Abstract:
One embodiment of the present invention is a unique gas turbine engine. Another embodiment is a unique cooling system for a gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for cooling one or more objects of cooling. 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 the benefit of U.S. Provisional Patent Application 61/291,540, filed Dec. 31, 2009, and is incorporated herein by reference. 
    
    
     GOVERNMENT RIGHTS 
     The present application was made with United States government support under Contract No. FA8650-07-C-2803 awarded by the United States Air Force. The United States government may have certain rights in the present application. 
    
    
     FIELD OF THE INVENTION 
     The present invention relates to gas turbine engines, and more particularly, to a cooling system for use in a gas turbine engine. 
     BACKGROUND 
     Cooling systems that effectively cool objects of cooling, such as fluids or devices in a gas turbine engine, 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 gas turbine engine. Another embodiment is a unique cooling system for a gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for cooling one or more objects of cooling. 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 depicts a non-limiting example of a gas turbine engine with a cooling system in accordance with an embodiment of the present invention. 
         FIG. 2  depicts a non-limiting example of a cooling system in accordance with an embodiment of the present invention. 
         FIG. 3  schematically depicts an end view of the cooling system of  FIG. 2 . 
     
    
    
     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 a gas turbine engine  10  in accordance with an embodiment of the present invention is schematically depicted. In one form, gas turbine engine  10  is an aircraft propulsion power plant in the form of a turbofan engine. In various embodiments, engine  10  may be any gas turbine engine configuration. Gas turbine engine  10  includes a fan  12 , a fan  14 , a compressor system  16 , a diffuser  18 , a combustor  20 , a turbine system with a high pressure (HP) turbine  22  and a low pressure (LP) turbine  24 , an exhaust nozzle system  26 , a bypass duct  28 , a bypass duct  30 , and a cooling system  32 . 
     Each of fan  12  and fan  14  include a plurality of fan blades that pressurize air received at the fan inlet. In one form, fan  12  includes a single stage of circumferentially spaced blades and a single stage of circumferentially spaced vanes. In other embodiments, fan  12  may not include vanes, or may include multiple stages of both blades and vanes. Likewise, in one form, fan  14  includes a single stage of circumferentially spaced blades and a single stage of circumferentially spaced vanes. In other embodiments, fan  14  may not include vanes, or may include multiple stages of both blades and vanes. In one form, gas turbine engine  10  includes a flow control system  34  to direct some of the pressurized air discharged from fan  12  into fan  14  and some of the pressurized air into bypass duct  30 . In some embodiments, flow control system  34  is configured to vary the amount of flow as between fan  14  and bypass duct  30 . In some embodiments, flow control system  34  may be an active means of directing flow, e.g., controlled by a control system (not shown). In other embodiments, flow control system  34  may be passive, e.g., controlled based on pressures and/or temperatures in one or more regions of engine  10 , or may be fixed. In still other embodiments, gas turbine engine  10  may not include a flow control system such as flow control system  34 . 
     Compressor system  16  includes a plurality of blades and vanes for compressing air. In one form, compressor system  16  is a single compressor having a plurality of stages of blades and vanes driven by a common shaft at a common speed. In other embodiments, compressor system  16  may include a plurality of compressors operating at the same or different speeds, each of which includes one or more stages of blades, and each of which may also include a desirable number of vane stages. For example, in some forms, compressor system  16  may include an LP compressor and/or an intermediate pressure compressor and/or an HP compressor. In one form, gas turbine engine  10  includes a flow control system  36  to direct some of the pressurized air discharged from fan  14  into compressor system  16  and some of the pressurized air into bypass duct  28 . In some embodiments, flow control system  36  is configured to vary the amount of flow as between compressor system  16  and bypass duct  28 . In some embodiments, flow control system  36  may be an active means of directing flow, e.g., controlled by a control system (not shown). In other embodiments, flow control system  36  may be passive, e.g., controlled by pressures and/or temperatures in one or more regions of engine  10 , or may be fixed. In still other embodiments, gas turbine engine  10  may not include a flow control system such as flow control system  36 . 
