangalore, India, 2 Mechanical Engg. Dept., National Institute of Technology Karnataka, Mangalore, India 3, 4 Mechanical Engg. Dept., The Oxford College of Engineering, Bangalore, India ABSTRACT This study aims at investigating the film cooling effectiveness using both experimentally and numerically for the scaled up gas turbine blade leading edge compound angle model. A compound angle gas turbine blade leading edge model having the five rows of holes, one at stagnation line, two rows of holes at 30 degrees on either side of stagnation line and two rows of holes at 60 degrees on either side of stagnation line. Each row has the five holes at a pitch of 21mm with the varied hole angles of 0, 30, 45, 55 and 60 degrees oriented with the stream line direction. The film cooling hole rows are arranged in a staggered manner to cover the more flow area on the blade surface and with each row consisting of 5 holes with the hole diameter of 4mm. Experiments are conducted by varying the blowing ratios (B.R) in the range of 1.0 to 2.5 at the density ratio (D.R) of 1.30, at a nominal flow Reynolds number of 1, 00,000 based on the leading edge diameter. The blowing ratio is varied by varying the coolant mass flow with the main flow maintained at a constant rate. The density ratio of 1.30 is maintained by allowing the coolant flow at 231K and main stream at room temperature. The film cooling effectiveness is also found using CFD (Fluent) simulation. The k-€ realizable turbulence model is used to solve the flow field. The generated CFD results are compared with the experimental results for the validation of CFD. The CFD results indicated the similar trends of the cooling effectiveness results as that of experimental values. Both the CFD and Experimental shown the increase in cooling effectiveness increases with the increase in blowing ratio upto 2.0 and found the decrease over the 2.0 for this model hence, the optimized blowing ratio can be considered as 2.0. The generated temperature and velocity contours using CFD shown the heat loads on the leading edge surface. Keywords: Blowing ratio, CFD , Compound angle, Density ratio, Gas turbines.  1.   INTRODUCTION Generally, the Gas turbines are operated at higher temperatures in the range of 1200-1800 degree Celsius so the materials used for gas turbine blades should with stand these high temperatures without any melting and thermal stresses. 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 leading edge surfaces is essential. Film cooling is used in many applications to reduce convective heat transfer to a surface. In order to increase the life of the blade and efficiency, the optimized cooling of gas turbine blade leading edge surfaces is essential. Film cooling is used in many applications to reduce convective heat transfer to a surface. Some of the examples are, film cooling of gas turbine combustion chamber liners, vanes and blades which are subjected to high heat loads from combustion gases.   INTERNATIONAL JOURNAL OF MECHANICAL ENGINEERING AND TECHNOLOGY (IJMET) ISSN 0976 – 6340 (Print) ISSN 0976 – 6359 (Online) Volume 5, Issue 9, September (2014), pp. 91-100 © IAEME: www.iaeme.com/IJMET.asp Journal Impact Factor (2014): 7.5377 (Calculated by GISI) www.jifactor.com   IJMET   © I A E M E    Proceedings of the 2 nd  International Conference on Current Trends in Engineering and Management ICCTEM -2014 17 – 19, July 2014, Mysore, Karnataka, India   92 Film cooling is the introduction of a secondary fluid at one or more discrete locations along a surface exposed to a high temperature environment to protect the surface not only in the immediate region of injection but also in the downstream region as shown in Fig.1. Turbine airfoil surfaces, shrouds, blades tips and end walls are cooled using discrete-hole film cooling. Film cooling protects the airfoil surface directly, compared to internal cooling technique that remove heat from the inside surface. Film cooling also removes heat from the blade surface through the film hole by internal convection, q= h o  ( T f   -T w ), and film cooling is defined as, η  = (T m- T f  ) / (T m -T c ). Fig. 1: Schematic of Film Cooling Concept The thermal protection is expected to provide reduced heat load to the airfoil surface. Designers need to know the net heat load into the component surface when film is injected. Due to complex nature of discrete hole injection, there is need to know the local wall temperature (T w ) under the film and the gas side heat transfer coefficient with film injection. Both these components are required to estimate reduced heat load to the surface. Film cooling primarily depends on the coolant-to hot mainstream pressure ratio (Pc/Pt), temperature ratio (Tc/Tg), and the film cooling hole location, configuration and distribution on a film cooled airfoil. The coolant-to-mainstream pressure ratio can be related to the coolant-to-mainstream mass flux ratio (blowing ratio) the coolant-to mainstream temperature ratio can be related to the coolant-to-mainstream density ratio. In a typical gas turbine airfoil, the blowing ratios vary approximately from 0.5 to 2.50 while the Tc/Tg values vary from 0.5 to 0.85, corresponding density ratios approximately in the range of 1.3 to 1.5. Sang Woo Lee and Young Beom Kim [1], Surface flow visualizations show that the increase in the orientation angle results in better film coverage, especially in the spanwise direction, but produces more flow disturbances such as flow reversal and recirculation. A near-wall flow model for the velocity ratio of 2.0 has been proposed from the visualizations. the strength of downstream secondary flow strongly depends on the velocity ratio. Srinath Ekkad and Chin Han JE [2], studied the recent development in turbine blade film cooling accurate and detailed local heat transfer and film cooling for turbine edge region would be useful to prevent blade failure due to local heat spot. Flow visualization/ measurement, and the CFD predictions would provide valuable information