Abstract:
One embodiment of the present invention is a unique aircraft propulsion gas turbine engine. Another embodiment is a unique gas turbine engine. Another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines with heat exchange systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    The present application claims benefit of U.S. Provisional Patent Application No. 61/581,850 filed Dec. 30, 2011, entitled AIRCRAFT PROPULSION GAS TURBINE ENGINE WITH HEAT EXCHANGE, which is incorporated herein by reference. 
     
    
     GOVERNMENT RIGHTS 
       [0002]    The present application was made with United States government support under Contract No. NNH08ZEA001 N Amendment #2, awarded by NASA. The United States government may have certain rights in the present application. 
     
    
     FIELD OF THE INVENTION 
       [0003]    The present invention relates to gas turbine engines, and more particularly, to gas turbine engines with heat exchange systems. 
       BACKGROUND 
       [0004]    Gas turbine heat exchange systems that effectively transfer heat from pressurized compressor air to fuel 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 
       [0005]    One embodiment of the present invention is a unique aircraft propulsion gas turbine engine. Another embodiment is a unique gas turbine engine. Another embodiment is another unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines with heat exchange systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0006]    The description herein makes reference to the accompanying drawings wherein like reference numerals refer to like parts throughout the several views, and wherein: 
           [0007]      FIG. 1  schematically depicts some aspects of a non-limiting example of a gas turbine engine in accordance with an embodiment of the present invention. 
           [0008]      FIG. 2  schematically illustrates some aspects of non-limiting examples of a heat exchanger and a gas turbine engine in accordance with an embodiment of the present invention. 
           [0009]      FIG. 3  schematically illustrates some aspects of non-limiting examples of a heat exchanger and a compressor system in accordance with an embodiment of the present invention. 
           [0010]      FIG. 4  schematically illustrates some aspects of a non-limiting example of a heat exchanger in accordance with an embodiment of the present invention. 
       
    
    
     DETAILED DESCRIPTION 
       [0011]    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. 
         [0012]    Referring to the drawings, and in particular  FIG. 1 , some aspects of a non-limiting example of an engine  10  in accordance with an embodiment of the present invention are schematically depicted. Engine  10  is an aircraft propulsion gas turbine engine. Engine  10  includes a compressor system  12 , a combustion system  14  in fluid communication with compressor system  12 , and a turbine system  16  in fluid communication with combustion system  14 . In one form, compressor system  12 , combustion system  14  and turbine system  16  are disposed about an engine centerline  18 , e.g., the axis of rotation of compressor system  12  and turbine system  16 . In other embodiments, other arrangements may be employed. In various embodiments, engine  10  may be a single spool engine or a multi-spool engine. In various embodiments, engine  10  may or may not have a turbine system, or may have additional turbomachinery components in addition to a compressor system and/or a turbine system, e.g., a fan system. In some embodiments, engine  10  may be a direct propulsion engine that produces thrust directly from combustion system  14 . In other embodiments, combustion system  14  may form a gas generator for a gas turbine propulsion system, or may be employed in a gas turbine engine topping cycle. In still other embodiments, engine  10  may be one or more of other types of gas turbine engines, hybrid engines and/or combined cycle engines. 
         [0013]    Compressor system  12  includes a compressor case  20  that houses stationary and rotating compressor system  12  components. In various embodiments, compressor case  20  may be formed of one or more individual compressor case structures, e.g., depending on the number, size and location of compressor stages and/or the number of spools employed in engine  10 . Combustion system  14  includes a combustor case  22 , a combustor  24  and a plurality of fuel injectors  26 . Combustor  24  receives pressurized air from compressor system  12 . Fuel injectors  26  are configured to inject fuel into combustor  24 . Combustor  24  is configured to combust the fuel injected therein by fuel injectors  26  with pressurized air received from compressor system  12 . Turbine system  16  includes a turbine case  28  that houses stationary and rotating turbine system  16  components. In various embodiments, turbine case  28  may be formed of one or more individual turbine case structures, e.g., depending on the number, size and location of turbine stages and/or the number of spools employed in engine  10 . 
