Patent Abstract:
A high temperature gas turbine component for use in the gas flow path that also is a specular optical reflector. A thin layer of a high temperature reflector is applied to the gas flow path of the component, that is, the surface of the component that forms a boundary for hot combustion gases. The component typically includes a thermal barrier coating overlying the high temperature metallic component that permits the component to operate at elevated temperatures. The thermal barrier coating must be polished in order to provide a surface that can suitably reflect the radiation into the gas flow path. A thin layer of the high temperature reflector the is applied over the polished thermal barrier coating by a process that can adequately adhere the reflector to the polished surface without increasing the roughness of the surface. The high temperature reflector can be applied to any surface aft of the compressor, such as on a turbine nozzle. The surface reflects radiation back into the hot gas flow path. The reflected radiation is not focused onto any other hardware component. The design of the component is such that the radiation is returned to the gas flow path rather than absorbed into a component wall that only serves to increase the temperature of the wall.

Full Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
   This Application is related to application Ser. No. 10/335,657, filed contemporaneously with this Application on Dec. 31, 2002, entitled “IMPROVED HIGH TEMPERATURE SPLASH PLATE FOR TEMPERATURE REDUCTION BY OPTICAL REFLECTION AND PROCESS FOR MANUFACTURING” assigned to the assignee of the present invention and which is incorporated herein by reference, to application Ser. No. 10/335,647, filed contemporaneously with this Application on Dec. 31, 2002, entitled “IMPROVED HIGH TEMPERATURE CENTERBODY FOR TEMPERATURE REDUCTION BY OPTICAL REFLECTION AND PROCESS FOR MANUFACTURING” assigned to the assignee of the present invention and which is incorporated herein by reference, and to application Ser. No. 10/335,442, filed contemporaneously with this Application on Dec. 31, 2002, entitled “IMPROVED HIGH TEMPERATURE COMBUSTOR WALL FOR TEMPERATURE REDUCTION BY OPTICAL REFLECTION AND PROCESS FOR MANUFACTURING” assigned to the assignee of the present invention and which is incorporated herein by reference. 

   FIELD OF THE INVENTION 
   The present invention is directed to gas turbine engines, and in particular, to modifications of components of such engines to reduce the temperature of boundary walls of the hot section portions of the components by optical radiation generated by combustion. 
   BACKGROUND OF THE INVENTION 
   In the compressor portion of an aircraft gas turbine engine, atmospheric air is compressed to 10-25 times atmospheric pressure, and adiabatically heated to about 800°-1250° F. (425°-675° C.) in the process. This heated and compressed air is directed into a combustor, where it is mixed with fuel. The fuel is ignited, and the combustion process heats the gases to very high temperatures, in excess of 3000° F. (1650° C.). These hot gases pass through the turbine, where rotating turbine wheels extract energy to drive the fan and compressor of the engine, and the exhaust system, where the gases supply thrust to propel the aircraft. To improve the efficiency of operation of the aircraft engine, combustion temperatures have been raised. Of course, as the combustion temperature is raised, steps must be taken to prevent thermal degradation of the materials forming the flow path for these hot gases of combustion. 
   Every aircraft gas turbine engine has a so-called High Pressure Turbine (HPT) to drive its compressor. The HPT sits just behind the compressor in the engine layout and experiences the highest temperature and pressure levels (nominally 2400° F. and 300 psia respectively) developed in the engine. The HPT also operates at very high speeds (10,000 RPM for large turbofans, 50,000 for small helicopter engines). In order to meet life requirements at these levels of temperature and pressure, today&#39;s HPT components are always air-cooled and constructed from advanced alloys. 
   While a straight turbojet engine will usually have only one turbine (an HPT), most engines today are of the turbofan, either high bypass turbofan or low bypass turbofan, or turboprop type and require one (and sometimes two) additional turbine(s) to drive a fan or a gearbox. The additional turbines are called the Low Pressure Turbines (LPT) and immediately follows the HPT in the engine layout. Since substantial pressure drop occurs across the HPT, the LPT operates with a much less energetic fluid and will usually require several stages (usually up to six) to extract the available power. 
