Design and Analysis of Gas Turbine Blade with Varying Pitch of Cooling Holes Lalit Dhamecha #1, Shubham Gharde *2, Ganraj More #3, M.J.Naidu *4 #1 Department of Mechanical Engineering Smt. Kashibai Navale College of Engineering, Pune, India. 2 Department of Mechanical Engineering, Smt. Kashibai Navale College of Engineering, Pune, India. 3 Department of Mechanical Engineering, Smt. Kashibai Navale College of Engineering, Pune, India. 4 Assistant Professor Department of Mechanical Engineering, Smt. Kashibai Navale College of Engineering, Pune, India Abstract Gas turbine play a vital role in the today s industrialized society, and as the demand for power increase, the power output and thermal efficiency of gas turbine must also increase. One method of increasing both the power output and thermal efficiency of the engine is to increase the temperature of the gas entering the turbine. In the advanced gas turbine, the inlet temperature of around 1500 C is used, however, this temperature exceeds the melting temperature of the metal airfoils. Therefore, along with high temperature material development, a refined cooling system must be developed for continuous safe operation of gas turbines with high performance. Gas turbine blades is cool internally and externally. This paper is mainly focus on external cooling of the turbine i.e. film cooling. In film cooling, relatively cool air is injected from the inside of the blade which travels through the entire blade length and form a protective film around the blade-surface. In present work attempt has been made to analyse the failure of the gas turbine blade through structural analysis. The analysis is conducted for two different blade configuration one is the base line configuration with film cooling cylindrical holes along the entire length and in another configuration is the holes along the leading edge are branched together for anti-vortex considerations. The two configurations is further study for two different pitch to diameter ratio(y/d) of the cooling holes. Keywords Gas Turbine, Film Cooling, Failure Analysis, Structural Analysis. I. INTRODUCTION Gas turbines is the main component in industries like power generation; processing plant and aircraft propulsion. In 1960s, by material property limited the turbine blade and gas turbine firing temperature around the 800C.But now a day s gas turbine engines operate at high temperatures (1200-1500 C) to improving thermal efficiency and power output. Due to increasing the gas temperature, the heat transfer to the blades will also increase significantly resulting in their thermal failure. With the existing materials, it is not possible to go to higher temperatures. Therefore a suitable cooling method must be developed for continuous safe operation of gas turbines with high performance. In order to employ high gas temperature in gas turbine stages, it is necessary to cool the casing, nozzles, rotor blades and discs. Cooling of these components can be achieved either by air or liquid cooling. By using the liquid cooling method to face the some disadvantages of these processes i.e. the problems of leakage, corrosion. Besides other side, air cooling method, it s allows to be discharged into the main flow without any problem. The blade metal temperatures can be reduced by around 200-300 C. By using suitable blade materials (nickel-based alloys) now available, an average blade temperature of 800 C can be used. This gives the permit maximum gas temperatures around 1200-1500 C. Air cooling method is briefly described in two main parts i.e. internal cooling & external cooling Internal cooling of the blade can be achieved by passing cooling air through internal cooling passage from hub towards the blade tips. The internal passages may be circular or elliptical and are distributed near the entire surface of a blade. The cooling of the blade is achieved by conduction and convection. Relatively hot air after traversing the entire blade length in the cooling passages escapes to the main flow from the blade tips. A part of this air can be usefully utilized to blow out thick boundary layers from the suction surface of the blades. Hollow blades can also be manufactured with a core and internal cooling passage. Cooling air enters the leading edge region in the form of jet and then turns towards the trailing edge. Cooling of the blade takes place due to both jet impingement (near the leading edge) and convection heat transfer. External cooling of the turbine blades is achieved in two ways. The cooling air enters the internal passage from hub towards the tips. On its way upward it s allow to flow over the blade surface through a number of small orifices inclined to the surface. A series of such holes is provide at various sections of the blade along their lengths. The cooling air flowing out of these small holes forms a film over the blade surface. Besides cooling the blade surface it decreases the heat transfer from the hot gases to the blade metal. Another variation of this method is the blade surface is made of a porous wall ISSN: 2231-5381 http://www.ijettjournal.org Page 383
