Report of THERMAL BARRIER COATING OF GAS TURBINE BLADES
A SEMINAR REPORT ON
THERMAL BARRIER COATING OF GAS TURBINE BLADES
SUBMITTED BY Ashutosh kumar tiwari Roll no.-0702740024 IIIrd B.Tech M.E
SUBMITTED TO Mr.Vivek Pansari Mr. J.S Behl
CREATION OF TBC
FAILURE OF TBC
1. INTRODUCTION The advances and improvements of current gas turbine technology have led to more efficient and more powerful engines. One very large improvement over the past fifty years has been the increase in maximum gas temperature that is produced in the gas turbine engine. Currently, temperatures in military aircraft turbines can reach upto 1500 degree Celsius and for commercial aircraft temperatures can reach upto 1400 degree Celsius. This increase in temperature has in turn increased efficiency, but has also introduced other problems such as the actual constraints in the materials itself. The super alloy that is used to create the turbines in a gas turbine engine has a melting point of around 1300 degree Celsius which as we can see is well below the operating temperature of the engine. To overcome these constraints three major advancements have been developed: (i)creating alloys that are more creep resistant and oxidation resistant, (ii) including channels in the blades so that air can flow through them, and (iii) development of a thermal barrier coating (TBC) on the turbine blade itself. Gas turbine engines use nickel- and cobalt-base super alloys in the turbine components, such as airfoils, combustors, transition ducts, and seals. Increased thermal efficiency demands of the newer engines require that the turbine inlet temperature be significantly increased. This must be accomplished without structural failure of components from melting, creep, oxidation, thermal fatigue, or other degradation mechanisms. Therefore, the surface temperature of these components must be maintained low enough to retain materials properties within acceptable limits. The demand is partially met by innovative component cooling schemes using compresor discharge air. However, cooling air is only available at the expense of loss of thrust and added fuel consumption. The major element in meeting the demand without significant performance loss comes from two parallel materials innovations: (i) improvement in the temperature capability of super alloys and (ii) development of thermal barrier coatings (TBCs) capable of providing thermal insulation equivalent to about 165 to 170 *C. TBCs are a bilayer system, consisting of a metallic layer (the bond coat) on the substrate on which a ceramic layer is deposited. In addition to (1)
Providing oxidation protection to the substrate, the metallic bond coat layer provides an adequately prepared surface to which the ceramic layer adheres. The ceramic layer composition is selected based on thermal conductivity, high-temperature stability, and thermal expansion compatibility with the substrate. Due to its low thermal conductivity and good thermo cyclic durability when stabilized against phase transformation, zirconia has become the ceramic of choice. The current bilayer design and composition of TBC have evolved from a multitude of designs and a number of compositions tested over a period of several decades.
Fig 1. Micrographs of La2Zr2O7 and YSZ coating
Some of products incorporating TBC: Turbine blades Cutting tool inserts Aero engine parts Gas turbine parts such as combustion liners, shroud, nozzles or blades IC engine parts Rocket nozzles and valve bodies Damper pin for turbine bucket Thermal
2. HISTORICAL BACKGROUND Pratt & Whitney first introduced TBCs on burner cans in JT8D engines in 1963. This TBC consisted of zirconia stabilized by the addition of 22 wt% MgO (22MSZ) to avoid the detrimental tetragonal (high-temperature phase) to monoclinic (lowtemperature phase) phase transformation. The ceramic was deposited on a flame sprayed Ni-Al bond coat. Subsequent evolution of TBCs incorporating refinements that allowed incremental increases in combustor exit temperature, is summarized in Table 1. In current systems, the ceramic consists of 7 wt% yttria partially stabilized zirconia (7YSZ). Depending on the application requirements, the bond coat is a variation of the NiCoCrAIY composition applied either by air plasma spray (APS) or EB-PVD method. Table 1: Material properties for a turbine blade
Density (kg/m3) Thermal
conductive(W/M*K) Specific heat (J/kg K)
coating system 3. DIFFERENT LAYERS
In an effort to protect the turbine blades from extremely high temperature gases; the thermal barrier coating (TBC) system has been implemented. The TBC system primarily consists of following layers: 1) Ceramic top-coat, 2) Thermally grown oxide (TGO), 3) Bond coat, and Ceramic top-coat: - The ceramic top-coat is the layer that provides thermal insulation for the blade. It has a very low thermal conductivity and has been designed using point defects to withstand thermal cycles. This layer is typically made of Y 2O3 which is stabilized with ZrO2 or in short YSZ. The thermal conductivity for this layer at a temperature of a thousand degrees Celsius is 2.3 W/(M*K) which is one of lowest conductivity of all the ceramics. In addition YSZ has a very high melting point (2700
degree Celsius) which makes it perfect for this application. Furthermore in an effort to reduce stresses in this material, cracks and porosity are intentionally incorporated into the material to make it “highly compliant (elastic modulus of 50 GPa) and strain tolerant.” Bond Coat: - Bond coat is a metallic one, whose function is, on the one side, to protect the basic material against oxidation and corrosion and, on the other side, to provide with a good adhesion to the thermal insulating ceramic layer. Such a ceramic coating is
(4) mostly made of yttria partially stabilized zirconia (YSZ), since this material has turned out particularly suitable during the last decades. Bond coat is about 75-150 micrometers thick. It is an oxidation-resistant metallic layer and is primary used to hold the ceramic top coat to the substrate. This layer is typically made of Ni and Pt and in some cases can
Thermally grown oxide: - TGO layer was not intended but was created when the ceramic top-coat reacts with the bond coat in very high temperatures. This layer is about one to ten micro meters thick and its growth is slow, uniform, and defect free.”
