Abstract:
A hybrid cooling system for a gas turbine engine includes a vapor cooling assembly and a cooling air cooling assembly. The cooling air cooling assembly is configured to remove thermal energy from cooling air used to cool a first component of the gas turbine engine. The vapor cooling assembly configured to transport thermal energy from a vaporization section to a condenser section through cyclical evaporation and condensation of a working medium sealed within the vapor cooling assembly. The vaporization section is located at least partially within a second component of the gas turbine engine, and the condenser section is located outside the second component.

Description:
BACKGROUND 
     The present invention relates to a hybrid system for cooling structures of gas turbine engines using a combination of vapor cooling and air cooling. 
     Known gas turbine engines have utilized superalloys, thermal barrier coatings (TBCs), and fluidic cooling schemes in order to provide engine structures that can operate efficiently at high temperatures and pressures while still maintaining a relatively long lifespan. Furthermore, “cooled” cooling air systems have been developed that reject thermal energy from air that is then used to provide cooling to various gas turbine engine components. However, the ability to provide cooled cooling air in a volume and with adequately low thermal energy to provide cooling to all of the static and rotating components of a gas turbine engine would be extremely demanding on the cooled cooling air systems, making suitable cooled cooling air systems undesirably large, heavy and complex. Therefore, it is desired to provide improved cooling capabilities for gas turbine engines, in order to better maintain engine components at temperatures below designated maximum operating temperature levels. 
     SUMMARY 
     A hybrid cooling system for a gas turbine engine includes a vapor cooling assembly and a cooling air cooling assembly. The cooling air cooling assembly is configured to remove thermal energy from cooling air used to cool a first component of the gas turbine engine. The vapor cooling assembly configured to transport thermal energy from a vaporization section to a condenser section through cyclical evaporation and condensation of a working medium sealed within the vapor cooling assembly. The vaporization section is located at least partially within a second component of the gas turbine engine, and the condenser section is located outside the second component. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a block diagram of a gas turbine engine having a hybrid cooling system according to the present invention. 
         FIGS. 2 and 3  are schematic cross-sectional views of portions of the gas turbine engine. 
     
    
    