which is equivalent to providing an infinite number of orifice. Cooling air is force through this porous wall which forms an envelope of a comparatively cooler boundary layer or film. This film around the blade prevents it from reaching very high temperatures. Besides the effusion of the coolant over the entire blade surface causes uniform cooling of the blade. The present paper attempts to study the effect of variation in blade configuration and pitch to diameter ratio(x/d) of the cooling holes on stress distribution. The blade is analyses for two x/d ratio of 3 and 4. The structural analysis of turbine blade is done using ansys 16 workbench Solver. The material used for the analysis is inconel718. II. LITERATURE SURVEY Theju V et.al had mainly done the research work on jet engines turbine blade; the study was done on two different materials Inconel -718 and Titanium T- 6; to investigate the effect of temperature and induced stresses on turbine blade. The study concluded that the Titanium T-6 would have lesser value of deformation and lower strength; however if cost of material is the primary issue then inconal-718 could be selected it would have little higher deformation and higher strength. Also Inconel have good material properties at higher temperature than titanium. V. Veeragavanet.al had done their research on aircraft turbine blades, his main focus was on 10C4/60C50 turbine blade models. Conventional alloys Such as titanium, zirconium, molybdenum were chosen for analysis.he studied the effect of temperature on different material for the certain interval of times and concluded that molybdenum had better temperature resistance capability. G Narasa Raju et.al had done research on different types of cooling techniques which maintain temperature of bade within allowable limits. Finite Element Analysis is used to study steady state thermal and structural performance for N155 and Inconel-718.four different models consisting of solid blade and blades with varying number of holes(5,9,13) were analysed to find out the number of cooling holes. They had used two materials Inconel 718 and Inconel 155 for their research and found out Inconel 718 has better thermal properties as blade temperature and stress induce is lesser. V. NagaBhushana Rao et.al had done research on turbine blade used in marine applications. The blade was investigated for structural analysis at elevated temperatures and under the action of large centrifugal force, the material used was nickel based super alloy, it was observed that maximum stresses and strains were observed near to the root of the turbine blade and upper surface along the blade root, maximum temperature is observed at the blade tip and minimum at the root of the blade; temperature distribution is decreasing from tip to root and temperatures observed are below the melting point of blade material. Bhatti et al.performed transient state stress analysis on an axial flow gas turbine blade and disk using finite element techniques. They have chosen Inconel 718, a high heat resistant alloy of chromium, nickel & niobium. The study was focused on centrifugal & thermal stress arising in the disk. Deepanraj et.al. had considered that the titanium aluminium alloy as the blade material and performed structural and thermal analysis with varying number of cooling passages. They further studied the effect of varying the cooling air temperature on the temperature distribution in the blades. It is concluded that the blade configuration with 8 holes gives an optimum blade temperature of 800C. A.K.Mattaet.al.studied the stressanalysis for N 155 & Inconel 718 material. On solid blades it is reported that Inconel 718 show better results for high temperature operation. Ervanet. al.had suggested that high turbine efficiency can be obtained by minimizing the air flow required for cooling by effectively utilizing its cooling potential. He suggested a cooling technology which has three main parts:- a) The leading edge is provided with impingement cooling; b) the middle section of blade contains cooling pipes with obstacles provided along the length to enhance turbulence in the cooling air and c)the trailing edge of the blade is provided with pin fins for effective cooling. G. Narendranath et.al.examine the first stage rotor blade off the gas turbine analyzed using ANSYS 9.0. The material of the blade was specified as N155. Thermal and structural analysis is done using