Abstract:
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.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     The present application claims benefit of U.S. Provisional Patent Application No. 61/428,795, filed Dec. 30, 2010, entitled Engine Hot Section Component And Method For Making The Same, which is incorporated herein by reference. 
    
    
     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. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The description herein makes reference to the accompanying drawings wherein like reference numerals refer to like parts throughout the several views, and wherein: 
         FIG. 1  schematically illustrates some aspects of a non-limiting example of a gas turbine engine in accordance with an embodiment of the present invention. 
         FIG. 2  depicts some aspects of non-limiting examples of a combustion system and turbine system that may be used in conjunction with embodiments of the present invention. 
         FIG. 3  schematically illustrates some aspects of a non-limiting example of a hot section component in accordance with an embodiment of the present invention. 
     
    
    
     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 particular  FIG. 1 , a non-limiting example of an engine  10  in accordance with an embodiment of the present invention is depicted. In one form, engine  10  is an aircraft propulsion gas turbine engine. In other embodiments, engine  10  may be a land-based or marine engine. In one form, engine  10  is a multi-spool turbofan engine. In other embodiments, gas turbine engine  10  may be a single or multi-spool turbofan, turboshaft, turbojet, turboprop gas turbine or combined cycle engine. In still other embodiments, engine  10  may be ramjet engine, scramjet engine, pulse detonation engine and/or any engine having components exposed to high temperatures. 
     Gas turbine engine  10  includes a fan system  12 , a compressor system  14 , a diffuser  16 , a combustion system  18  and a turbine system  20 . Combustion system  18  is fluidly disposed between compressor system  14  and turbine system  20 . Fan system  12  includes a fan rotor system  22 . Compressor system  14  includes a compressor rotor system  24 . Turbine system  20  includes a turbine rotor system  26 . Turbine rotor system  26  is driving coupled to compressor rotor system  24  and fan rotor system  22  via a shafting system  28 . Combustion system  18  and turbine system  20  are considered hot sections of gas turbine engine  10 , and components of combustion system  18  and turbine system  20  are considered hot section components. 
     During the operation of gas turbine engine  10 , air is drawn into the inlet of fan  12  and pressurized by fan  12 . Some of the air pressurized by fan  12  is directed into compressor system  14 , and the balance is directed into a bypass duct (not shown). Compressor system  14  further pressurizes the air received from fan  12 , which is then discharged in to diffuser  16 . Diffuser  16  reduces the velocity of the pressurized air, and directs the diffused airflow into combustion system  18 . Fuel is mixed with the air in combustion system  18 , which is then combusted in a combustion liner (not shown). The hot gases exiting combustor  18  are directed into turbine system  20 , which extracts energy in the form of mechanical shaft power to drive fan system  12  and compressor system  14  shafting system  28 . The hot gases exiting turbine system  20  are directed into a nozzle (not shown), and provide a component of the thrust output by gas turbine engine  10 . 
     Referring now to  FIG. 2 , a non-limiting example of some aspects of combustion system  18  and turbine system  20  is schematically depicted. Combustion system  18  includes a combustion liner  30 . Turbine system  20  includes a plurality of turbine vanes  32 , and a plurality of turbine blades  34  operationally disposed within a plurality of blade tracks  36 . Disposed between vanes  32  and liner  30  are transition duct walls  38  and  40 . Transition duct walls  38  and  40  are operative to guide hot gases from combustion system  18  into turbine vanes  32 . During the operation of engine  10 , combustion liner  30  contains one or more combustion flames. Each of combustion liner  30 , turbine vanes  32 , turbine blades  34 , blade tracks  36  and transition duct walls  38  and  40  have surfaces that are in line-of-sight radiative communication with a combustion flame  42  during the operation of gas turbine engine  10 , resulting in radiative heat transfer from combustion flame  42  to those surfaces of combustion liner  30 , turbine vanes  32 , turbine blades  34 , blade tracks  36  and transition duct walls  38  and  40  during the operation of engine  10 . Each of combustion liner  30 , turbine vanes  32 , turbine blades  34 , blade tracks  36  and transition duct walls  38  and  40  have surfaces that are also exposed to radiation from combustion flame  42  via 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 system  18 , e.g., combustion flame  42 , 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 to  FIG. 3 , a non-limiting example of a gas turbine hot section component  50  is depicted. In one form, component  50  is a combustion liner, such as combustion liner  30 . In other embodiments, gas turbine engine  10  may include a plurality of components  50  in the form of one or more of turbine vanes  32 , turbine blades  34 , blade tracks  36  and transition duct walls  38  and  40 . In still other embodiments, component  50  may 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 system  18 , during the operation of an engine. 
