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 1000m, bond coat 50 to 125m, and TGO 0.5 to about 10m,
the last depending on the duration of exposure of the TBS in the high-temperature
environment of the coating process and in operation [1].
x
Ceramic
Topcoat
Bond Coat
125-1000m
50-125m
TGO 0.5-10m
Substrate
Figure 1: Thermal Barrier Coatings system
Temperature Reduction by TBCs [1]
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  jQy  kQz / t   KAiT / x  jT / y  kT / 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 firststage turbine blade may experience a larger temperature reduction than a secondstage blade for identical thickness of TBC. [1]
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 [1].
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 [1]. 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 [1].