   One well-known solution that has been undertaken to protect the metals that form the flow path for the hot gases of combustion, including those of the HPT and LPT, have included application of protective layers having low thermal conductivity. These materials are applied as thermal barrier coating systems (TBCs), typically comprising a bond coat that improves adhesion of an overlying ceramic top coat, typically a stabilized zirconia, to the substrate. These systems are known to improve the thermal performance of the underlying metals that form the flow path in the hot section of the engine. However, as temperatures of combustion have increased, even these TBCs have been found to be insufficient. 
   Another solution that has been used in conjunction with TBCs is active cooling of metal parts. Initially, active cooling provided a flow of air from the compressor to the back side of metal parts comprising the flow gas path. As temperatures increased even further, serpentine passageways were formed in the metallic components and cooling air was circulated through the parts to provide additional cooling capability, the cooling air exiting through apertures positioned in the gas flow side of the component, providing an additional impingement film layer along the gas flow path. This method is referred to as film cooling. Even though the air from the compressor is adiabatically heated to perhaps as high as 1250° F. (675° C.), the compressor air is still significantly cooler than the combustion gases moving along the gas flow path of the engine. However, as the temperatures of the combustion process have continued to increase, even these tried and true methods of cooling are reaching their limitations. In particular, the turbine nozzles of high efficiency, advanced cycle turbine engines are prone to failure as a result of thermal degradation. 
   While some modifications of the traditional flow path surfaces have been applied in the past, such as the application of materials over the TBC, these modifications have been directed to reducing the emissions of pollutants such as unburned hydrocarbons (UHC) and carbon monoxide (CO). One such modification is set forth in U.S. Pat. No. 5,355,668 to Weil, et al., assigned to the assignee of the present invention, which teaches the application of a catalyst such as platinum, nickel oxide, chromium oxide or cobalt oxide directly over the flow path surface of the thermal barrier coating of a component such as a turbine nozzle. The catalyst layer is applied to selected portions of flow path surfaces to catalyze combustion of fuel. The catalytic material is chosen to reduce air pollutants such as unburned hydrocarbons (UHC) and carbon monoxide (CO) resulting from the combustion process. The catalytic layer is applied to a thickness of 0.001 to 0.010 inches and is somewhat rough and porous, having a surface roughness of about 100 to 250 micro inches, in order to enhance the surface area available to maximize contact with the hot gases in order to promote the catalytic reaction. The rough surface assists in creating some turbulence that promotes contact with the catalytic surface. 
   The prior art solutions are either directed to problems that are unrelated to the problem of high temperature experienced by turbine nozzles, such as the Weil patent, or provide different solutions to the problem of high temperatures resulting from the combustion process. The present invention provides a different approach to the problem of high temperatures experienced by turbine nozzles. 
   SUMMARY OF THE INVENTION 
   The present invention is a high temperature gas turbine component for use in the gas flow path that also is a specular optical reflector. The gas turbine component is positioned in the hot section of the engine, behind the compressor section and reflects heat radiation, for example infrared radiation having wavelengths in the range of about 1 micron to about 10 microns, from the combustor region back into the hot gas flow path. The reflected radiation is focused away from any other hardware component in the combustor region so that the radiative heat passes into the turbine portion of the engine. The design of the component is such that the radiation is returned to the gas flow path rather than absorbed into a component wall which only serves to increase the temperature of the wall. 
