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Heat engines are based on considering various factors such as durability, performance and efficiency with the objective of minimizing the life cycle cost. For example, the turbine inlet temperature of a gas turbine having advanced air cooling and improved component materials is about 1500oC. Metallic coatings were introduced to sustain these high temperatures. The trend for the most efficient gas turbines is to exploit more recent advances in material and cooling technology by going to engine operating cycles which employ a large fraction of the maximum turbine inlet temperature capability for the entire operating cycle. Thermal Barrier Coatings (TBC) performs the important function of insulating components such as gas turbine and aero engine parts operating at elevated temperatures. Thermal barrier coatings (TBC) are layer systems deposited on thermally highly loaded metallic components, as for instance in gas turbines. TBC’s are characterized by their low thermal conductivity, the coating bearing a large temperature gradient when exposed to heat flow. The most commonly used TBC material is Yttrium Stabilized Zirconia (YSZ), which exhibits resistance to thermal shock and thermal fatigue up to 1150oC. YSZ is generally deposited by plasma spraying and electron beam physical vapour deposition (EBPVD) processes. It can also be deposited by HVOF spraying for applications such as blade tip wear prevention, where the wear resistant properties of this material can also be used. The use of the TBC raises the process temperature and thus increases the efficiency.



Thermal Barrier Coating consists of two layers (duplex structure).

The first layer, a metallic one, is called bond coat, whose function is to protect the basic material against oxidation and corrosion. The second layer is an oxide ceramic layer, which is glued or attached by a metallic bond coat to the super alloy. The oxide that is commonly used is Zirconia oxide (ZrO2) and Yttrium oxide (Y2O3). The metallic bond coat is an oxidation/hot corrosion resistant layer. The bond coat is empherically represented as MCrAlY alloy where

M Y CrAl


Metals like Ni, Co or Fe. Reactive metals like Yttrium. base metal.

Coatings are well established as an important underpinning technology for the manufacture of aeroengine and industrial turbines. Higher turbine combustion temperatures are desirable for increased engine efficiency and environmental reasons (reduction in pollutant emissions, particularly NOx), but place severe demands on the physical and chemical properties of the basic materials of fabrication.

In this context, MCrAlY coatings (where M = Co, Ni or Co/Ni) are widely applied to first and second stage turbine blades and nozzle guide vanes, where they may be used as corrosion resistant overlays or as bond-coats for use with thermal barrier coatings. In the first and second stage of a gas turbine, metal temperatures may exceed 850°C, and two predominant corrosion mechanisms have been identified:

Accelerated high temperature oxidation (>950°C) where reactions between the coating and oxidants in the gaseous phase produce oxides on the coating surface as well as internal penetration of oxides/sulphides within the coating, depending on the level of gas phase contaminants

Type I hot corrosion (850 - 950°C) where corrosion occurs through reaction with salts deposited from the vapour phase (from impurities in the fuel). Molten sulphates flux the

3 oxide scales, and non-protective scales, extensive internal suplhidation and a depletion zone of scale-forming elements characterize the microstructure.

The choice of base material (Co or Ni) is dependent on the primary corrosion mechanism, but as engine temperatures increase, the trend is towards CoNiCrAlY compositions. Cr and Al are present in the MCrAlY composition because they form highly tenacious protective oxide scales, whilst Y promotes formation of these stable oxides. MCrAlY coatings may be applied by a number of processes including:

i. ii. iii.

Physical vapour deposition (PVD). Low pressure (LPPS), vacuum plasma (VPS) or air plasma spraying (APS). High velocity oxyfuel (HVOF) spraying.

PVD and VPS offer high quality in terms of minimal oxidation of the coating during the deposition process, but are the most expensive. Approvals have been granted for the use of APS and HVOF coatings on certain components with significant cost savings. It is common for MCrAlY coatings to be deposited onto components pre-coated with Al, PtAl or Cr, which have been produced by vapour deposition techniques or diffusion processes.

The concept of TBC is illustrated in Fig 1

Fig 1. Schematic concept of Thermal Barrier Coating.

4 Barrier coat is defined as an exterior coating applied to a composite filament, wound structure to provide protection. In fuel tanks, a coating applied to the inside of the tank prevents fuels from permeating the sidewalls.