ANSYS 9.0 Finite element analysis software. The temperature variations from leading edge the trailing edge on the blade profile is varying from 839.531C to 735.162C at the tip of the blade. It is observed that the maximum thermal stress is 1217 and the minimum thermal stress is the less than the yield strength value i.e., 1450. K haribrahmaiah et.al. Examined the thermal analysis of gas turbine with four different models consisting of blade with and without holes and blades with varying number of holes(5,9&13) were analyzed. Transfer rate and temperature distribution, the blade with 13 holes is considered as optimum. Steady state thermal and structural analysis is carried out using ANSYS software with different blade materials of Chromium steel and Inconel- 718. While comparing these materials Inconel-718 is better thermal properties and induced stresses are lesser than the Chromium steel. V.Vijaya Kumar et.al.examine the preliminary design of a power turbine for maximization of an existing turbojet engine. For proper understanding of the combined mechanical and the thermal stresses for the mechanical axial and centrifugal forces. The ISSN: 2231-5381 http://www.ijettjournal.org Page 384
peripheral speed of rotor and flows velocities is kept in the reasonable range so to minimize losses. In which the base profiles is analyzed later for flow condition through any of the theoretical flow analysis method such as potential flow approach. III. Blade Design Gas turbine is a power producing mechanical device. Combustion gases from the combustor directly goes to turbine blade. Blade is the main component of any gas turbine that extract the power from high temperature and high pressure gas. Blade is made of material that can withstand high temperature. Inconel alloy and titanium are generally used. Blade cooling is studied by taking Inconel as material as it is less costlier than titanium and give performance in the similar way. A. Material properties Blade material:-inconel (Ni-Cr alloy) Properties Unit Inconel 718 Young s MPa 2E5 modulus Density kg/m3 8190 Poisson s ratio 0.284 Tensile yield MPa 1069 strength Allowable stress MPa 641.8 Allowable Shear MPa 385.08 Stress Specific heat J/kg-K 586 Thermal Conductivity W/m-K 25 B. Steps In Blade Design The Blade model profile is generated by using CATIA V5 software. Key Points are created along the profile in working plane. The points are joined by drawing Spline curves to obtain a smooth contour. The contour(2d) model is then converted into area and then volume(3d) model was generated by extrusion. The profile was extruded in y direction up to 117mm i.e. the blade height. The hub is also generated similar grounds. These two volumes are then combined into single volume. The blade with two pitch to diameter ratios (y/d) 3 and 4 are modeled and analyzed for structural failure. Where d= Diameter of cooling holes. y= Distance between two cooling holes. Fig. 1 Catia Model Of The Blade C. List Of Selected Key points SR. X Z NO 1 0.07-0.768 2 80.761 5.864 3 15.055 9.472 4 21.565 12.672 5 28.282 15.41 6 25.188 17.631 7 42.251 19.282 8 49.43 20.318 9 56.672 20.712 10 63.921 20.452 11 71.118 19.55 12 78.211 18.036 13 85.16 15.954 14 91.33 13.356 15 98.51 10.296 16 104.883 6.83 17 104.883 3.83 18 97.195 5.446 19 89.384 6.283 20 81.528 6.423 21 73.687 5.92 22 65.913 4.783 23 58.274 2.95 24 50.886 0.283 25 43.888-3.284 26 37.151 7.331 27 30.187-10.964 28 92.782 13.572 29 15.054-14.95 30 7.216-14.78 31 0.18-11.169 IV.FINITE ELEMENT ANALYSIS OF A GAS TURBINE BLADE The turbine blade is analysed for its structural performance under the action of various forces. The structural analysis of a gas turbine is carried out ISSN: 2231-5381 http://www.ijettjournal.org Page 385
using ANSYS 16.0 software. Single blade is taken into consideration for analysis as turbine blades and is mounted on the periphery of hub symmetrically along the axis of rotation of the blade. The cross section of the blade is in the X-Z plane and the length of the blade is along the Y axis. Centrifugal forces generated during service by rotation of the disc were also taken into consideration. Static analysis was carried out to determine the mechanical stresses and elongation experienced by the gas turbine blade. In this analysis, the gas forces are assumed to be distributed evenly, the tangential and axial forces acts along the centroid of the blade. The centrifugal force also acts through the centroid of the blade in the radial direction. Tangential Force (Ft) = 980 N Axial Force (Fa) = 250 N Centrifugal Force (Fc) = 282661.8 N Fig. 3 Total Deformation Of Blade With Y/d=4 V. RESULTS AND DISCUSSION Blade failures can be caused by a number of mechanisms under the turbine operating conditions of high rotational speed at elevated