   A thin layer of a high temperature reflector metal is applied to the flow path surface of the component, that is, the surface of the component that forms a boundary for hot combustion gases. The high temperature reflector must be applied as an optically smooth coating. The component typically includes a thermal barrier coating overlying the high temperature metallic component that permits the component to operate at elevated temperatures. The thermal barrier coating (TBC) applied to the component typically is rough and must be polished in order to provide a sufficiently smooth surface that can suitably reflect the radiation into the gas flow path. A thin layer of the high temperature reflector then is applied by a process that can adequately adhere the reflector to the polished TBC surface without increasing the roughness of the surface. The high temperature reflector can be applied to any surface aft of the compressor, but is most beneficially used in the combustor portion of the engine, for instance, the combustor wall, and the high pressure turbine portion of the engine. For military aircraft, the high temperature reflector metal would also be beneficially used in the augmentor portion of the engine. 
   An advantage of the present invention is that the radiation from the combustion process is reflected back into the gas flow path. This radiative heat, rather than being absorbed by the component in the combustor or HPT portion of the engine, is absorbed by the fluid and carried back into portions of the engine further aft that currently operate at cooler temperatures. The result is that the component does not become as hot. At a given temperature of operation of the engine, the component, because it is operating at a cooler temperature, will not deteriorate as rapidly due to thermal degradation, resulting in longer component life and less mean time between repair or refurbishment. 
   Another advantage of the present invention is that the fluid stream will be heated to a higher temperature as the reflected radiation is absorbed by the materials comprising the gaseous fluid and carried from the combustor portion of the engine into the aft turbine portions of the engine. This increased fluid temperature translates into increased engine efficiency, as the available energy in the fluid stream for both extraction by the turbine to operate the engine and for thrust to propel the aircraft is greater. 
   Still another advantage of the present invention is that the engine can be operated at an even higher temperature than currently experienced using the current invention if shortened component life and increased repair rates can be tolerated in exchange for even greater efficiency. 
   Other features and advantages of the present invention will be apparent from the following more detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings which illustrate, by way of example, the principles of the invention. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
       FIG. 1  is a schematic representation of a high bypass turbofan gas turbine engine. 
       FIG. 2  is a schematic representation of a low bypass turbofan gas turbine engine equipped with an augmentor. 
       FIG. 3  is a schematic representation of the combustor and high pressure turbine sections of a gas turbine engine. 
       FIG. 4  is a cross-section of an as-manufactured high pressure turbine vane of a gas turbine engine after application of a conventional thermal barrier system. 
       FIG. 5  is a cross-section of the high pressure turbine vane of the gas turbine engine after the outer surface of the ceramic topcoat has been smoothed to achieve a surface finish of 50 micro inches or finer; and 
       FIG. 6  is schematic representation of the optical reflector of the present invention applied over a smooth ceramic topcoat. 
   

   DETAILED DESCRIPTION OF THE INVENTION 
   In accordance with the present invention, hot section components of a gas turbine engine which form the boundary of the gas flow path or which are located in the gas flow path are coated with a thin layer of a specular optical reflective material that has a high temperature capability. The material as applied has a smooth surface finish so as to reflect the heat back into the fluid path and away from other hot section components. 
   A high bypass aircraft gas turbine engine  10  is shown schematically in FIG.  1 . During operation, air is forced through the fan  12 . A portion of the air bypasses the core of the engine and is used to contribute to the thrust that propels the engine. A portion of the air is compressed in the booster  14  and compressor  16  portions of the engine up to 10-25 times atmospheric pressure, and adiabatically heated to 800°-1250° F. in the process. This heated and compressed air is directed into the combustor portion of the engine  18 , where it is mixed with fuel supplied through a fuel nozzle system  20 . The fuel is ignited, and the combustion process heats the gases to temperatures on the order of 3200°-3400° F. These hot gases pass through the high pressure  22  and low pressure  24  turbines, where rotating discs extract energy to drive the fan and compressor of the engine. The gases then are passed to the exhaust system  26 , where they contribute to thrust for aircraft propulsion. 
   Operation of a low bypass gas turbine engine, shown schematically at  30  in  FIG. 2 , is similar, except that operational requirements may dictate omission of the booster  14  and addition of an augmentor  28  in the exhaust system shown at  26  in FIG.  1 . To emphasize the conceptual similarity, the same identification numerals are employed in both figures. 