The conventional metallic coated component is air-cooled and thus can operate even under conditions where the gas temperature exceeds its melting point. By applying TBC, the difference between gas and metal temperature can be further increased because of large temperature drop (∆T) established across the ceramic thermal barrier layer. The magnitude of ∆T is a function of the coating thickness, heat flux and oxide thermal conductivity. A 0.4 mm thick ceramic layer can typically give a 100 to 300 oC ∆T. Both the bond coat and oxide layer shown in the figure were deposited by plasma spray process the most common current technique for applying such coatings.



Some of the innumerable design options with regard to TBC are

given below.

Fuel Flexibility Corrosion resistance Alternative fuels No derating for heavy fuels

Availability and Reliability Corrosion / Erosion resistance Lower metal temperature Lower transient thermal stress

Efficiency Reduce coolant flow Increase the turbine inlet temperature

Capital cost
Easily cast super alloy Simplified cooling

The options stressed above depend on the application of TBC. For aircraft turbines, emphasis has been placed on efficiency, durability and capital cost. For example, calculations have shown that the application of 1 mm oxide coating to the first two stages of an aircraft gas turbine can reduce cooling air consumption by 6.1% yielding a net thrust specific fuel consumption improvement of 1.3%. Alternatively, metal temperature and transient thermal stresses can be reduced significantly with more than a four-fold improvement in blade life.

For stationary gas turbines and diesel engines, emphasis has been placed on fuel flexibility and durability. In some cases, ceramics are more corrosion resistant than potential metallic coated materials, thus permitting firing with minimally

6 processed fuels. They also result in lower metal temperatures; improve creep and thermal fatigue resistance of the substrate metal.

For small aircraft and land vehicles on IC engines, efficiency improvement has been emphasized.



Nowadays, in the industrial production there are two different

processes for the manufacturing of thermal barrier coatings: the plasma spraying (PS) and the electron beam - physical vapour deposition (EB-PVD). Whilst EB-PVD process is usually used to deposit an aluminium layer as bond coat, the plasma spraying technology applies vacuum plasma-sprayed bond coats, which are mostly made of MCrAlY (M=Ni, Co). Recently, the high velocity oxy-fuel technology (HVOF) is also increasingly used for this task. The ceramic layer is usually deposited by means of the atmospheric plasma spraying (APS).

The most common is the Plasma Spray process. Fig 2 shows a plasma spray torch.

Fig 2. Plasma Spray Process

It consists of tungsten cathode and water-cooled copper anode. The gas used is Argon, Helium, Hydrogen and Nitrogen. Argon is usually chosen as the base gas due to its ionization characteristics and chemical inertness. An electric arc is initiated between the two electrodes using a high frequency discharge and then sustained using dc power. The arc ionizes the gas, creating high-pressure gas plasma. The resulting increase in gas temperature, which may exceed 30,000oC in turn, increases the volume of the gas and, hence, its pressure and velocity as it exits the nozzle. Power levels in plasma spray torches re usually in the range of 30 kW to 80 kW, but they can be as high as 120 kW.

8 The powder velocities usually range from 300 m/s to 550 m/s. Temperatures are usually at or slightly above the melting point. Powder is usually introduced in to the gas stream either just outside the torch or in the diverging exit region of the nozzle. Prior to coating deposition by plasma spraying the substrate is degreased and roughened by grit blasting to facilitate bonding with MCrAlY bond coat. Some of the important factors to be considered in plasma spray process are:

i. ii. iii. iv. v. vi.

The power settings of the plasma spray gun. The gun to substrate distance. The composition of plasma gas. The motion and orientation of the gun relative to the item being coated. The particle size distribution. The thermal properties and temperature of the substrate.

These factors control the temperature of the hot particles as they deposit on the surface, their impact velocity, the rates at which they cool, the time available for solid state reactions, and thus the structure which develops are some of the critical variables controlling coating performance.

Coating thickness usually ranges from about 0.05 mm to 0.5 mm but may be thicker for some applications. In the as deposited condition, plasma sprayed coatings are composed of many layer of splattered particles. MCrAlY bond coat alloys must have a surface roughness of the order of 7 micron to produce a good mechanical bond with the ceramic.

