Engine hot section component and method for making the same

One embodiment of the present invention is a unique engine hot section component having a coating system operative to reduce heat transfer to the hot section component. Another embodiment is a unique method for making a gas turbine engine hot section component with a coating system. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines, hot section components and coating systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

FIELD OF THE INVENTION

The present invention relates to engines, e.g., gas turbine and other engines, and more particularly, to an engine hot section component and method for making the same.

BACKGROUND

Engine hot section components and coating systems for engine hot section components remain an area of interest. Some existing systems have various shortcomings, drawbacks, and disadvantages relative to certain applications. Accordingly, there remains a need for further contributions in this area of technology.

SUMMARY

One embodiment of the present invention is a unique engine hot section component having a coating system operative to reduce heat transfer to the hot section component. Another embodiment is a unique method for making an engine hot section component with a coating system. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines, hot section components and coating systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

DETAILED DESCRIPTION

For purposes of promoting an understanding of the principles of the invention, reference will now be made to the embodiments illustrated in the drawings, and specific language will be used to describe the same. It will nonetheless be understood that no limitation of the scope of the invention is intended by the illustration and description of certain embodiments of the invention. In addition, any alterations and/or modifications of the illustrated and/or described embodiment(s) are contemplated as being within the scope of the present invention. Further, any other applications of the principles of the invention, as illustrated and/or described herein, as would normally occur to one skilled in the art to which the invention pertains, are contemplated as being within the scope of the present invention.

Referring now to the drawings, and in particularFIG. 1, a non-limiting example of an engine10in accordance with an embodiment of the present invention is depicted. In one form, engine10is an aircraft propulsion gas turbine engine. In other embodiments, engine10may be a land-based or marine engine. In one form, engine10is a multi-spool turbofan engine. In other embodiments, gas turbine engine10may be a single or multi-spool turbofan, turboshaft, turbojet, turboprop gas turbine or combined cycle engine. In still other embodiments, engine10may be ramjet engine, scramjet engine, pulse detonation engine and/or any engine having components exposed to high temperatures.

Gas turbine engine10includes a fan system12, a compressor system14, a diffuser16, a combustion system18and a turbine system20. Combustion system18is fluidly disposed between compressor system14and turbine system20. Fan system12includes a fan rotor system22. Compressor system14includes a compressor rotor system24. Turbine system20includes a turbine rotor system26. Turbine rotor system26is driving coupled to compressor rotor system24and fan rotor system22via a shafting system28. Combustion system18and turbine system20are considered hot sections of gas turbine engine10, and components of combustion system18and turbine system20are considered hot section components.

During the operation of gas turbine engine10, air is drawn into the inlet of fan12and pressurized by fan12. Some of the air pressurized by fan12is directed into compressor system14, and the balance is directed into a bypass duct (not shown). Compressor system14further pressurizes the air received from fan12, which is then discharged in to diffuser16. Diffuser16reduces the velocity of the pressurized air, and directs the diffused airflow into combustion system18. Fuel is mixed with the air in combustion system18, which is then combusted in a combustion liner (not shown). The hot gases exiting combustor18are directed into turbine system20, which extracts energy in the form of mechanical shaft power to drive fan system12and compressor system14shafting system28. The hot gases exiting turbine system20are directed into a nozzle (not shown), and provide a component of the thrust output by gas turbine engine10.

Referring now toFIG. 2, a non-limiting example of some aspects of combustion system18and turbine system20is schematically depicted. Combustion system18includes a combustion liner30. Turbine system20includes a plurality of turbine vanes32, and a plurality of turbine blades34operationally disposed within a plurality of blade tracks36. Disposed between vanes32and liner30are transition duct walls38and40. Transition duct walls38and40are operative to guide hot gases from combustion system18into turbine vanes32. During the operation of engine10, combustion liner30contains one or more combustion flames. Each of combustion liner30, turbine vanes32, turbine blades34, blade tracks36and transition duct walls38and40have surfaces that are in line-of-sight radiative communication with a combustion flame42during the operation of gas turbine engine10, resulting in radiative heat transfer from combustion flame42to those surfaces of combustion liner30, turbine vanes32, turbine blades34, blade tracks36and transition duct walls38and40during the operation of engine10. Each of combustion liner30, turbine vanes32, turbine blades34, blade tracks36and transition duct walls38and40have surfaces that are also exposed to radiation from combustion flame42via reflection.