     DETAILED DESCRIPTION 
     In general, the present invention relates to a hybrid cooling system for a gas turbine engine that utilizes both a “cooled” cooling air subsystem and a vapor cooling subsystem in order to help maintain engine components at temperatures below designated maximum operating temperature levels. Such a hybrid cooling system allows the size and complexity of the cooled cooling air subsystem to be offset by use of the vapor cooling subsystem. Cooled cooling air can be provided, for example, primarily to rotatable engine components, while the vapor cooling subsystem can be used to primarily cool static (i.e., non-rotating) engine components. Such a hybrid cooling system can substantially reduce cooling air expenditure and overall cooling system weight, each of which helps provide engine efficiency gains and cycle power increases. 
     As used herein, the term “static” as applied to gas turbine engine components generally refers to non-rotating components, although such components may be subject to some movement, for instance, when installed in an engine of a movable vehicle. 
       FIG. 1  is a block diagram of a gas turbine engine  10  that includes rotatable engine components  12  (e.g., turbine blades, rotors, etc.) and static engine components  14  (e.g., vanes, shroud rings, etc.). The representation of the gas turbine engine  10  in  FIG. 1  is simplified for clarity, but those of ordinary skill in the art will recognize that the present invention can be applied to essentially any type of gas turbine engine. 
     In the illustrated embodiment, a “cooled” cooling air assembly  16  provides cooled cooling air to the rotatable engine components  12 . The cooled cooling air assembly  16  includes a heat exchanger  18  and a fuel stabilization unit (FSU)  20 . The heat exchanger  18  accepts bleed air (e.g., compressor bleed air) from a bleed air source  22  and transfers thermal energy from the bleed air to a liquid fuel from a fuel supply  24  via the FSU  20 . The heat exchanger  18  provides a means to transfer thermal energy between two fluids while maintaining physical separation of those fluids. Here, the liquid fuel passing through the heat exchanger  18  acts as a heat sink to accept thermal energy from the gaseous bleed air. The bleed air passing through the heat exchanger  18  is cooled, in order to produce the cooled cooling air that is then routed to the rotatable engine components  12  to provide a desired cooling effect. Cooled cooling air that is heated as a result of cooling the rotatable engine components  12  can then be rejected as exhaust.The heat exchanger  18  can be of a conventional configuration, or can be configured as disclosed in commonly-assigned U.S. Patent Application Publication No. 2008/0142189, entitled “Vapor Cooled Heat Exchanger,” filed Dec. 19, 2006, which is hereby incorporated by reference in its entirety. 
     The fuel that accepts thermal energy from the bleed air in the heat exchanger  18  has limits as to how much heat can be accepted before that fuel degenerates or auto-ignites. Therefore, the FSU  20  is provided, which help prevent fuel degeneration. In general, the FSU  20  can act as a fuel deoxygenator to reduce oxygen concentration in the fuel, which can reduce undesired “coking” effects, and thereby allows the engine  10  to operate with elevated fuel temperatures. The FSU  20  can have a known configuration. Fuel leaving the FSU  20  is ultimately routed to a combustor assembly (not shown in  FIG. 1 ) where it is burned to power the engine  10 . 
     The static engine components  14  are cooled using a vapor cooling assembly  26 . In general, the vapor cooling assembly is configured to transport thermal energy from a vaporization section to a condenser section at a relatively high rate through cyclical evaporation and condensation of a working medium sealed within the vapor cooling assembly  26 . Thermal energy can be transferred from the vapor cooling assembly  26  to a gaseous heat sink  28 , for example, relatively cool air in a fan bypass stream. The gases of gaseous heat sink  28  are ultimately ejected from the engine  10  as exhaust. 
       FIG. 2  is a schematic cross-sectional view of a portion of the gas turbine engine  10 , showing one embodiment of the cooled cooling air assembly  16 . As shown in  FIG. 2 , the gas turbine engine  10  includes a compressor section  30 , a combustor assembly  32  and a turbine section  34 , all arranged relative to an engine centerline C L . In the illustrated embodiment, two heat exchangers  18 A and  18 B provide low-pressure and high-pressure cooling circuits, respectively, both of which are operatively connected to the FSU  20 . Low-pressure bleed air  36  is routed to the heat exchanger  18 A from the compressor section  30 , and high pressure bleed air  38  is routed to the heat exchanger  18 B from the combustor assembly  32  (e.g., from a plenum surrounding an annular combustor liner  40 ). Furthermore, low-pressure cooled cooling air  42  is routed from the heat exchanger  18 A to desired areas (e.g., areas that are radially outside of a primary gas flowpath of the engine  10 ), and high-pressure cooled cooling air  44  is routed from the heat exchanger  18 B to other desired areas (e.g., areas that are radially inward from a primary gas flowpath of the engine  10 ). Ultimately, the high-pressure cooled cooling air  44  and/or the low-pressure cooled cooling air  42  can be used to cool various components, such as the rotating components including turbine blades and rotors in the turbine section  34 . It should be noted that the high-pressure cooled cooling air  44  would generally not be exposed to the primary gas flowpath of the engine  10 , but, as necessary for particular applications, would be routed through suitable conduits to radially cross the primary flowpath (typically at locations upstream from the combustor assembly  32 ). The low-pressure cooling circuit of associated with the low-pressure bleed air  36  and the low-pressure cooled cooling air  42  can be isolated from the high-pressure cooling circuit associated with the high-pressure bleed air  38  and the high-pressure cooled cooling air  44 , such that gases in those two circuits do not mix or interact. In the illustrated embodiment, the heat exchangers  18 A and  18 B are separate units, although in alternative embodiments, a single heat exchanger unit can be utilized that is configured to maintain separation between the low and high-pressure cooling circuits. 
       FIG. 3  is a schematic cross-sectional view of another portion of the gas turbine engine  10 . As shown in  FIG. 3 , the gas turbine engine  10  includes a vane  46  extending into a primary flowpath  48 , a fan bypass duct  50 , and a vapor cooling assembly  26 . In the illustrated embodiment, the vapor cooling assembly  26  includes a vaporization section that extends into the vane  46  and a condenser section  54  that is exposed to airflow in the fan bypass duct  50 . An optional flow guide  38  positioned in the fan bypass duct  50  functions to direct air in the fan bypass duct  50  toward and past the condenser section  54  of the vapor cooling assembly  26 , and can then direct air heated by the condenser section  56  back to the fan bypass flowpath. 
     The vapor cooling assembly  26  functions as a heat pipe that uses an evaporative cooling cycle to transfer thermal energy through the evaporation and condensation of a working medium, such as disclosed in commonly-assigned U.S. patent application Ser. No. 11/654,472, entitled “Vapor Cooled Static Turbine Hardware,” filed Jan. 17, 2007 and commonly-assigned U.S. patent application Ser. No. 11/642,010, entitled “Vapor Cooling of Detonation Engines,” filed Dec. 19, 2006, which are both hereby incorporated by reference in their entireties. In general, the vapor cooling assembly  26  utilizes an evaporative cooling cycle to transfer thermal energy from the vane  46  to air passing through the fan bypass duct  50 . Thermal energy absorbed by the vane  46  from the hot gases in the combustion gas flowpath  48  heats the vaporization section  52 , which causes the working medium in the vaporization section  52  to evaporate. Moreover, the relatively cool air in the fan bypass duct  50  absorbs thermal energy from the condenser section  54 , and causes the vaporized working medium to condense. The working medium physically moves between the vaporization section  52  and the condenser section  54 , in order to transfer thermal energy between the locations where evaporation and condensation occur. The composition of the working medium used in the vapor cooling assembly  26  is selected according to the particular operating conditions at which heat transfer is desired. Thermal energy added to air in the fan bypass duct  50  raises the temperature and pressure of that air, which contributes to thrust output of the engine  10  and lessens energy loss due to the vapor cooling assembly  26 . 
     In traditional gas turbine engine cooling systems using cooling air to cool both static and rotating engine components, approximately twice as much cooling air (by volume) goes toward cooling static components as toward cooling rotating components. The present invention allows cooling air expenditures for static components to be reduced or eliminated, thereby allowing approximately two-thirds savings in cooling air expenditure over prior art systems. Furthermore, by decreasing cooling air expenditure, the present invention allows the use of significantly smaller-volume cooled cooling air assemblies than would otherwise be needed, allowing the overall weight of a hybrid cooling system of the present invention to be approximately half the weight of cooling systems that would utilize only cooling air to cool both static and rotatable components of an engine. For example, smaller volume cooled cooling air assemblies can utilize smaller heat exchangers, which can greatly contribute to weight reductions. These features of the present invention help provide engine efficiency gains and cycle power increases. 
     Although the present invention has been described with reference to preferred embodiments, workers skilled in the art will recognize that changes may be made in form and detail without departing from the spirit and scope of the invention. For instance, the rotatable components can be cooled with both air cooling and vapor cooling as disclosed in commonly assigned U.S. patent application Ser. No. 11/542,097, entitled “Hybrid Vapor and Film Cooled Turbine Blade,” filed Oct. 3, 2006, which is hereby incorporated by reference in its entirety. Moreover, the particular manner in which cooled cooling air is routed through an engine, and the configuration of vapor cooling assemblies can vary as desired for particular applications.