When sprayed under ambient conditions, the MCrAlY particles are partially oxidized. The resultant deposit has oxide stringers at splat boundaries and some porosity as illustrated in Fig 1. The amount of porosity and oxidation, however, can be reduced by inert gas shrouding and by the use of recently developed supersonic spray guns. Nearly 100% dense, oxide free MCrAlY coatings can be deposited by plasma spray in partial vacuum.

Micro cracks and pores are generally beneficial to ceramic coating durability since they provide a means for arresting cracks that tend to be generating during

9 thermo mechanical stress. Coating porosity also increases its insulating capability. In practice, porosity is typically maintained at about 10 to 15% of Zirconia (Zro2) coatings. Fig 3 shows the mechanism of life increase.




Fig 3. Mechanism of life increase

Plasma prayed coatings are applied adopting CNC machine tool. Here a microprocessor controlled plasma spray system has been coupled with an optical probe, which provides feed back measurements of coating thickness. This system has the capability to apply thermal barrier coatings to turbine blades with much uniformity and reproducibility than is possible manually.




Thermal barrier coatings have been extensively used to protect the internal surfaces of the combustion chambers in aircraft gas turbines. Due to their low absorptance and low thermal conductivity, such coatings yield a substantial reduction in metal temperature.

In turbines, the demands placed on TBC are far more stringent than in the combustor. The high convective heat fluxes encountered in the turbine results in large thermo mechanical stress in ceramic coatings under both transient and steady state conditions. The average and high metal temperature on the hot spots of the turbine range from 50 to 100oC higher than in the combustor. This places great demand on the environmental resistance of the bond coat. At these high surface temperatures, plasma sprayed coatings are subjected to process such as sintering and phase changes.

Earlier, thermal barrier coatings for turbine blade applications involved calcia and magnesia stabilized Zirconia and nichrome (Ni-20Cr) bond coats (all compositions are in weight percent). The first major advance was the identification of ZrO2-12Y2O3 / Ni-16Cr-6Al-0.6Y system. This new generation TBC offered superior durability compared to the oxide coat mentioned above. The bond coat operating temperature limit for this system is 850 to 900oC a significant improvement attributed to the improved environmental resistance of NiCrAlY.

Some of the tests performed on turbine blades of the research engines used in aircraft are given below. These tests were performed at National Aeronautical and Space Administration (NASA), USA.


A second milestone in TBC technology was the successful completion of an engine test. In this test, all but two of the turbine blades of a J-75 research engine first stage rotor were coated with ZrO2-12Y2O3 / Ni-16Cr-6Al-0.6Y coating systems. After 17 hours of operation between full power and flame out, all coated blades were in good condition. The results are given below;

11 At full power, engine conditions were 1370oC 3 atm 8300 1080oC 930oC

Inlet gas temperature Pressure RPM


Coating surface temperature Blade metal temperature -

At flame out condition 730oC 1 atm 3300 530oC

Inlet gas temperature Pressure RPM Blade metal temperature



In a Burner Rig instrument the temperature is monitored at the bond coat by 1/16” sheathed K-type thermocouple. The thermocouple is made of 0.127 mm chromel and alumel wires that are imbedded at various depths during spraying. Data acquisition was done with a high-speed analog to digital converter connected to a computer. The program records the maximum cycle temperature.

Fig 4 shows the result of Mach I burner rig test. The original system survived 850 hours at 1450oC. The new system was removed from the test without failure after 2000 hours at a slightly higher surface temperature. Another specimen was exposed to a much higher temperature of 1570oC and still survived 685 hours. It should be noted that these surface temperatures greatly exceed the melting point of the super alloys used for gas turbine blades and vanes. They also exceed the turbine inlet temperature of present gas turbines.

Additional results show that ceramic doubling the bond can approximately double coating life coat thickness. Increasing the bond coat Cr content from 16 to 25 w/o or increase the bond coat Al content from 6 to 10 w/o can also double life.

12 With these improved coatings, the bond coat operating temperature has been significantly increased to 950 – 1000oC range.

ZrO2-12 w/o Y2O3/NiCrAlY Substrate temp: 880oC, Surface temp: 1450oC

ZrO2-8 w/o Y2O3/NiCrAlY Substrate temp: 920oC, Surface temp: 1470oC

ZrO2-8 w/o Y2O3/NiCrAlY Substrate temp: 950oC, Surface temp: 1570oC






Fig 4. Mach I Burner Rig test results of ZrO2-Y2O3 / NiCrAlY TBC.