   The combustor  18  and high pressure turbine  22  sections of an engine such as in  FIG. 1  or  FIG. 2  are shown in greater detail in FIG.  3 . Compressed air from the compressor is introduced through a diffuser  40  into an annular cavity defined by outer combustor case  42  and the inner combustor case  44 . A portion of the compressed air passes through a swirl nozzle  46 , where it is mixed with fuel supplied through a fuel tube  48 . The swirl nozzle and fuel tube are components of the fuel nozzle system  20 . The fuel/air mixture is self-igniting under normal operating conditions, except for those transient conditions where flame instability or flame-out occurs. The flame is confined and directed toward the turbine by the outer combustor liner  50  and the inner combustor liner  52 . These liners are oriented about a central axis  55  and are substantially symmetrical about this central axis  55  forming the gas flow path. Each combustor liner additionally is provided with a plurality of cooling holes  54 , through which compressed air supplied by the compressor is forced to pass. Compressed air in annulus between liners  50 ,  52  and combustor cases  42 ,  44  provides back side cooling to liners  50 ,  52  before exiting cooling holes  54 . The combustor liners  50  and  52  are described as having an inner side, and an outer side. 
   The hot gases of combustion then leave the combustor and enter the high pressure turbine  22 , which may comprise a single stage, as shown in  FIG. 3 , or multiple stages, each stage being comprised of a nozzle  60  and a rotor  70 . For the purposes of this discussion, the combustor is presumed to be of a single stage configuration, for simplicity of discussion, but the concepts of the present invention are fully applicable to gas turbines of other configurations and designs with additional turbine stages. The nozzle  60  is comprised of a plurality of vanes  62  disposed between and secured to an inner band  64  and an outer band  66 . Vanes  62  are substantially stationary, although they may be capable of rotating about their axes in limited motion in variable guide vane configurations. The turbine nozzle  60  preferably includes a plurality of circumferentially adjoining segments  80  collectively forming a complete 360° assembly. Each segment  80  has two or more circumferentially spaced vanes  62  (one shown in FIG.  3 ), each having an upstream leading edge and a downstream trailing edge, over which the combustion gases flow. As the temperature of the hot gas in the gas flow path can easily exceed the melting point of the materials forming the boundaries of the gas flow path, it is necessary to cool the components forming the flow path, first by passing the air coming from the compressor at about 1000°-1250° F. (535°-675° C.) over the outer surfaces of the nozzles, then by using the same air after it passes through the cooling holes  102  (shown in  FIGS. 4 ,  5 , and  6 ) to direct a thin film of air between the surface of the vanes  62  and the hot gases. The thin film of air forming a boundary layer assists in protecting the vanes  62  from being heated to even higher temperatures by a process referred to as film cooling. Components of at least one turbine stage are often provided with cooling air through cooling holes. Additionally, the surface of the vanes  62  are also coated with thermal barrier coating systems, which are comprised of a bond coat applied between an underlying superalloy base material and an overlying ceramic layer, to create a thermal barrier coating system that reduces the flow of heat to the substrate material. 
   The rotor  70  is comprised of a plurality of blades, each having an airfoil section  72  and a platform  74 , which are securely attached to the periphery of a rotating disk  78 . Important associated structures to support the rotor are not shown. The blades cooperate with a stationary shroud  76  to effect a gas seal between rotor  70  and the stationary components of the engine. 
   Downstream of the fuel nozzle  46 , the gas flow path is defined by the inner surfaces of the inner combustor liner  52  and the outer combustor liner  50 , and portions of the turbine or turbines including the inner and outer bands  64  and  66 , the vanes  62 , which direct the flow of gas, the airfoil  72 , which extracts energy from the hot gas, the shrouds  76 , as well as the exhaust system  26  and/or augmentor  28  aft or downstream of the turbine section of the engine. The present invention is specifically applicable to those components which define the gas flow path downstream of the swirl nozzle  46 . Systems for providing cooling air and thermal barrier coating systems are well-known in the gas turbine engine art. 