     Diffuser  18  and combustor  20  are fluidly disposed between compressor system  16  and HP turbine  22 . Compressor system  16 , diffuser  18 , combustor  20 , HP turbine  22  and LP turbine  24  form an engine core. HP turbine  22  and LP turbine  24  extract power from the airflow exiting combustor  20 . LP turbine  24  is drivingly coupled to fan  12  via an LP shaft  38 . HP turbine  22  is drivingly coupled to compressor system  16  via an HP shaft  40 . Compressor system  16 , HP shaft  40  and HP turbine  22  form, in part, an HP spool. Fan  12 , LP shaft  38  and LP turbine  24  form, in part, an LP spool. In one form, fan  14  is driven by LP turbine  24 , which may be a direct coupling via LP shaft  38  in some embodiments. In other embodiments, fan  14  may be coupled to LP turbine  24  via a system that allows fan  14  to operate at a different speed than LP turbine  24 , e.g., a fixed speed ratio or a variable ratio gear train. In still other embodiments, fan  14  may be powered by HP turbine  22 . 
     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 fan  14  by flow control system  34 , and the balance is directed into bypass duct  30 . Bypass duct  30  channels the pressurized air to exhaust nozzle system  26 , which provides a component of the thrust output by gas turbine engine  10 . The air directed into fan  14  is further pressurized by fan  14 . Some of the air pressurized by fan  14  is directed into compressor system  16  by flow control system  36 , and the balance is directed into bypass duct  28 . Bypass duct  28  channels the pressurized air to exhaust nozzle system  26 , which provides a component of the thrust output by gas turbine engine  10 . Exhaust nozzle system  26  is operative to control the pressure of the air streams in exhaust nozzle system  26 , including balancing pressures as between bypass duct  28  and bypass duct  30 . 
     Compressor system  16  receives the pressurized air from fan  14 , which is compressed and discharged in to diffuser  18 . Diffuser  18  diffuses the core flow that is discharged from compressor system  16 , reducing its velocity and increasing its static pressure. The diffused airflow is directed into combustor  20 . Fuel is mixed with the air in combustor  20 , which is then combusted in a combustion liner (not shown). The hot gases exiting combustor  20  are directed into HP turbine  22 , which extracts energy from the hot gases in the form of mechanical shaft power to drive compressor system  16  via HP shaft  40 . The hot gases exiting HP turbine  22  are directed into LP turbine  24 , which extracts energy in the form of mechanical shaft power to drive fan  12  and fan  14  via LP shaft  38 . The hot gases exiting LP turbine  24  are directed into nozzle  26 , and provide a component of the thrust output by gas turbine engine  10 . 
     The airflow that passes through compressor system  16  and subsequently into combustor  20  is referred to herein as core flow. The pressurized airflow exiting fan  14  and received into bypass duct  28  is referred to herein as a second stream flow; and the pressurized airflow exiting fan  14  and received into bypass duct  30  is referred to herein as a third stream flow. Each of the core flow, second stream flow and third stream flow are working fluid streams. Working fluid in the context of the present application is the air that is compressed (pressurized) and/or expanded in one or more of the engine&#39;s fan, compressor and turbine stages to produce the thrust output by the engine. 
     Cooling system  32  is in fluid communication with compressor system  16 . Cooling system  32  is operative to receive core airflow from compressor system  16 , and extract heat therefrom to reduce the temperature of the received core airflow. In one form, the airflow that is received by cooling system  32  is a small portion of the core airflow, and is returned to the engine  10  core for use in cooling turbine blades and vanes of HP turbine  22 . In another form, the core air that is received by cooling system  32  represents a larger amount of air, e.g., substantially all of the core airflow in some embodiments, and is returned to compressor system  16  for additional compression. For example, in such embodiments, cooling system  32  may serve as an intercooler system. In still other embodiments, cooling system  32  may be employed to reduce the temperature of other objects of cooling. An object of cooling, as used herein, is the fluid, whether in liquid or gas form, and/or component and/or system that is sought to be cooled. For example, in other forms, objects of cooling may be one or more of hydraulic fluid and/or related systems, electrical and/or electronic circuits and/or systems, mechanical components and/or systems, and/or other components and/or systems, such as refrigeration components and/or systems. 