In order to protect one or more surfaces of components that are exposed to high temperatures resulting from the combustion of fuel and air in combustion system18, e.g., combustion flame42, it is desirable to provide coatings on components so exposed. In accordance with embodiments of the present invention, the coatings include radiation barrier coatings. In other embodiments, the protective coatings also include thermal barrier coatings (TBC). In still other embodiments, one or more bond coats and adhesion aid coatings are layered between the protective coating layers, e.g., to increase coating durability. Embodiments of the present invention are applicable to gas turbine engine hot section components and hot section components of other engines, including stationary components, rotating components, translating components and reciprocating components.

Referring now toFIG. 3, a non-limiting example of a gas turbine hot section component50is depicted. In one form, component50is a combustion liner, such as combustion liner30. In other embodiments, gas turbine engine10may include a plurality of components50in the form of one or more of turbine vanes32, turbine blades34, blade tracks36and transition duct walls38and40. In still other embodiments, component50may be any hot section component that is exposed to hot gases from the combustion process that takes place in a combustion system, e.g., such as combustion system18, during the operation of an engine.

In the depicted embodiment, component50includes a substrate52and a coating system53. In one form, substrate52is the base material from which component50is formed. In other embodiments, substrate52may be any portion or material used in the construction of component50.

Coating system53is operative to reduce heat transfer to substrate52. In one form, coating system53is configured to reduce radiative heat transfer to substrate52. In one form, coating system53is configured to reduce radiative and convective heat transfer to component50. In other embodiments, coating system may also be configured to reduce conductive heat transfer to component50.

In one form, coating system53includes a bond coat54, a TBC coating56, an adhesion aid58and a radiation barrier coating system60. However, in an elemental form, component50may include only substrate52and radiation barrier coating system60. Embodiments may include, in addition to radiation barrier coating system60, other coatings in addition or in place of, such as bond coat54, TBC coating56and adhesion aid58, and may contain a greater or lesser number of coatings and coating types. In other embodiments, component50may include more layers than those illustrated and described herein or may employ less layers.

Bond coat54is configured to provide an adherent surface for adhering TBC coating56to substrate52. In one form, bond coat54is also configured as an oxidation and corrosion resistant layer to protect substrate52from environmental degradation in the presence of hot combustion gases. In other embodiments, bond coat54may be configured only or primarily for adhering TBC coating56to substrate52. In one form, bond coat54is a MCrAlY, wherein M may be Co, Ni or Co/Ni. In other embodiments, other bond coat materials may be employed, for example and without limitation, other aluminum rich layers operative to form protective alumina scale over substrate52, such as a low sulphur platinum aluminide. In one form, bond coat54is applied using large area filtered arc deposition (LAFAD) processing with ion assisted arc deposition (IAAD). In other embodiments, other processes may be employed to apply bond coat54to substrate52, e.g., directed vapor deposition (DVD), low pressure plasma spray (LPPS) and/or pack cementation. Some embodiments may not include bond coat54.

TBC coating56is operative to reduce heat transfer from the hot combustion gases supplied by combustion system18. In one form, TBC coating56is a yttria stabilized zirconia (YSZ) layer. In a particular form, TBC coating56is an 8% YSZ layer. In one form, TBC coating56is applied to component50, e.g., on top of bond coat54, using large area filtered arc deposition (LAFAD) processing with ion assisted arc deposition (IAAD). In other embodiments, other processes may be employed to apply TBC coating56to bond coat54, e.g., directed vapor deposition (DVD), air plasma spray (APS) or another thermal spray system and/or electron beam physical vapor deposition (EB-PVD). Some embodiments may not include TBC coating56.

In one form, TBC coating56is processed with a surface finish enhancing treatment that is operative to enhance the reflectivity of TBC coating56. In one form of surface finish enhancing treatment, TBC coating56is diamond polished. In other embodiments, other surface finish enhancing treatments may be employed. Some surface finish enhancing treatments may include, for example and without limitation, tumbling in a vibratory finishing machine, and laser finishing. In other embodiments, TBC coating56may not receive any surface finish treatment processes.

Adhesion aid58is operative to provide an adherent surface for adhering radiation barrier coating system60to component50, e.g., to TBC coating56, which in some embodiments increases the durability of the component50coating system. In one form, adhesion aid58is alumina. In other embodiments, other adhesion aid materials may be employed, for example, mullite, silicates and/or zircon. Some embodiments may not include adhesion aid58.

In one form, adhesion aid58is processed with a surface finish enhancing treatment to improve the reflectivity of adhesion aid58. In one form of surface finish enhancing treatment, adhesion aid58is diamond polished. In other embodiments, other surface finish enhancing treatments may be employed. Some surface finish enhancing treatments may include, for example and without limitation, tumbling in a vibratory finishing machine, and laser finishing. In other embodiments, adhesion aid58may not receive any surface finish treatment processes.