         [0014]    Engine  10  includes a heat exchanger  30  fluidly disposed between two compressor stages and in fluid communication with fuel injectors  26 . Heat exchanger  30  is configured to cool pressurized airflow in compressor system  12  by heat exchange with the fuel supplied to fuel injectors  26 , and to heat the fuel by heat exchange with the pressurized airflow in compressor system  12 . 
         [0015]    Referring to  FIG. 2 , in conjunction with  FIG. 1 , some aspects of non-limiting examples of heat exchanger  30  and engine  10  in accordance with an embodiment of the present invention are schematically depicted. Heat exchanger  30  is an air/fuel heat exchanger that is configured to exchange heat between compressor system  12  air and the fuel supplied to fuel injectors  26  without the use of an intermediate heat transfer fluid. Heat exchanger  30  is disposed within engine  10 . That is, disposed within one or more engine  10  cases, as opposed to being disposed external to gas turbine engine  10 , which requires external ducting to duct the pressurized air to and from heat exchanger  30  from and to compressor stages within engine  10 . By being internal to engine  10 , heat exchanger  30  does not increase the frontal area of engine  10 , which would otherwise adversely impact the flight characteristics and frontal area drag of the aircraft or air-vehicle into which engine  10  is installed as a propulsion power plant. In one form, heat exchanger  30  is an annular heat exchanger, extending annularly around engine centerline  18 . In other embodiments, heat exchanger  30  may take other forms. 
         [0016]    Heat exchanger  30  is in fluid communication with and fluidly disposed between a compressor stage  32  and a compressor stage  34 . Combustor  24  is fluidly disposed downstream of compressor stage  34 . Compressor stage  32  is a lower pressure compressor stage than compressor stage  34 . Compressor stage  32  is configured to produce a pressurized airflow, which is received by compressor stage  34  after having passed through heat exchanger  30 . In one form, compressor stage  34  is a final compressor stage, and combustor  24  is configured to receive compressor discharge air from compressor stage  34  for combustion, e.g., via a diffuser. In other embodiments, compressor stage  34  may not be a final compressor stage. 
         [0017]    Heat exchanger  30  is also in fluid communication with a fuel supply  36  and fuel injectors  26 . Fuel supply  36  is operative to supply fuel to heat exchanger  30  for subsequent delivery to fuel injectors  26  after having performed heat exchange between pressurized air from compressor stage  32  and the fuel prior to delivery of the fuel to fuel injectors  26 . Heat exchanger  30  is configured to receive the pressurized air flow from compressor stage  32 , to discharge the pressurized air flow to compressor stage  34 ; to heat the fuel by heat exchange with the pressurized air flow prior to delivery of the fuel to fuel injector  26 ; and to cool the pressurized air flow by heat exchange with the fuel prior to delivery of the pressurized air flow to compressor stage  34 . 
         [0018]    Referring to  FIG. 3 , some aspects of non-limiting examples of heat exchanger  30  and compressor system  12  in accordance with an embodiment of the present invention are schematically depicted. In one form, compressor stage  32  is an axial compressor stage, whereas compressor stage  34  is a centrifugal compressor stage. In other embodiments, compressor stages  32  and  34  take other forms, e.g., including both compressor stages  32  and  34  being axial compressor stages; both compressor stages  32  and  34  being centrifugal compressor stages; or compressor stage  32  being a centrifugal compressor stage and compressor stage  34  being an axial compressor stage. 
         [0019]    Disposed between and fluidly coupling compressor stage  32  and compressor stage  34  is a primary flowpath  38 . In one form, primary flowpath  38  is annular, extending annularly around engine centerline  18  and forming an annulus therein. In other embodiments, primary flowpath  38  may take other forms. Heat exchanger  30  is disposed within primary flowpath  38 , between compressor stage  32  and compressor stage  34 . Primary flowpath  38  is configured to deliver the pressurized air flow from compressor stage  32  to heat exchanger  30  and then to compressor stage  34 . Primary flowpath  38  is disposed within engine  10 , that is, disposed within one or more engine  10  cases, as opposed to being disposed external to gas turbine engine  10 . By being disposed internal to engine  10 , primary flowpath  38  does not include any engine  10  external ducting to duct pressurized air to and from heat exchanger  30  from and to compressor stages  32  and  34  within engine  10 . By being internal to engine  10 , primary flowpath  38  does not increase the frontal area of engine  10 , which would otherwise adversely impact the flight characteristics or frontal area drag of the aircraft or air-vehicle into which engine  10  is installed as a propulsion power plant. 