     In the depicted embodiment, component  50  includes a substrate  52  and a coating system  53 . In one form, substrate  52  is the base material from which component  50  is formed. In other embodiments, substrate  52  may be any portion or material used in the construction of component  50 . 
     Coating system  53  is operative to reduce heat transfer to substrate  52 . In one form, coating system  53  is configured to reduce radiative heat transfer to substrate  52 . In one form, coating system  53  is configured to reduce radiative and convective heat transfer to component  50 . In other embodiments, coating system may also be configured to reduce conductive heat transfer to component  50 . 
     In one form, coating system  53  includes a bond coat  54 , a TBC coating  56 , an adhesion aid  58  and a radiation barrier coating system  60 . However, in an elemental form, component  50  may include only substrate  52  and radiation barrier coating system  60 . Embodiments may include, in addition to radiation barrier coating system  60 , other coatings in addition or in place of, such as bond coat  54 , TBC coating  56  and adhesion aid  58 , and may contain a greater or lesser number of coatings and coating types. In other embodiments, component  50  may include more layers than those illustrated and described herein or may employ less layers. 
     Bond coat  54  is configured to provide an adherent surface for adhering TBC coating  56  to substrate  52 . In one form, bond coat  54  is also configured as an oxidation and corrosion resistant layer to protect substrate  52  from environmental degradation in the presence of hot combustion gases. In other embodiments, bond coat  54  may be configured only or primarily for adhering TBC coating  56  to substrate  52 . In one form, bond coat  54  is 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 substrate  52 , such as a low sulphur platinum aluminide. In one form, bond coat  54  is 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 coat  54  to substrate  52 , e.g., directed vapor deposition (DVD), low pressure plasma spray (LPPS) and/or pack cementation. Some embodiments may not include bond coat  54 . 
     TBC coating  56  is operative to reduce heat transfer from the hot combustion gases supplied by combustion system  18 . In one form, TBC coating  56  is a yttria stabilized zirconia (YSZ) layer. In a particular form, TBC coating  56  is an 8% YSZ layer. In one form, TBC coating  56  is applied to component  50 , e.g., on top of bond coat  54 , 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 coating  56  to bond coat  54 , 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 coating  56 . 
     In one form, TBC coating  56  is processed with a surface finish enhancing treatment that is operative to enhance the reflectivity of TBC coating  56 . In one form of surface finish enhancing treatment, TBC coating  56  is 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 coating  56  may not receive any surface finish treatment processes. 
     Adhesion aid  58  is operative to provide an adherent surface for adhering radiation barrier coating system  60  to component  50 , e.g., to TBC coating  56 , which in some embodiments increases the durability of the component  50  coating system. In one form, adhesion aid  58  is alumina. In other embodiments, other adhesion aid materials may be employed, for example, mullite, silicates and/or zircon. Some embodiments may not include adhesion aid  58 . 
     In one form, adhesion aid  58  is processed with a surface finish enhancing treatment to improve the reflectivity of adhesion aid  58 . In one form of surface finish enhancing treatment, adhesion aid  58  is 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 aid  58  may not receive any surface finish treatment processes. 
     Radiation barrier coating system  60  is 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 system  60  to reflect radiant energy from substrate  52 . In one form, the refraction indexes are selected to configure radiation barrier coating system  60  to reflect radiant energy from combustion flame  42  away from component  50  to reduce radiative heat transfer from combustion flame  42  to component  50 . Accordingly, in some embodiments radiation barrier coating system  60  includes 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 system  60  may 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 system  60  includes alternating high refractive index materials and low refractive index materials. By alternating high and low refractive index radiation barrier coatings, radiation barrier coating system  60  increases the reflection of radiant energy from component  50  relative 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 system  60  to 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 flame  42  away from substrate  52 , e.g., radiant energy at preselected wavelengths, such as the wavelengths of combustion flame  42  that would otherwise result in undesirable heat transfer to substrate  52 . 