Radiation barrier coating system60is a multi-layered radiation barrier coating system formed of radiation barriers that are selected for have differing refraction indexes. The refraction indexes of each layer are selected to configure radiation barrier coating system60to reflect radiant energy from substrate52. In one form, the refraction indexes are selected to configure radiation barrier coating system60to reflect radiant energy from combustion flame42away from component50to reduce radiative heat transfer from combustion flame42to component50. Accordingly, in some embodiments radiation barrier coating system60includes materials that are selected to refract and reflect radiant energy in the wavelength range of 200-700 nm. The range of 200-700 nm was determined to be appropriate, including by testing. In other embodiments, radiant barrier coating system60may be configured to refract and reflect radiant energy at desired wavelengths within and/or without the range of 200-700 nm.

In one form, radiation barrier coating system60includes alternating high refractive index materials and low refractive index materials. By alternating high and low refractive index radiation barrier coatings, radiation barrier coating system60increases the reflection of radiant energy from component50relative to systems that do not so alternate high and low index of refraction coatings. As used herein with respect to refraction index, the terms “high” and “low” pertain to the refractive indexes of the coatings materials in the comparative sense, not in the absolute sense. The terms “high” and “low,” as used herein with respect to refraction indexes are not to be construed as limiting radiation barrier coating system60to any particular materials or indexes of refraction.

The radiation barrier coatings are selected for their refraction indexes, among other things, e.g., temperature capability, oxidation resistance, hot corrosion resistance. In particular, the radiation barrier coatings are selected for having different refraction indexes, e.g., as between adjacent radiation barrier coating layers. In one form, the index of refraction is based on the coating elemental composition. In some embodiments, the index of refraction is based not only on the material composition, but also based on how the composition is manufactured and processed, how the composition is applied to the component, and/or any processing after the composition is applied to the component. The radiation barrier coatings are configured to reflect radiant energy from combustion flame42away from substrate52, e.g., radiant energy at preselected wavelengths, such as the wavelengths of combustion flame42that would otherwise result in undesirable heat transfer to substrate52.

In one form, radiation barrier coating system60includes a plurality of radiation barrier coatings, depicted as radiation barrier coatings60A,60B,60C and60D. Radiation barrier coating60A is applied onto component50, e.g., onto adhesion aid58, and radiation barrier coating60B is applied over radiation barrier coating60B. Radiation barrier coating60C is applied onto radiation barrier coating60B, and radiation barrier coating60D is applied over radiation barrier coating60C. In other embodiments, radiation barrier coating system60may include only two radiation barrier layers, e.g., radiation barrier coating60A and radiation barrier coating60B. In still other embodiments, radiation barrier coating system60may include only three layers, e.g., radiation barrier coating60A, radiation barrier coating60B and radiation barrier coating60C. In yet still other embodiments, more than four of radiation barrier coatings may be employed.

In the embodiment illustrated inFIG. 3, there are no adhesion aids between the layers60A,60B,60C and60D of radiation barrier coating materials. In other embodiments, adhesion aids may be employed between some or all layers of radiation barrier coatings in order to aid the adhesion of one layer to the other. In still other embodiments, other coatings between two or more of radiation barrier coatings60A-60D and/or between any other coating layers deposited onto component50may be employed to reduce the adverse effects of differential thermal expansion between layers60A-60D and/or other coating layers.

Also, in the embodiment ofFIG. 3, the individual radiation barrier coating layers60A-60D do not receive any surface finish enhancing treatment. In other embodiments, surface finish enhancing treatments to improve the reflectivity of one or more or radiation barrier coatings60A-60D and/or of any adhesion aid disposed between radiation barrier coatings may be employed. Surface finish enhancing treatments may include, for example and without limitation, diamond polishing, tumbling in a vibratory finishing machine, and laser finishing.

Examples of materials for radiation barrier coatings60A and60B include aluminum oxide (Al2O3), mullite, SiO2, tantala, rutile (TiO2) and niobium oxide (Nb2O5). Other materials may be employed in addition to or in place of those mentioned herein. Of those listed above, tantala, rutile (TiO2) and Nb2O5have high indexes of refraction relative to the index of refraction of each of Al2O3, mullite and SiO2.

In one form, radiation barrier coating60A has a lower index of refraction than radiation barrier coating60B; radiation barrier coating60C has a lower index of refraction than radiation barrier coating60B, and radiation barrier coating60D has a higher index of refraction than radiation barrier coating60C. In other embodiments, other relative variations in index of refraction between the layers may be employed. For example, in some embodiments, radiation barrier coating60A has a higher index of refraction than radiation barrier coating60B.