Figure 3. A schematic illustration of a modern thermal barrier coating system consisting of a thermally insulating thermal barrier coating, a thermally grown oxide (TGO) and an aluminum rich bond coat. The temperature gradient during engine operation is overlaid.
structure b)EB-PVD columnar structure
(6) 4. CRAEATION OF TBC
The bond coat is primarily deposited by “electroplating in conjunction with diffusionaluminizing or chemical-vapour deposition.”
Recently, the high velocity oxy-fuel
technology (HVOF) is also increasingly used for this task.
The Ceramic top-coat on other hand is currently being deposited in two different ways: 1) Air plasma- sprayed (APS) deposition and 2) Electron-beam physical-vapour deposition (EB-PVD). A picture of the different microstructures created by each method are shown in Figure 4. Air plasma- sprayed (APS) deposition: - In the APS powder application method, the ceramic material is in the form of a flowable powder that is fed into a plasma torch and sprayed molten onto the surface of the metallic substrate. Droplets of molten material form “splats” on the metallic substrate. Sprayed coatings have half the thermal conductivity of the EB-PVD coatings and are therefore better isolators.
Fig 5. Schematic microstructure of thermal spray coating, showing only a few layers of particles
(7) In an APS top coat, the orientation of the cracks and pores are normal to the flow of heat, which reduces the thermal conductivity from 2.3 W/ (m*K) to about 0.8-1.7 W/ (m*K). However, because of the orientation of the micro structure and the roughness of the interface, this method generally produces a shorter thermal cycle compared to EBPVD. This can be seen in Figure 4.The plasma spraying technology applies vacuum plasma-sprayed bond coats, which are mostly made of MCrAlYs (M=Ni, Co). The ceramic layer is usually deposited by means of the atmospheric plasma spraying (APS). In the institute IWV1, thermal barrier coating systems are being produced through plasma spraying, whereas also the HVOF process is also adopted for the deposition of bond coats. The main goal of the activities by the IWV1 in the research field of the thermal barrier coatings is the development of systems for higher application temperatures and with a longer lifetime.
Electron-beam physical-vapour deposition:- EB-PVD is the process currently recommended for high-quality coatings. In this technique, a cylindrical ingot of the coating material is vaporized with an electron beam, and the vapor uniformly condenses on the surface on the turbine blade. Contrary to APS structure, the EB-PVD structure is columnar which prevents the build up of tensile stresses and increases the life span of this material as noted earlier. The porosity and the cracks within this structure also help reduce the thermal conductivity to about 1.5-2 W/(m*K) which as we can see is higher than APS. Overall because of EB-PVD life span this deposition method is used to coat turbine blades in a jet engine. EB-PVD process is usually used to deposit an aluminum layer as bond coat.
Fig 6. Schematic EBPVD process, the whole assembly would be under vacuum. Rotation of the electron beam is obtained by a magnetic field perpendicular to the drawing
(8) 5. FAILURE OF TBC Failure of thermal barrier coating systems under cyclic thermomechanical loading:The failure mechanisms of thermal barrier coating (TBC) systems applied on gas turbine blades and vanes are investigated using thermomechanical fatigue (TMF) tests and finite element (FE) modeling. TMF tests were performed at two levels of applied mechanical strain, namely five times and three times the critical in-service mechanical strain of an industrial gas turbine. TMF testing under the higher mechanical strain of air plasma-sprayed (APS) and electron beam-physical vapor deposition (EBPVD) coated samples showed that both systems failed after a similar number of cycles by cracks that initiated at the bond coat/thermally grown oxide (TGO) interface and propagated through the bond coat to the substrate. When the applied mechanical strain
was decreased, cracking of the bond coat in EB-PVD coated systems was suppressed, the life of the coated system increased significantly and delamination of the top-coat was observed. A subsequent FE analysis showed that, by subjecting the system to the higher mechanical strain, significant tensile stresses develop in the TGO and the bond coat that are thought to be responsible for the observed crack initiation and propagation. The FE model also predicts that cracking initiates at specific geometric features of the rough interface of a PS coated system, which was confirmed by metallographic examination of failed samples. The decrease of the applied mechanical strain and hence of the developed stresses led to the suppression of failure by bond coat cracking and activate delamination. These results outline the importance of designing TMF tests and selecting the appropriate mechanical loading in order to accelerate testing and still trigger the same failure mechanisms as observed in-service.
