Thermal and Environmental Barrier Coating Development for Advanced Propulsion Engine Systems

Ceramic thermal and environmental barrier coatings (TEBCs) are used in gas turbine engines to protect engine hot-section components in the harsh combustion environments, and extend component lifetimes. Advanced TEBCs that have significantly lower thermal conductivity, better thermal stability and higher toughness than current coatings will be beneficial for future low emission and high performance propulsion engine systems. In this paper, ceramic coating design and testing considerations will be described for turbine engine high temperature and high-heat-flux applications. Thermal barrier coatings for metallic turbine airfoils and thermal/environmental barrier coatings for SiC/SiC ceramic matrix composite (CMC) components for future supersonic aircraft propulsion engines will be emphasized. Further coating capability and durability improvements for the engine hot-section component applications can be expected by utilizing advanced modeling and design tools

Ceramic thermal barrier coatings for electric utility gas turbine engines

Research and development into thermal barrier coatings for electric utility gas turbine engines is reviewed critically. The type of coating systems developed for aircraft applications are found to be preferred for clear fuel electric utility applications. These coating systems consists of a layer of plasma sprayed zirconia-yttria ceramic over a layer of MCrAly bond coat. They are not recommended for use when molten salts are presented. Efforts to understand coating degradation in dirty environments and to develop corrosion resistant thermal barrier coatings are discussed

Microstructure Based Material-Sand Particulate Interactions and Assessment of Coatings for High Temperature Turbine Blades

Gas turbine engines for military/commercial fixed-wing and rotary wing aircraft use thermal barrier coatings in the high-temperature sections of the engine for improved efficiency and power. The desire to further make improvements in gas turbine engine efficiency and high power-density is driving the research and development of thermal barrier coatings and the effort of improving their tolerance to fine foreign particulates that may be contained in the intake air. Both commercial and military aircraft engines often are required to operate over sandy regions such as in the Middle-East nations, as well as over volcanic zones. For rotorcraft gas turbine engines, the sand ingestion is adverse during take-off, hovering near ground, and landing conditions. Although, most of the rotorcraft gas turbine engines are fitted with inlet particle separators, they are not 100 percent efficient in filtering fine sand particles of size 75 microns or below. The presence of these fine solid particles in the working fluid medium has an adverse effect on the durability of turbine blade thermal barrier coatings and overall performance of the engine. Typical turbine blade damages include blade coating wear, sand glazing, Calcia-Magnesia-Alumina-Silicate (CMAS) attack, oxidation, plugged cooling holes, all of which can cause rapid performance deterioration including loss of aircraft. The objective of this research is to understand the fine particle interactions with typical ceramic coatings of turbine blades at the microstructure level. A finite-element based microstructure modeling and analysis has been performed to investigate particle-surface interactions, and restitution characteristics. Experimentally, a set of tailored thermal barrier coatings and surface treatments were down-selected through hot burner rig tests and then applied to first stage nozzle vanes of the Gas Generator Turbine of a typical rotorcraft gas turbine engine. Laser Doppler velocity measurements were performed during hot burner rig testing to determine sand particle incoming velocities and their rebound characteristics upon impact on coated material targets. Further, engine sand ingestion tests were carried out to test the CMAS tolerance of the coated nozzle vanes. The findings from this on-going collaborative research to develop the next-gen sand tolerant coatings for turbine blades are presented in this paper

Thermal barrier coatings for gas turbine and diesel engines

The present state of development of thin thermal barrier coatings for aircraft gas turbine engines and thick thermal barrier coatings for truck diesel engines is assessed. Although current thermal barrier coatings are flying in certain gas turbine engines, additional advances will be needed for future engines. Thick thermal barrier coatings for truck diesel engines have advanced to the point where they are being seriously considered for the next generation of engine. Since coatings for truck engines is a young field of inquiry, continued research and development efforts will be required to help bring this technology to commercialization

A Review of the Erosion of Thermal Barrier Coatings.

The application of thermal barrier coatings (TBCs) to components with internal cooling in the hot gas stream of gas turbine engines has facilitated a steep increase in the turbine entry temperature and the associated increase in performance and efficiency of gas turbine engines. However, TBCs are susceptible to various life limiting issues associated with their operating environment including erosion, corrosion, oxidation, sintering and foreign object damage (FOD). This is a review paper that examines various degradation and erosion mechanisms of TBCs, especially those produced by electron beam physical vapour deposition (EB-PVD). The results from a number of laboratory tests under various impact conditions are discussed before the different erosion and FOD mechanisms are reviewed. The transitions between the various erosion mechanisms are discussed in terms of the D/d ratio (contact area diameter/column diameter), a relatively new concept that relates the impact size to the erosion mechanism. The effects of ageing, dopant additions and calciumâ  magnesiumâ  aluminaâ  silicates on the life of TBCs are examined. It is shown that while ageing increases the erosion rate of EB-PVD TBCs, ageing of plasma sprayed TBCs in fact lowers the erosion rate. Finally modelling of EB-PVD TBCs is briefly intr