   Materials employed in the combustor, turbine and exhaust system sections of aircraft gas turbines are typically high temperature superalloys based on nickel, cobalt, iron or combinations thereof. All of these superalloys are believed to be suitable substrate materials for the present invention. Also, monolithic ceramic materials and fiber reinforced ceramic matrix composite materials, described herein collectively as ceramic materials, may be employed in the combustor, turbine and exhaust systems sections of an aircraft gas turbine. Such ceramic materials are specifically contemplated for use in the present invention, and may have higher temperature limits than the high temperature superalloys used for combustors. 
   Even for gas turbine engines designed for commercial airliners, gas velocity through the engine may approach the speed of sound. Thus, the total gas residence time in the engine is but a small fraction of a second, during which time air coming through the compressor is mixed with liquid fuel, and combustion of the mixture occur. As the mixture is combusted to form a gas, heat, including radiant heat, is generated. Even with the most recent advances in cooling measures used in gas turbine engines such as active cooling controls and advanced thermal barrier coating systems which reduce the amount and/or rate of heat transferred to components due to convective and conductive heat transfer, the temperatures of the components along the flow path surface are still elevated to very high temperatures. The present invention assists in reducing the amount of heat transferred to these components by radiation transfer. 
   The present invention utilizes a high temperature specular optical reflector applied directly over existing ceramic materials such as thermal barrier systems utilized to protect the substrate material. These specular optical reflectors are applied as a very thin coating and in a manner so that they do not adversely affect the cooling holes in the surfaces of the components along the gas flow path. Conventional and well known techniques for applying thermal barrier coatings provide surfaces that are much too rough for the thin coatings to act as optical reflectors. When these specular reflectors are applied over conventional thermal barrier coatings having surface finishes of 100 micro inches and greater, the rough surface causes the radiation to be scattered in multiple of different directions and are ineffective in transferring heat back into the rapidly moving fluid. When the coatings are porous, such as when used for as a catalytic coating, the radiation is reabsorbed into the substrate, so it cannot be used as an optical reflector. 
   In one embodiment of the present invention, the specular optical coating of the present invention is applied to the surface of an afterburner liner. In another embodiment of the present invention, the specular optical coating of the present invention is applied to the surface of a flameholder. 
   In another embodiment of the present invention, a turbine nozzle is manufactured in accordance with standard manufacturing methods. A standard ray-tracing program could be used to optimize the geometry of the turbine nozzle to be coated with the specular optical coating of the present invention. In addition, there may be surfaces that have heat reflected onto them from reflection or refraction from neighboring turbine blades that could also benefit from the specular optical coating. Referring to  FIG. 4 , turbine vane  62  is comprised of a substrate  110  suitable for use at high temperatures. As discussed above, the substrate can be selected from several materials. However, as illustrated in  FIG. 4 , substrate  110  is a high temperature nickel base superalloy. A bond coat  112  is applied over the nickel base alloy substrate. Overlying bond coat  112  is a ceramic layer  114  having a surface  115  that has a rough surface finish. As used herein, the term “rough surface finish” is one that is greater than about 100 micro inches. When the substrate is selected from one of the available different materials, such as a ceramic matrix composite material, the bond coat  112  may be omitted. 