temperature in corrosive environments. To identify the causes of this failures mechanical analysis of the turbine blade was conducted. This work focuses on failure analysis of gas turbine blades through structural analysis. The present work deals with the modelling and analysis of gas turbine blades. The structural analysis was performed on the turbine blades using ANSYS 16.0 software and results were discussed. The force of high temperature gases impinging on the turbine blade is divided in two components: tangential force (Ft) and axial force (Fc). The Centrifugal, Axial and Tangential forces acting on the blade are considered as loads in structural analysis. The maximum mechanical stresses and elongations induced in the turbine blade are observed to be within the safe limit. Maximum elongations are observed at the tip section of the blade and minimum elongations at the root Section of the blade. Fig. 4 Von-Mises Stress Of Blade With Y/d=3 A. Blade With Cylindrical Holes Fig-5: Von-Mises Stress Of Blade With Y/d=4 Fig. 2 Total Deformation Of Blade With Y/d=3 ISSN: 2231-5381 http://www.ijettjournal.org Page 386
B. Blade With Branched Cylindrical Holes At Leading Edge Fig. 6 Total Deformation Of Blade With Y/d=3 Fig. 7 Total Deformation Of Blade With Y/d=4 Fig. 8 Von-Mises Stress Of Blade With Y/d=3 Fig. 9 Von-Mises Stress Of Blade With Y/d=4 Above analysis Shows that the total deformation due to action of all forces such as axial, tangential and centrifugal forces is maximum at the tip section of the turbine blade material and minimum deformation is observed at the root section. The maximum deformation observed is 0.22522 mm for blade having branched holes at leading edge and having y/d=4. The minimum deformation observed is 0.13485 mm for blade having cylindrical holes and y/d=4.the Von-Mises due to action of all forces such as axial, tangential and centrifugal forces is maximum at the root section of the turbine blade material and minimum stress is observed at the tip section. The maximum Stress observed is 1878.5 Mpa for blade having branched holes at leading edge and having y/d=4. The minimum stress observed is 1122.6 Mpa for blade having cylindrical holes and y/d=3. It is Observed in all the cases that the maximum stress occurring at the tip is at a very small area and that the majority of the blade section is having stress within the allowable limits. The high stress zones can be avoided by reducing the number of holes at the root section or by providing a fillet at the root to avoid abrupt change in cross-section and reduce stress concentration in that area. VI. CONCLUSION Gas turbine blade with film cooling holes is studied for its structural performance considering Inconel 718 as the blade material. The analysis was conducted on ANSYS 16.0 software. Maximum elongations are observed at the tip section and minimum at the root section of the blade. The elongations observed are within the safe limit. Maximum stresses are observed near the root section and minimum at the tip section of the blade. The maximum stresses are observed near the film cooling holes due to high stress concentration along the holes. The maximum stress of 1878.5 Mpa occurs at the trailing edge nearer to the root of the ISSN: 2231-5381 http://www.ijettjournal.org Page 387
blade with y/d=4. This stress exceeds the yield stress of the material and which might lead to the failure of the turbine blade. At all other parts of turbine blade, the stresses induced are within the same limits. [6] Amjed Ahmed Jassim,A L Luhaibi,Mohammad Tariq Thermal Analysis Of Cooling Effect On Gas Turbine Blade,International Journal Of Research in Engineering And Technology, ISSN:2319-1163. VII. REFERENCES [1] Theju V,Uday P S,PLV Gopinath Reddy,C.J. Manjunath, Design And Analysis Of Gas Turbine Blade, International Journal of Innovative Research in Science, Engineering and Technology, ISSN: 2319-8753. [2] V.NagaBushana Rao,N. Niranjan Kumar,N. Madhulata,A. Abhijeet, Mechanical Analysis Of 1 st Stage Marine Gas Turbine Blade,International Journal Of Advanced Science And Technology vol.68(2014).pp.57-64. [3] R D V Prasad,G Narasa Raju,M S Srinivasa Rao,N Vasudeva Rao, Steady State Thermal And Structural Analysis Of Gas Turbine Blade Cooling System,International journal Of Engineering Research And Technology, ISSN:2278-0181. [4] K Hari Brahmaiah,M lava Kumar, Heat Transfer Analysis Of Gas Turbine Blade Through Cooling holes, International Journal Of Computational Engineering Research, ISSN:2250-3005. [5] Aqeel Jomma Athab,Dr, Naga Sarada, CFD Analysis Of A Gas Turbine Blade Cooling In The Presence Of Holes,International Journal And Magazine Of Engineering,Technology,Management And Research, ISSN: 2348-4845. ISSN: 2231-5381 http://www.ijettjournal.org Page 388