     Referring now to  FIG. 2 , cooling system  32  is schematically depicted. Cooling system  32  includes a heat exchanger  42  having a heat exchanger core  44 , a cooling medium inlet  46  for heat exchanger  42 , and a cooling medium outlet  48  for heat exchanger  42 . In one form, heat exchanger  42  is located in bypass duct  28 , although other locations are contemplated herein. 
     Heat exchanger  42  structured to remove heat from an object of cooling. In one form, heat exchanger  42  is operative to cool core airflow, in which case cooling system  32  includes a plurality of passages  50 ,  52 . Passages  50 ,  52  are structured to conduct core airflow to and from heat exchanger  42 . For example, where cooling system  32  is an intercooler system, core airflow from compressor system  16  may be directed to heat exchanger  42  via passage  50 , and then returned to compressor system  16  via passage  52 . In one form, passages  50 ,  52  include pipes that deliver core airflow to and from heat exchanger  42 . In other embodiments, other types of passages may be employed in addition to and/or in place of pipes. In other forms, cooling system  32  may not include passages  50 ,  52 , e.g., where the object of cooling is an electronic component. In one form, heat exchanger  42  is a parallel flow heat exchanger. In other embodiments, other heat exchanger types may be employed, e.g., counter flow heat exchangers, cross flow heat exchangers and/or mixed flow heat exchangers. 
     Cooling medium inlet  46  is in fluid communication with heat exchanger core  44  and bypass duct  28 . In one form, the air pressure in bypass duct  28  is greater than the air pressure in bypass duct  30 . Cooling medium inlet  46  is structured to receive a cooling medium in the form of air from a working fluid stream of gas turbine engine  10 , namely, from the second stream flow in bypass duct  28 . In one form, cooling medium inlet  46  receives a portion of the second stream flow from bypass duct  28 . In other embodiments, cooling medium inlet  46  may receive all or substantially all of the second stream flow. Cooling medium outlet  48  is in fluid communication with heat exchanger core  44  and bypass duct  30 . Cooling medium outlet  48  is structured to discharge the cooling medium to a third working fluid stream of gas turbine engine  10 , namely from the third stream flow in bypass duct  30 . Heat exchanger  42  removes heat from the object of cooling, e.g., core airflow, using the cooling medium (pressurized air) received from fan bypass duct  28  through cooling medium inlet  46 . The pressurized air is discharged into fan bypass duct  30  through cooling medium outlet  48 . 
     In one form, the pressure differential between the second stream flow and the third stream flow governs the flow rate of the cooling medium through heat exchanger  42 . In the depicted embodiments, the pressure differential is determined by the operation of exhaust nozzle system  26 , although other systems may determine the pressure differential in other embodiments. 
     It will be understood that more than one such cooling system  32  may be employed. For example, referring now to  FIG. 3 , in one form cooling system  32  may include heat exchangers, inlets and outlets that are distributed circumferentially and/or axially around gas turbine engine  10 , e.g., in bypass duct  28 , bypass duct  30  and/or other locations. It is also contemplated that in some forms, heat exchanger  42  may extend circumferentially around engine  10 , e.g., around bypass duct  28 , bypass duct  30 , and/or inboard of bypass duct  28  and/or outboard of bypass duct  30 . In other embodiments, only a single cooling system  32  having a single heat exchanger may be employed. 
     In other embodiments, the air pressure in bypass duct  30  may be greater than the air pressure in bypass duct  28 , in which case cooling medium inlet  46  is positioned to receive the cooling medium from bypass duct  30 , and cooling medium outlet  48  is positioned to discharge the cooling medium into bypass duct  28 . It is also contemplated that in some embodiments, the direction of pressure differential between bypass duct  28  and bypass duct  30  may alternate. In some such embodiments, cooling system  32  may employ valves or other flow directing systems to maintain flow through heat exchanger core  44  in a given direction. In other such embodiments, heat exchanger  42  may be configured to operate under flow reversing conditions. 