In one form, radiation barrier coating60A is formed of Al2O3. In other embodiments, radiation barrier coating60A is made from mullite, SiO2and/or other material(s) that have a refraction index less than that of radiation barrier coating60B, in addition to or in place of Al2O3. In one form, radiation barrier coating60B is formed of rutile (TiO2). In other embodiments, radiation barrier coating60B is made from tantala, rutile (TiO2) and Nb2O5and/or other material(s) that have a refraction index greater than that of radiation barrier coating60A, in addition to or in place of rutile. In one form, radiation barrier coating60C is formed of Al2O3. In other embodiments, radiation barrier coating60C is made from mullite, SiO2and/or other material(s) that have a refraction index less than that of radiation barrier coating60B, in addition to or in place of Al2O3. In one form, radiation barrier coating60D is formed of rutile (TiO2). In other embodiments, radiation barrier coating60D is made from tantala, rutile (TiO2) and Nb2O5and/or other material(s) that have a refraction index greater than that of radiation barrier coating60C, in addition to or in place of rutile.

In one form, radiation barrier coatings60A-60D applied to component50, e.g., on top of bond coat54, using large area filtered arc deposition (LAFAD) processing with ion assisted arc deposition (IAAD). In other embodiments, other processes may be employed to apply radiation barrier coatings60A-60D, e.g., directed vapor deposition (DVD), air plasma spray (APS) or another thermal spray system and/or electron beam physical vapor deposition (EB-PVD).

Embodiments of the present invention include a method for manufacturing a gas turbine engine hot section component, comprising: selecting a first radiation barrier coating for a first refraction index; selecting a second radiation barrier coating for a second refraction index different from the first refraction index; applying the first radiation barrier coating onto the gas turbine engine hot section component; and applying the second radiation barrier coating over the first radiation barrier coating.

In a refinement, the method further comprises applying the first radiation barrier coating over a first application of the second radiation barrier coating.

In another refinement, the method further comprises applying the second radiation barrier coating over a second application of the first radiation barrier coating.

In yet another refinement, the second refraction index is greater than the first refraction index.

In still another refinement, a first radiation barrier coating material includes one or more of aluminum oxide (Al2O3), mullite and SiO2.

In yet still another refinement, a second radiation barrier coating material includes one or more of tantala, rutile (TiO2) and niobium oxide (Nb2O5).

In a further refinement, the method further comprises performing a surface finish enhancing treatment on the gas turbine engine hot section component prior to applying at least one of the first radiation barrier coating and the second radiation barrier coating.

Embodiments of the present invention include a method for manufacturing an engine hot section component, comprising: applying a thermal barrier coating (TBC) to the engine hot section component; applying a first radiation barrier coating onto the TBC; and applying a second radiation barrier coating material over the first radiation barrier coating, wherein the second radiation barrier coating material is selected for having a different index of refraction than the first radiation barrier coating.

In a refinement, the TBC includes Yttria-Stabilized Zirconia (YSZ).

In another refinement, the method further comprises applying a bond coat to the component prior to applying the TBC.

In yet another refinement, the bond coat is an MCrAlY bond coat, and where M=Co, Ni or Co/Ni.

In still another refinement, the method further comprises performing a surface finish treatment on the TBC prior to application of the first radiation barrier coating, wherein the surface finish treatment is operative to enhance reflective properties of the component.

In yet still another refinement, the method further comprise applying an adhesion aid to the TBC prior to application of the first radiation barrier coating.

In a further refinement, the adhesion aid includes one or more of alumina, mullite, silicates and zircon.

In a yet further refinement, the method further comprises performing a surface finish treatment on the adhesion aid prior to application of the first radiation barrier coating, wherein the surface finish treatment is operative to enhance reflective properties of the component.

In a still further refinement, a first radiation barrier coating material includes one or more of aluminum oxide (Al2O3), mullite and SiO2.

In a yet still further refinement, the second radiation barrier coating material includes one or more of tantala, rutile (TiO2) and niobium oxide (Nb2O5).

In an additional refinement, the first radiation barrier coating is applied using large area filtered arc deposition (LAFAD) processing with ion assisted arc deposition (IAAD).

In another additional refinement, the first radiation barrier coating is applied using directed vapor deposition (DVD) processing.

In yet another additional refinement, an inert gas is used to transport a first radiation barrier coating material to the component.

Embodiments of the present invention include an engine hot section component, comprising: a substrate having a surface in line-of-sight radiative communication with a combustion flame during operation of an engine; a first radiation barrier coating positioned between the substrate and the combustion flame; a second radiation barrier coating positioned between the first radiation barrier coating and the combustion flame, wherein the first radiation barrier coating has a first index of refraction; the second radiation barrier coating has a second index of refraction different from the first index of refraction, and wherein the first radiation barrier coating and the second radiation barrier coating are configured to reflect radiant energy at preselected wavelengths from the combustion flame away from the substrate.