for designing effective cooled blades for advanced gas turbines. D.G. Bogard and K. A. Thole [3] studied the gas turbine film cooling; the film-cooling performance is closely linked to whether the coolant jet has separated from the surface. For nominal conditions of a flat surface, low Freestream turbulence, and cylindrical holes, the film-cooling performance is reasonably predictable with empirical correlations. However, surface curvature, high Freestream turbulence, and shaping of the hole exit can greatly change film-cooling performance by significantly affecting the blowing ratio at which the coolant jet separates. CFD predictions, though very useful in providing insight in the spatial details of the film-cooling process, are also limited by the very complex flow conditions that occur for film cooling, particularly when the coolant jets begin to separate. Consequently, the film-cooling performance for actual turbine conditions is often difficult to predict precisely, and this remains a major constraint in the design for the durability of the turbine section. This study aims at investigating the film cooling effectiveness using both experimentally and numerically for the scaled up gas turbine blade leading edge compound angle model. The film cooling hole with the hole diameter of 4mm. Experiments are conducted by varying the blowing ratios in the range of 1.0 to 2.5 at the density ratio of 1.30, at a nominal flow Reynolds number of 1, 00,000 based on the leading edge diameter. 2.   EXPERIMENTAL SETUP AND PROCEDURE 2.1 Model Description Only the semicircular leading edge portion of gas turbine blade is taken in to our study, the model is generated using solid works and fabricated using rapid prototyping, using the model description as shown in Table 1, ABS-M30  Proceedings of the 2 nd  International Conference on Current Trends in Engineering and Management ICCTEM -2014 17 – 19, July 2014, Mysore, Karnataka, India   93 material is use to fabricate the leading edge model which is a low thermal conductivity material to avoid heat losses from the gas path non-contact side of the model. TABLE 1: Fabricated Turbine Blade Leading Edge Model Description SI.No. Model Description Dimensions 1 Leading Edge Outer Diameter 89 mm 2 Leading Edge Inner Diameter 65 mm 3 Film Cooling Hole Diameter and Pitch D=4 mm and Pitch= 21mm 4 Leading Edge Model Height 210 mm 5 No. of Rows 5 6 Compound Hole angles 0, 30, 45, 55 and 60 Deg. with Stream line direction 7 Hole Orientation Angle 0, 30 and 60 Deg. Angles from stagnation line The test models are prepared half cylindrical with flat downstream surfaces by attaching the coolant chamber. Hard foam is filled in the model slots to have the further low thermal conductivity. The Fig.2 and Fig.3 shows a fabricated model of turbine blade leading edge scaled up configuration. Stainless steel sheet having thickness of 0.15mm with a required film cooling hole geometry, machined by water jet cutting is wound over the model. The SS sheet with an area of 260 x 160mm is connected in series by brass bus bars to supply the high current at low voltage to heat the model. The reference thermocouples are soldered underside of the SS sheet for applying the correction factor to the thermo gram data and these are routed through the model slots. Fig. 2: Compound Angled Leading Edge Model The Fig.2 shows gas turbine leading edge models with the thermocouples, etc. respectively and the varied coolant hole angles are shown in Fig.3 and the models are prepared to withstand the experimental conditions.    Proceedings of the 2 nd  International Conference on Current Trends in Engineering and Management ICCTEM -2014 17 – 19, July 2014, Mysore, Karnataka, India   94 Fig. 3: Details of 0, 30, 45, 55, and 60 Deg. Hole Inclination Angles w.r.to Stream Line Direction 2.2 Experimental Setup Gas turbine leading edge model is mounted in the test module, consisting of rectangular duct with a size of 320mm x 230mm x 700 mm. Experimental test facility consists of compressed air unit, settling chamber, air filter, control valve, orifice meter and rectangular duct test sections where gas turbine leading edge models are placed. Air is selected as working fluid for both mainstream and coolant. Main stream flows through the settling chamber to the test section. The main flow is controlled by the gate valve placed much ahead of settling chamber. The coolant air to the model is passed through the heat exchanger, where the controlled liquid nitrogen is used cool the coolant air to have the required coolant temperature. The static and total pressures of mainstream flow to the inlet of test section are measured and maintained to have the required Reynolds number. The coolant flow passing through the orifice meter is also maintained by monitoring the upstream and differential pressures across the orifice meter. The required coolant flow is maintained to have the blowing ratios of 1.00, 1.50, 2.00 and 2.50. The mainstream and coolant temperatures are monitored and maintained to have the required density ratio. To measure the pressure and temperature of main stream and coolant air, pressure ports and thermocouples are incorporated at inlet and outlet of rectangular duct and inlet coolant chamber. Flir make Infra-red Camera is used for the non-contact type temperature measurement of the test surface as shown in Fig.4. The calibrated reference thermocouples are placed on the test model to correct thermo gram test surface data obtained by the Infra-red camera. Pressure net scanner is used for measuring pressures from pressure ports and the Fluke data acquisition is used to measure the temperatures of Thermocouples For these experiments. Fig. 4: Experimental Test Setup  The air is drawn from the centralized air compressor facility of NAL, with the continuous pressure of 60psi. The gate valve is used to maintain experimental pressure at the inlet to the test rig. The rectangular test section is fully
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