The durability of the ZrO2-12Y2O3 / Ni-16Cr-6Al-0.6Y coating system is greatly diminished when inorganic contaminants such as Na, V and S are present in gas turbine fuels. The Mach 0.3 burner rig test was conducted at NASA. The results are shown in Fig 5. The liquid salts like Na2SO4 (Sodium Sulphate) are believed to enter the coating where the temperature exceeds the melting point of the salt. Since Na 2SO4 does not react with ZrO2-Y2O3 under the experiment conditions, the impurity adversely affects the ability of the ceramic to accommodate cyclic thermal stresses.

In order to determine an improved thermal barrier coating, a series of coating systems were tested in 0.3 burner rig test with fuel impurity level of 5 ppm Na + 2ppm V. The coating thickness was maintained at 13 micron. The most promising duplex

13 coating was identified as 1.8CaO.SiO2 / Ni-16Cr-6Al-0.6Y which survived 600 hours of working before spalling.


0.2 ppm V

5 ppm Na + 2 ppm V

5 ppm Na

0.5 ppm Na




Fig 5. Effect of fuel impurities on ZrO2-12Y2O3 / Ni-16Cr-6Al-0.6Y TBC.



Barrier Coatings developed primarily for turbine

applications can also be used in diesel engines benefits derived from using this technology includes

i. ii. iii. iv. v. vi.

An improvement in the diesel engine fuel economy. Increase in engine power density. 40% heat reduction in in-cylinder heat rejection. Reduction in the ring groove temperatures. Reduction in the metal piston temperatures. Reduction in the lubricant temperatures

The coating selected for engine tests was a fully stabilized Zirconia coating consists of multiple ceramic metallic layers as shown in Fig 6.


1.524 mm


0.508 mm


0.508 mm 0.127 mm

Fig 6. Multilayer thermal barrier coating system

Plasma sprayed multiplayer TBC consisting of discrete layers comprised of a NiCrAlY band coat followed by 40/60 Zirconia /CoCrAlY, 85/15 Zirconia/CoCrAlY and with a fully stabilized 20% w/o yttria-zirconia bond coat were prepared for engine evaluation. After 14000 hours of service in a towboat engine operating

15 on inland waterways, the piston crowns and fire decks exhibited excellent durability and had performed smoothly. The thermal and mechanical efficiency of the engine had increased by the use of TBC. Calculations has shown that a coating less than 2.54 mm thick on pistons can withstand a pressure of1.03 MPa brake mean effective pressure without eroding.

The thermal structure analysis predicted performance improvement with following coatings system features.


Multilayered coatings were preferred over single layer coatings as they reduce strain discontinuities in the bond line region.


Decreased top layer coating thickness for reduced overall coating state of stress.


Prestressing of the coating can be used to control peak stresses particularly in the bond line region.




Thermal barrier coatings can also be applied on furnace components, heat-treating equipments, chemical processing equipments, heat exchangers rocket motor nozzles, exhaust manifolds, jet engine parts and nuclear power plant components. It has also been applied to uncoated blades used to drive the high-pressure hydrogen turbo pump of the Space Shuttle Main Engine. Here the coating is applied to provide thermal lag.




The main disadvantage is ceramic bond coat interface region which was determined to be the weakest link in the ZrO2-12Y2O3 / NiCrAlY system. Adhesive plus cohesive failure of the ceramic layer in this region occurred when the coating was subjected to a normal tensile load at room temperature as well as when applied on bars subjected to axial tension or compression at elevated temperature.



Some of the work related to Thermal Barrier Coating undergoing at

NASA-Lewis are


A study of solid-state reactions between the ceramic layer, the bond coat metal oxides, the bond coat and substrate.


A study to determine the kinetics of phase transformation in the ZrO212Y2O3 system and the effect of such transformations on mechanical properties.


Detailed analysis of coating stresses and controlled process, plasma spray technology has significantly improved the reliability of TBC turbines, diesel engines and other heat engines. Processing improvement in the control and development of TBC are required. Further study on, the mechanisms controlling coating adherence and degradation in clean and dirty environments, the effects of coating composition and structure on coating properties and correlation of models of engine tests are necessary to obtain thermal barrier coating that have even better tolerance to high temperature and thermo mechanical stresses.




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