Oxidation-Induced Failure of Thermal Barrier Coatings:On prolonged high-temperature exposure in air, thermal barrier coatings (TBCs) on bond-coated super alloys fail by spalling. It is a new discovered TBC failure mode, in which failure is associated with moisture-enhanced sub-critical crack-growth along the bond-coat/ thermally grown oxide interface. By making concurrent piezospectroscopy measurements, the interfacial fracture energy was determined to be ^ 10 J/m(exp 2) - a considerably smaller value than that of sapphire/metal interfaces prepared in the laboratory but consistent with measurements of the effects of segregation on metal/ceramic interfaces. New insights into the mechanism underlying failure of the thermally grown oxide have come from direct optical microscopy. These indicate that failure is associated with surface roughness of the bond-coat and specifically that the thermally-grown oxide separates from the bond-coat on cooling at the concave ("crests") surface features. These locally separated regions grow with oxidation time and are seen to link-up. These events are believed to be the precursor events that grow to provide the critical-sized flaws from which buckling and spalling of thermal barrier coatings occur.
6. MATHEMATICAL BACKGROUND Because of the complexity of this model, the majority of the results are done through experimentation. In order to get a physical feel for what is happening, we have made several assumptions in our analysis: 1) Tgas = T1 = 1873 degrees Kelvin which is kept constant. 2) T4 = Tair = 298 degrees Kelvin which is kept constant. 3) Assume that a small portion of the air foil that I represent as a rectangle is a good model of the entire air foil. 4) Properties of state stay constant throughout the model because boundary conditions stay constant. 5) No heat flux through the top and bottom of our model (insulated). 6) TGO layer has no effect on temperature distribution This information combined with table 1 was inputted into COMSOL and solved. The results are shown in figure 6. Table 1: Material properties for a turbine blade
Density (kg/m3) Thermal
conductive(W/M*K) Specific heat (J/kg K)
Figure 7: 2-D Diagram of the blade
Figure 8: Comsol model of the temperature distribution in a turbine blade
As one can see in Figure 6, the temperature distribution drops dramatically through the ceramic top coat because of its very low thermal conductivity. Furthermore, these results show that the temperature distribution through the supper alloy is well below the melting point of the metal which is what was expected. (12)
7. PROPOSED RESEARCH The only way to improve efficiency in a gas turbine engine is to increase the temperatures within the system. Because of this more research needs to be done in improving reliably and thermal conductively of the thermal barrier coat. Focus needs to be placed on the current faults of TBC system and how they can be resolved. Some problems are with oxidation of the material, and lack of modelling to predict the life span of the thermal barrier coating. In addition, new research in different ways of cooling the blade needs to be done. A reasonable area to focus on would be water cooling and nitrogen cooling. With the collected thermal diffusivity values, the specific heat and thermal conductivity values are remaining to be calculated. Along with more testing, the thermal conductivity values would be ready for comparison to yttria-stabilized zirconia with the goal of creating a potential new thermal barrier coating. The main goal of the activities by the institute IWV1 in the research field of the thermal barrier coatings is the development of systems for higher application temperatures and with a longer lifetime. For that purpose, works are being carried out in following basic areas. These areas are1) Thermal spraying 2)
3) Thermal cycling
Thermal spraying: - One main area is thermal spraying by means of the atmospheric and the vacuum plasma spraying, as well as the high velocity flame spraying (HVOF). Special attention deserves the reproducible fabrication of coatings. The corresponding equipment to guarantee a high reproducibility has been implemented and is being continuously improved. Another part of the activities in the institute corresponds to the diagnostics works, which cover the plasma (enthalpy probe) as well as the particle diagnostics (DPV2000). The collected data are being used also in the modelling of the plasma spraying process. Within the scope of a European Union project, highly efficient Solid Oxide Fuel Cells (SOFC) was produced using Atmospheric Plasma Spraying. (13)
Modelling: - Modelling constitutes a second main research area. The principal goal in the modelling of the plasma spraying process is the improvement of our understanding about the relation between the process parameters and the structure and effectiveness of the coating systems. Modelling yields basic information, which can lead to the optimization of the process. A similar function corresponds to the modelling of thermal barrier coating life. The achieved development of a lifetime model based on the microstructure gives important indications for an optimized layer system. Thermal cycling: - A further research field is the characterization of layers. Beside the determination of the microstructure by means of metallography, Hg-porosimetry and specific surface analysis, the thermal cycling of TBC-systems has an outstanding importance for the efficiency evaluation. Ceramic