     In one form, radiation barrier coating system  60  includes a plurality of radiation barrier coatings, depicted as radiation barrier coatings  60 A,  60 B,  60 C and  60 D. Radiation barrier coating  60 A is applied onto component  50 , e.g., onto adhesion aid  58 , and radiation barrier coating  60 B is applied over radiation barrier coating  60 B. Radiation barrier coating  60 C is applied onto radiation barrier coating  60 B, and radiation barrier coating  60 D is applied over radiation barrier coating  60 C. In other embodiments, radiation barrier coating system  60  may include only two radiation barrier layers, e.g., radiation barrier coating  60 A and radiation barrier coating  60 B. In still other embodiments, radiation barrier coating system  60  may include only three layers, e.g., radiation barrier coating  60 A, radiation barrier coating  60 B and radiation barrier coating  60 C. In yet still other embodiments, more than four of radiation barrier coatings may be employed. 
     In the embodiment illustrated in  FIG. 3 , there are no adhesion aids between the layers  60 A,  60 B,  60 C and  60 D 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 coatings  60 A- 60 D and/or between any other coating layers deposited onto component  50  may be employed to reduce the adverse effects of differential thermal expansion between layers  60 A- 60 D and/or other coating layers. 
     Also, in the embodiment of  FIG. 3 , the individual radiation barrier coating layers  60 A- 60 D 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 coatings  60 A- 60 D 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 coatings  60 A and  60 B include aluminum oxide (Al 2 O 3 ), mullite, SiO 2 , tantala, rutile (TiO 2 ) and niobium oxide (Nb 2 O 5 ). Other materials may be employed in addition to or in place of those mentioned herein. Of those listed above, tantala, rutile (TiO 2 ) and Nb 2 O 5  have high indexes of refraction relative to the index of refraction of each of Al 2 O 3 , mullite and SiO 2 . 
     In one form, radiation barrier coating  60 A has a lower index of refraction than radiation barrier coating  60 B; radiation barrier coating  60 C has a lower index of refraction than radiation barrier coating  60 B, and radiation barrier coating  60 D has a higher index of refraction than radiation barrier coating  60 C. In other embodiments, other relative variations in index of refraction between the layers may be employed. For example, in some embodiments, radiation barrier coating  60 A has a higher index of refraction than radiation barrier coating  60 B. 
     In one form, radiation barrier coating  60 A is formed of Al 2 O 3 . In other embodiments, radiation barrier coating  60 A is made from mullite, SiO 2  and/or other material(s) that have a refraction index less than that of radiation barrier coating  60 B, in addition to or in place of Al 2 O 3 . In one form, radiation barrier coating  60 B is formed of rutile (TiO 2 ). In other embodiments, radiation barrier coating  60 B is made from tantala, rutile (TiO 2 ) and Nb 2 O 5  and/or other material(s) that have a refraction index greater than that of radiation barrier coating  60 A, in addition to or in place of rutile. In one form, radiation barrier coating  60 C is formed of Al 2 O 3 . In other embodiments, radiation barrier coating  60 C is made from mullite, SiO 2  and/or other material(s) that have a refraction index less than that of radiation barrier coating  60 B, in addition to or in place of Al 2 O 3 . In one form, radiation barrier coating  60 D is formed of rutile (TiO 2 ). In other embodiments, radiation barrier coating  60 D is made from tantala, rutile (TiO 2 ) and Nb 2 O 5  and/or other material(s) that have a refraction index greater than that of radiation barrier coating  60 C, in addition to or in place of rutile. 
     In one form, radiation barrier coatings  60 A- 60 D applied to component  50 , e.g., on top of bond coat  54 , 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 coatings  60 A- 60 D, 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 (Al 2 O 3 ), mullite and SiO 2 . 
     In yet still another refinement, a second radiation barrier coating material includes one or more of tantala, rutile (TiO 2 ) and niobium oxide (Nb 2 O 5 ). 
     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 (Al 2 O 3 ), mullite and SiO 2 . 
     In a yet still further refinement, the second radiation barrier coating material includes one or more of tantala, rutile (TiO 2 ) and niobium oxide (Nb 2 O 5 ). 
     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. 
     While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment(s), but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims, which scope is to be accorded the broadest interpretation so as to encompass all such modifications and equivalent structures as permitted under the law. Furthermore it should be understood that while the use of the word preferable, preferably, or preferred in the description above indicates that feature so described may be more desirable, it nonetheless may not be necessary and any embodiment lacking the same may be contemplated as within the scope of the invention, that scope being defined by the claims that follow. In reading the claims it is intended that when words such as “a,” “an,” “at least one” and “at least a portion” are used, there is no intention to limit the claim to only one item unless specifically stated to the contrary in the claim. Further, when the language “at least a portion” and/or “a portion” is used the item may include a portion and/or the entire item unless specifically stated to the contrary.