         [0020]    In one form, primary flowpath  38  has a maximum radial extent  40 , relative to engine centerline  18 , and heat exchanger  30  has a maximum radial extent  42 , relative to engine centerline  18 , that do not exceed the maximum radial extent  44  relative to engine centerline  18 , ( FIG. 1 ), of turbine case  28 . In one form, maximum radial extent  40  of primary flowpath  38  and/or maximum radial extent  42  of heat exchanger  30  do not exceed the maximum radial extent  46 , relative to engine centerline  18  ( FIG. 1 ), of a turbine blade tip diameter of turbine system  16 . In one form, maximum radial extent  40  of primary flowpath  38  and maximum radial extent  42  of heat exchanger  30  do not exceed the maximum radial extent  48 , relative to engine centerline  18  ( FIG. 1 ), of combustor case  22 . In one form, maximum radial extent  40  of primary flowpath  38  and maximum radial extent  42  of heat exchanger  30  do not exceed the maximum radial extent  50 , relative to engine centerline  18  ( FIG. 1 ), of compressor case  20 , e.g., a high pressure (HP) compressor case  52 , which surrounds and houses compressor stage  34 . In other embodiments, maximum radial extent  40  of primary flowpath  38  and maximum radial extent  42  of heat exchanger  30  may be disposed within the radial extents of other engine  10  components. 
         [0021]    Primary flowpath  38  includes a diffuser portion  54  and a converging portion  56 . Diffuser portion  54  is fluidly disposed upstream of heat exchanger  30 . Diffuser portion  54  is configured to diffuse the air pressurized by compressor stage  32  prior to entry of the pressurized air into heat exchanger  30 . Converging portion  56  is fluidly disposed downstream of heat exchanger  30 . Converging portion  56  is configured to reduce the flow area in primary flowpath  38  and to increase the velocity of the air pressurized by compressor stage  32  after the pressurized air has passed through heat exchanger  30 , prior to delivery of the pressurized air to compressor stage  34 . 
         [0022]    In one form, disposed within diffuser portion  54  is a flow splitter  58 . Some embodiments may not include a flow splitter. Flow splitter  58  is configured to prevent or reduce separation of the pressurized air flow from the walls of diffuser portion  54  upstream of heat exchanger  30 . In one form, flow splitter  58  is configured to enable a more aggressive diffusion angle in diffuser portion  54  than the diffusion angle of a diffuser portion not having a flow splitter, e.g., which allows a reduction in the length of diffuser portion  54  relative to embodiments not equipped with flow splitter  58 . In one form, flow splitter  58  is positioned proximate to heat exchanger  30 , e.g., immediately adjacent to heat exchanger  30 , to prevent recirculation of the pressurized air downstream of flow splitter  58  (between flow splitter  58  and heat exchanger  30 ), e.g., owing to potential pressure differentials between locations above and below splitter  58 , e.g., which may otherwise yield an effective flow blockage. In some embodiments, a seal  60  is disposed between flow splitter  58  and heat exchanger  30  in order to further prevent recirculation downstream of flow splitter  58 . Seal  60  may take any form suitable for fitment and sealing between flow splitter  58  and heat exchanger  30 . 
         [0023]    Referring to  FIG. 4  in conjunction with  FIG. 3 , some aspects of a non-limiting example of heat exchanger  30  in accordance with an embodiment of the present invention are schematically depicted. For clarity of illustration, only heat exchanger  30  is illustrated in  FIG. 4 . As set forth previously, in one form, heat exchanger  30  is an annular heat exchanger. In a particular form, heat exchanger  30  is formed of a plurality of individual heat exchanger modules  62 . In one form, heat exchanger modules  62  are equally spaced apart circumferentially and arranged annularly within primary flowpath  38 , between diffuser portion  54  and converging portion  56 . In other embodiments, heat exchanger modules  62  may be arranged differently. 
