Chapter - Thermal Barrier Coatings
Thermal Barrier Coatings This is a coating system whose function is to reduce component temperature and thereby increase life. Thermal Barrier Coatings are generally a combination of multiple layers of coatings, with each layer having a specific function and requirement. The topmost layer provides thermal insulation and consists of a ceramic, with low thermal conductivity, typically ZrO2, known as zirconia. The ceramic insulating layer is deposited on the substrate alloy with an intervening oxidation resistant metallic layer called the “bond coat.” The metallic coating is either diffusion aluminide, such as platinum aluminide, or an overlay coating of general composition NiCoCrAlY, conforming somewhat to the substrate alloy composition. During the ceramic coating deposition, a thermally grown oxide (TGO), predominantly Al2O3, forms on the bond coat surface at the ceramic-bond coat interface. In effect, the TGO binds the ceramic layer to the bond coat. Thus, TBCs are the systems consisting of ceramic coating, the TGO, and a metallic bond coat on the substrate alloy (see figure 1). The approximate thickness range of each layer is as follows: ceramic 125 to 1000m, bond coat 50 to 125m, and TGO 0.5 to about 10m, the last depending on the duration of exposure of the TBS in the high-temperature environment of the coating process and in operation . x
Ceramic Topcoat Bond Coat
Figure 1: Thermal Barrier Coatings system
Temperature Reduction by TBCs  In order for TBCs to be effective in reducing surface temperature, the coated component needs to be actively cooled to remove heat conducted through the TBC. Turbine blades and vanes are good examples of systems effectively cooled by TBCs. The mechanism by which TBCs reduce component surface temperature is explained in figure 2, which is applicable to all the components just listed.
Active Figure 2: Mechanism by which TBCs reduce temperature of turbine blade Here, a turbine blade is given an example because it one of the most demanding of applications. The blade, coated with TBC, is actively cooled by cold air from the compressor of gas turbine engines. The cold air is routed through the internal cavity of the blade to remove the heat being conducted through the coatings and the blade walls. The cooling air needs exit passages, which may be provided by fine cooling holes drilled through the wall. The diameters of these holes are of the order of a few hundred microns, the length being proportional to the thickness of the blade wall. The holes are drilled either before the ceramic coating is deposited by the process of electrical discharge machining, of after the deposition of ceramic by lasers. The left side of Figure 2 depicts the concave outer wall of the blade, which sees higher pressure the convex outer wall depicted on the right side. Further away from the left outer wall, the gas temperature is very high. As one approaches the coated external surface of the blade, the
temperature is reduce somewhat by convection due to a film created by the relatively cooler air exiting the holes drilled through the walls. Across the ceramic, the temperature drops significantly due to its lower thermal conductivity. Conductive heat flow in three dimensions is given by the vector equation iQx jQ y kQz / t KA iT / x jT / y kT / z ,
where K = thermal conductivity, (i, j, k) are the unit vectors in Cartesian x, y, and z directions, and Q x / t , etc., are the respective heat fluxes passing across the cross section A. In one dimension, the heat transfer equation reduces to dQ / dt KAdT / dx
where dQ / dt is the heat flux and dT / dx is the local temperature gradient; the negative sign balances the negative slope of temperature or, in effect, indicates a temperature reduction. For thin coatings dT / dx can be approximated to T / x , where T is the total temperature drop across the ceramic coating of thickness x . Rearranging the equation, the temperature reduction due to the ceramic coating is given by T 1 / KA x dQ / dt
Several important conclusions can be drawn from the foregoing equation:
Because of the relatively high thermal conductivity, the metallic coatings have no significant impact in reducing surface temperature; that is T across the metallic coating is negligible.
The temperature drop across the ceramic coating is inversely proportional to the thermal conductivity of the ceramic. Therefore, the lower the thermal conductivity of the ceramic layer, the cooler the metallic surface will be. This provides a strong incentive to identify and develop lower conductivity ceramic coatings.
The temperature drop is proportional to the thickness of the coating. Thus, a thicker TBC is preferable. However, an unlimited increase of the thickness is not beneficial for three reasons. First, it adds parasitic weight to the part, which is detrimental to rotating components such as turbine blades. Second, thickness increased residual stress and total strain energy, making the ceramic coating prone to failure. Third, the more insulting the ceramic layer, the higher its surface temperature, leading to the possibility of sintering and associated structural and thermal problems.
The temperature drop is proportional to the heat flux. In other words, components that experience more heat flux going through, effected by the cooling air, would have a larger drop in temperature for the same TBC thickness. This explains why a first-stage turbine blade may experience a larger temperature reduction than a second-stage blade for identical thickness of TBC. 
The Benefits of TBC Figure 3 shows a map of temperature reduction on Pratt & Whitney PW2000 gas turbine engine that is used in nowadays aircraft, e.g. Boeing 757. The test compared the temperature on turbine blade with TBC and the one with metallic coating only. This 125-m-thick zirconia-based TBC was coated by electron beam physical vapor deposition (EB-PVD). In figure 3, the reduction of surface temperature at the local hot spots is about 139°C (250°F).
Figure 3: Surface temperature reduced on concave side of Pratt & Whitney PW2000 turbine blade. The efficiency of these engines can be directly correlated with the temperature capability of turbine blades. From 1965 through 1985, the temperature capability gain had been approximately 80°C (144°F) (Fig 4, Demasi Macin and Gupta, 1994), accomplished by the development of several generations of nickel base single crystals of increasing creep and fatigue strength.
Figure 4: Schematic representation of temperature capability improvement over the years. Further improvements of similar magnitude through improvement of the alloys appear unlikely. However, since the introduction of TBC in the early 1980s on the turbine blades in Pratt & Whitney JT9D engines powering jumbo jets, the effective systems temperature capability has been improved by as much as 200°C (360°F). The direct benefits of the temperature reduction due to the TBC are reflected in reduction in oxidation. For rotating components, creep damage is reduced significantly. The substrate alloy remains a higher percentage of the room-temperature structural properties. Additionally, compressor air used to cool components may be rerouted to produce thrust, thus increasing efficiency of engines . There are a few significant but indirect benefits of TBC to aircraft components. TBCs tend to reduce local variation in temperature, which results in reduced distortion and improved thermal fatigue life . During takeoff, climb, and landing of an aircraft, a turbine blade is facing transients in temperature which originate from a sudden change in
energy input to high-temperature. These transients directly affect to the thermal fatigue life of the metallic components. Here with the presence of ceramic coating with lower thermal conductivity, these transients will be smoothed out to improve thermal fatigue life of the components. Materials Requirement for TBCs Designing a better TBC for the environment or specific mission of an aircraft is always a big challenge for engineers. Metallic bond coat requires oxidation-resistant including the capability to form and maintain pure alumina scale with good bonding when exposing to oxidizing environment at high temperature. Because reduce temperature of the underlying metal structure is the most critical application of TBC and is inversely proportional to the thermal conductivity, the ceramic needs to have low thermal conductivity. Table 1 shows materials requirement for ceramic Thermal Barrier Coating. Table 1: Materials requirement for ceramic Thermal Barrier Coating
Place Table 7.1 from Mr. Bose’s Book Page 159
In gas turbine engine, the TBC surface may be exposed to hot gas with temperatures in excess of 3000°F (1649°C). In order to survive in such environments without melting, the ceramic topcoat of the TBC should also have a high melting point. Additionally, it needs to have good oxidation resistance and thermal expansion coefficient closer to that of metallic coatings. The last requirement reduces the thermal stresses arising due to a mismatch of the thermal expansion coefficient between the metallic coated substrate and the ceramic coating .