Matrix Composites (CMCs). Further increases in temperature are likely to require the development of ceramic matrix composites. A number of simply shaped static components for military and civil applications are in the engine development phase and guide vanes for axial compressors have been manufactured to demonstrate process capability, such techniques involve advanced textile handling and chemical vapor infiltration that provide the ultimate challenge. It will eventually appear because the rewards are so high, but it will take much longer to bring it to a satisfactory standard than was anticipated a couple of decades back. Ceramic matrix composites are at the forefront of advanced materials technology because of their lightweight, high strength and toughness, high temperature capabilities, and graceful failure under loading. Research work has concentrated for some years on fiber reinforced ceramics for this application, as opposed to monolithic materials which possess adequate strength at high temperatures but the handicap of poor impact resistance. Today's commercially available ceramic composites employ silicon carbide fibers in a ceramic matrix such as silicon carbide or alumina. These materials are capable of uncooled operation at temperatures up to 1200°C, barely beyond the capability of the current best coated nickel alloy systems. Uncooled turbine applications will require an all oxide ceramic material system, to ensure the long term stability at the very highest (14)
temperatures in an oxidizing atmosphere. An early example of such a system is alumina fibers in an alumina matrix. To realize the ultimate load carrying capabilities at high temperatures, single crystal oxide fibers may be used, giving the possibility to operate under temperatures of 1400°C. Higher operating temperatures for gas turbine engines are continuously sought in order to increase their efficiency. However, as operating temperatures increase, the high temperature durability of the components of the engine must correspondingly increase. Significant advances in high temperature capabilities have been achieved through formulation of iron, nickel and cobalt-base super alloys. While super alloys have found wide use for components throughout gas turbine engines, alternative materials have been proposed. Materials containing silicon, particularly those with silicon carbide (SiC) as a matrix material and/or as a reinforcing material, are currently being considered for high temperature applications, such as combustor and other hot section components of gas turbine engines; like, combustion chambers, transition ducting (which takes the combustion products and directs them towards the turbine section), the nozzle guide vanes, the surrounding shroud section, and others.
8. CONCLUSION Gas turbines engine constitute a wide and good option for power generation used for both, industrial and aeronautical applications. This technology is requesting for better and more reliable materials to use mostly in those areas in which temperatures are extremely high; like, fist row of turbines blade and combustion chamber. Blades materials for turbine area in gas turbines have advanced rapidly in the last 2 decades. These blades are constructed using special alloys and are covered by special coats. Modifications are intended to increase the allowed temperature up to 1500OC without cooling. To increase efficiency. Ceramic coating is applied to the surface of the turbine blade using several methods. The most important ones are: Electron Beam Physical Vapor Deposition (EBPVD) and Arc Plasma Sprayable (APS) powder method. Besides the technology aimed to produce better coats, materials science is currently working extensible in Ceramic Matrix Composites, formed basically by silicon carbide fibers and special fabrics in order to increase the temperature gap in locations specially sensible for gas turbine operation. The last and, at the same time, largest area deals with the development of new thermal barrier coatings. In this wide activity, oxidic materials are principally being studied as candidates for thermal barrier coatings and, where applicable, developed to new TBC-systems. The experience and know-how in the other research areas (processing, characterization, modelling) are being intensely used for the further development of new systems.
(16) 10. REFERENCES
1) Clarke, D. R. and Levi, C. G. (2003). “Materials Design for the Next Generation Thermal Barrier Coatings.” Annual Reviews of Materials Research: Vol. 33, pp. 383417. 2) Nitin P. Padture, Maurice Gell, Eric H. Jordan. "Thermal Barrier Coatings for GasTurbine Engine Applications." Science 12 April 2002: Vol. 296. no. 5566, pp. 280 284 3) O’Donoghue Lisa. "Why don’t Gas Turbines Blades burn?" 4) n/a. "How does a modern gas turbine engine work?" December 4, 2007. < http://www.mtu.de/en/take-off/how_engines_work/index.html> 5) Marshall, Brian. "How Gas Turbine Engines Work." 6) W.G. Marijnissen and A. van Lieshout. "The evolution of thermal barrier coatings status and upcoming solutions for today's key issues." Surface and Coatings Technology. Vol 120- 121, November 1999, pp 61-67 7) Dr. Clarke and C.G. Levi. “Materials Design for the Next Generation Thermal Barrier Coatings.” Annual Review of Materials research. Vol 33:383-417. August 2003 8) K.A. Khor, Y.W. Gu. “Effects of residual stress on the performance of plasma sprayed functionally graded Zr02/NiCOCrAlY coatings”. Materials Science and Engineering A277 (2000) 64–76