     Embodiments of the present invention include a cooling system for use with a gas turbine engine. The cooling system includes a heat exchanger operative to cool air of a first working fluid stream of the gas turbine engine; a cooling medium inlet for said heat exchanger structured to receive a cooling medium from a second working fluid stream of the gas turbine engine; and a cooling medium outlet for said heat exchanger structured to discharge the cooling medium to a third working fluid stream of the gas turbine engine. The cooling system is structured to transfer heat from the air of the first working fluid stream to the cooling medium and discharges the cooling medium into the third working fluid stream. 
     In a refinement, a pressure differential between the second working fluid stream and the third working fluid stream determines the flow rate of the air received from the second working fluid stream through the heat exchanger. 
     In another refinement, the cooling system includes a passage structured to conduct air of the first working fluid stream to the heat exchanger. In one form, the passage includes a pipe. 
     In yet another refinement, the heat exchanger is structured to cool only a portion of the air of the first working fluid stream. 
     In still another refinement, the first working fluid stream is a core air flow of the gas turbine engine. In a further refinement, the third working fluid stream is the output a fan of the gas turbine engine, and the second working fluid stream is the output of another fan of the gas turbine engine. 
     Embodiments of the present invention also include a cooling system for use with a gas turbine engine. The cooling system includes a heat exchanger structured to remove heat from an object of cooling; a heat exchanger inlet in fluid communication with a first fan bypass duct of the gas turbine engine; and a heat exchanger outlet in fluid communication with a second fan bypass duct of the gas turbine engine. The heat exchanger removes heat from the object of cooling using pressurized air received from the first fan bypass duct through the heat exchanger inlet and discharged into the second fan bypass duct through the heat exchanger outlet. 
     In a refinement, air pressure in the first fan bypass duct is greater than air pressure in the second bypass duct, and a pressure differential between first fan bypass duct and the second fan bypass duct determines a cooling air flow rate through the heat exchanger. In one further refinement, the pressure differential between the first fan bypass duct and the second fan bypass duct varies with a thrust output of the gas turbine engine. 
     In another refinement, the first fan bypass duct channels a first thrust component of the gas turbine engine thrust output, the second fan bypass duct channels a second thrust component of the gas turbine engine thrust output, and a pressure differential between first fan bypass duct and the second fan bypass duct determines a cooling air flow rate through the heat exchanger. 
     In a further refinement, the heat exchanger is in fluid communication with an engine core of the gas turbine engine, and the object of cooling is air received from the engine core. 
     Embodiments of the present invention further include a gas turbine engine. The gas turbine engine includes a first fan; a second fan in fluid communication first fan; a compressor system in fluid communication with at least one of the first fan and the second fan; a combustor in fluid communication with the compressor system and structured to receive a first working fluid stream discharged by the compressor system; a turbine system in fluid communication with the combustor to received and extract power from the first working fluid stream, the turbine system being drivingly coupled to the first fan, the second fan and the compressor system; a first bypass duct in fluid communication with the second fan and structured to conduct a second working fluid stream; a second bypass duct in fluid communication with the first fan and structured to conduct a third working fluid stream; and a heat exchanger in fluid communication with the second working fluid stream and the third working fluid stream and operable to remove heat from an object of cooling. 
     In a refinement, the heat exchanger is in fluid communication with the compressor system, and the object of cooling is air of the first working fluid stream. In a further refinement, the heat exchanger is structured to cool only a portion of the air of the first working fluid stream. In an additional refinement, the gas turbine engine includes a passage is structured to conduct the air of the first working fluid stream to the heat exchanger. In one form, the passage includes a pipe. 
     In another refinement, the first working fluid stream, the second working fluid stream and the third working fluid stream are discharged as thrust output components of the gas turbine engine. 
     In still another refinement, the gas turbine engine includes an exhaust nozzle in fluid communication with at least one of the first bypass duct and the second bypass duct, the exhaust nozzle being operable to control a pressure in at least one of the first bypass duct and the second bypass duct. 
     In yet another refinement, the third working fluid stream has a lower pressure than the second working fluid stream, and the heat exchanger cools the object of cooling using flow generated by the pressure differential between the second working fluid stream and the third working fluid stream. 
     In yet still another refinement, the heat exchanger is located in at least one of the first bypass duct and the second bypass duct. 
     In a further refinement, the gas turbine engine includes an exhaust nozzle structured to control a pressure differential between the first bypass duct and the second bypass duct. 
     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.