   The surface finish of the thermal barrier coating system is typically too rough to act as a specular optical reflector because of the manufacturing techniques used to apply the ceramic top coat. Given the complex geometry of the turbine nozzle and vane, the forward facing surface of the upstream leading edge of the turbine vane, that is, the outer surface of the thermal barrier coating overlying the substrate surface, is then polished. In one embodiment, the forward facing surface of the upstream leading edge of the turbine vane is polished by hand using fine emery paper so that the surface  115  of the ceramic layer  114 , as shown in  FIG. 5 , has a surface finish of no greater than about 50 micro inches, preferably about 32 micro inches and smoother. It is undesirable to polish the other surfaces of the turbine nozzle or to coat the other surfaces of the turbine nozzle with the coating of the present invention as such polishing and coating would only serve to reflect the heat of the gas flow path into a nearby vane or wall. This smooth surface of the forward facing surface of the turbine nozzle leading edge is required to achieve the reflective properties required for the present invention to be effective. Additionally, the smooth surface assists in maintaining a smooth laminar-like flow of the cooling layer adjacent to the surface of the component by minimizing turbulence. In production, well known polishing techniques such as lapper wheels with diamond paste and tumbling can be employed to speed the polishing process and increase throughput. 
   Next, the forward facing surface of the upstream leading edge of the turbine nozzle vane is coated with a very thin specular reflective coating  116  of a material, as shown in  FIG. 6 , that will reflect the radiation away from the surface. A standard ray-tracing program could be used to optimize the geometry of the turbine nozzle vane to be coated with the specular optical coating of the present invention. In addition, certain surfaces of the turbine nozzle vanes have heat reflected onto them from reflection or refraction from neighboring turbine blades that could also benefit from the specular optical coating. The coating  116  is applied by a process that deposits material so that a very smooth surface finish is maintained. A preferred method is a chemical vapor deposition (CVD) process that deposits a coating to a thickness of about 1 micron (0.0004″). Other acceptable methods for depositing this thin specular coating to a thickness of about 1 micron include sputtering, liquid phase infiltration and physical vapor deposition. However, not all methods for depositing a coating produce coatings consistent with this invention. Other methods such as thermal spray methods do not produce an acceptable coating for specular reflection, as the coatings deposited by these processes are too thick and too rough. The thickness of the specular layer can be greater, for example 10 microns or less, but is maintained at about 1 micron because of the great expense of the material used as the specular reflector. 
   A preferred specular reflector coating material is platinum, although palladium or multiple dielectric mirrors comprising tantalum oxide (Ta 2 O 5 ), silica (SiO 2 ), titanium dioxide (TiO 2 ) and combinations thereof. It is fundamental that the material used as a coating material remain highly reflective as the hot gas stream  120  passes over the surface. Thus, thick non-adherent oxide scales cannot form, as the formation of these scales destroy the effectiveness of the coating as a reflector. Also, the very thin coating, in addition to being less expensive, is extremely adherent to the polished TBC, and, due to its thinness, does not peel off in layers, which peeling can adversely affect the surface finish. The thin layer does not provide a severe weight penalty for the components to which it is added. In addition, the layer is maintained as a thin layer to allow the surface finish to be of high reflective, optical quality. 
   Testing of other reflective combustor components has indicated that a specular reflective layer can reflect at least about 80% of the incident radiation, an amount of radiation sufficient to lower the temperature of a component by up to about 100° F. when the temperature of a ceramic coating adjacent to the fluid stream is at 2300° F. (1260° C.) as compared to a component having a ceramic coating but without the specular reflective layer. These components have displayed an improvement of 95° F., as measured by thermocouples attached to deflectors in a high pressure sector test for approximately 100 hours, as compared to a substantially identical deflector that lacked a coating such as described by the present invention. 
   While the present invention has been described as an improvement to a turbine nozzle, the present invention can be applied to any other surface along the gas flow path of a turbine engine or other high temperature devices, such as a continuous furnace or a burner. For example, the specular reflective coating can be applied to the combustor walls, so that any incident radiation is reflected away from the combustor walls and into the gas flow path. Because at least a portion of the energy is reflected from the components comprising the gas flow path, thereby lowering their temperature, the radiation is absorbed by the gases in the gas flow path, thereby raising its temperature. 
   While the invention has been described with reference to a preferred embodiment, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this invention, but that the invention will include all embodiments falling within the scope of the appended claims.

Technology Classification (CPC): 5