Summary of NASA research on thermal-barrier coatings

A durable, two-layer, plasma-sprayed coating consisting of a ceramic layer over a metallic layer was developed that has the potential of insulating hot engine parts and thereby reducing metal temperatures and coolant flow requirements and/or permitting use of less costly and complex cooling configurations and materials. The investigations evaluated the reflective and insulative capability, microstructure, and durability of several coating materials on flat metal specimens, a combustor liner, and turbine vanes and blades. In addition, the effect on the aerodynamic performance of a coated turbine vane was measured. The tests were conducted in furnaces, cascades, hot-gas rigs, an engine combustor, and a research turbojet engine. Summaries of current research related to the coating and potential applications for the coating are included

Erosion, corrosion and erosion-corrosion of EB PVD thermal barrier coatings

Electron beam (EB) physical vapour deposited (PVD) thermal barrier coatings (TBCs) have been used in gas turbine engines for a number of years. The primary mode of failure is attributed to oxidation of the bond coat and growth of the thermally grown oxide (TGO), the alumina scale that forms on the bond coat and to which the ceramic top coat adheres. Once the TGO reaches a critical thickness, the TBC tends to spall and expose the underlying substrate to the hot gases. Erosion is commonly accepted as a secondary failure mechanism, which thins the TBC thus reducing its insulation capability and increasing the TGO growth rate. In severe conditions, erosion can completely remove the TBC over time, again resulting in the exposure of the substrate, typically Ni-based superalloys. Since engine efficiency is related to turbine entry temperature (TET), there is a constant driving force to increase this temperature. With this drive for higher TETs comes corrosion problems for the yttria stabilised zirconia (YSZ) ceramic topcoat. YSZ is susceptible to attack from molten calciumâ  magnesiumâ  aluminaâ  silicates (CMAS) which degrades the YSZ both chemically and micro-structurally. CMAS has a melting point of around 1240 à °C and since it is common in atmospheric dust it is easily deposited onto gas turbine blades. If the CMAS then melts and penetrates into the ceramic, the life of the TBC can be significantly reduced. This paper discusses the various failure mechanisms associated with the erosion, corrosion and erosionâ  corrosion of EB PVD TBCs. The concept of a dimensionless ratio D/d, where D is the contact footprint diameter and d is the column diameter, as a means of determining the erosion mechanism is introduced and discussed for E

Further industrial tests of ceramic thermal barrier coatings

The NASA Lewis Research Center made technical assistance arrangements (contracts) with several commercial organizations under which Lewis designed plasma-sprayed thermal-barrier coatings (TBC) for their products. Lewis was then furnished with the test conditions and evaluations of coating usefulness. The coating systems were developed and sprayed at Lewis. All of the systems incorporated a two-layer, ceramic-bond coating concept. Coating thickness and chemical composition were varied to fit three applications: the leading edges of first-stage turbine vanes for an advanced gas turbine engine; the flame impingement surfaces of a combustor transition section; and diesel engine valves and head surfaces. The TBC incorporated yytria-stabilized zirconia, which lowered metal temperatures, protected metal parts, and increased metal part life. In some cases metal burning, melting, and warping were eliminated. Additional benefits were realized from these endeavors: hands-on experience with thermal-barrier coatings was provided to industry; the success of these endeavors encourages these and other organizations to accelerate the implementation of TBC technology

Potential benefits of a ceramic thermal barrier coating on large power generation gas turbine

Thermal barrier coating design option offers benefit in terms of reduced electricity costs when used in utility gas turbines. Options considered include: increased firing temperature, increased component life, reduced cooling air requirements, and increased corrosion resistance (resulting in increased tolerance for dirty fuels). Performance and cost data were obtained. Simple, recuperated and combined cycle applications were considered, and distillate and residual fuels were assumed. The results indicate that thermal barrier coatings could produce large electricity cost savings if these coatings permit turbine operation with residual fuels at distillate-rated firing temperatures. The results also show that increased turbine inlet temperature can result in substantial savings in fuel and capital costs

Advanced aircraft engine materials trends

Recent activities of the Lewis Research Center are reviewed which are directed toward developing materials for rotating hot section components for aircraft gas turbines. Turbine blade materials activities are directed at increasing metal temperatures approximately 100 C compared to current directionally solidified alloys by use of oxide dispersion strengthening or tungsten alloy wire reinforcement of nickel or iron base superalloys. The application of thermal barrier coatings offers a promise of increasing gas temperatures an additional 100 C with current cooling technology. For turbine disk alloys, activities are directed toward reducing the cost of turbine disks by 50 percent through near net shape fabrication of prealloyed powders as well as towards improved performance. In addition, advanced alloy concepts and fabrication methods for dual alloy disks are being studied as having potential for improving the life of future high performance disks and reducing the amount of strategic materials required in these components