         [0024]    In one form, heat exchanger  30  includes sixteen heat exchanger modules  62 . In other embodiments, the number and size of heat exchanger modules  62  may vary with the needs of the particular application. In one form, heat exchanger modules  62  are plate-and-fin heat exchanger modules. In other embodiments, heat exchanger modules  62  may take other forms. In one form, heat exchanger modules  62  are configured for cross-flow heat exchange between the fuel and the air pressurized by compressor stage  32 . In other embodiments, other heat exchange configurations may be employed in place of or in addition to cross-flow, e.g., counter-flow, parallel flow and/or mixed flow. 
         [0025]    Disposed between heat exchanger modules  62  are fuel distribution manifolds  64 . Fuel distribution manifolds  64  are in fluid communication with adjacent heat exchanger modules  62 , and are configured to transmit fuel between the adjacent heat exchanger modules  62 . In one form, fuel distribution manifolds are pie-shaped, owing to the shape of heat exchanger modules  62 . In other embodiments, other suitable shapes may be employed. In one form, heat exchanger  30  is effectively split into two parallel heat exchanger halves with a fuel inlet  66  and a fuel outlet  68  for distributing fuel in a generally circumferential direction  70  through one side of heat exchanger  30 ; and with a fuel inlet  72  and a fuel outlet  74  for distributing fuel in a generally circumferential direction  76  through the other side of heat exchanger  30 . By effectively splitting heat exchanger  30  into two parallel heat exchangers, the circumferential variation in heat transfer to the pressurized air flow provided by compressor stage  32  is reduced. In other embodiments, only a single fuel inlet and a single fuel outlet may be employed, e.g., for distributing the fuel around the entire heat exchanger  30 . In still other embodiments, a plurality of fuel inlets and/or fuel outlets may be employed to distribute fuel in parallel through smaller segments of heat exchanger  30 , e.g., to further reduce the circumferential variation in heat transferred to the pressurized air flow provided by compressor stage  32 . 
         [0026]    In one form, disposed immediately upstream of each fuel distribution manifold  64  is a leading transition  78 . Leading transitions  78  are configured to guide the pressurized airflow around fuel distribution manifolds  64  and into heat exchanger modules  62 , which reduces pressure losses in the pressurized air flow from compressor stage  32 . In some embodiments, trailing transitions may also be positioned downstream of fuel distribution manifolds  64  to reduce pressure losses in air flow exiting heat exchanger  30 . 
         [0027]    In one form, the fuel used by engine  10  is an endothermic fuel, and combustor  24 , fuel injectors  26  and air/fuel heat exchanger  30  are configured for use with the endothermic fuel. Endothermic fuel is a fuel having the fuel molecules pre-split in a manner that does not adversely affect the latent heating value of the fuel. In one form, endothermic fuel has a temperature limit of approximately 900° F. As the allowable fuel temperature increases, more heat can be transferred to the fuel from the pressurized air flow provided by compressor stage  32  via heat exchanger  30 , which increases the specific fuel consumption (SFC) benefit to engine  10  from the use of heat exchanger  30 , relative to the use of lower temperature-capable fuels. Additionally, by using a high temperature capable fuel, purging of heat exchanger  30  after engine  10  shutdown may not be required to avoid fuel coking at high temperatures, e.g., high operating temperatures and hot soak-back conditions. Hence, such embodiments may not require a purge system, reducing the cost and weight of engine  10  relative to systems that do require a purge system. 
         [0028]    In other embodiments, the fuel used by engine  10  is a deox fuel, and combustor  24 , fuel injectors  26  and air/fuel heat exchanger  30  are configured for use with the deox fuel. Deox fuel is a fuel that has been processed to remove oxygen from the fuel. In one form, deox fuel has a temperature limit of approximately 600° F. As the allowable fuel temperature increases, more heat can be transferred to the fuel from the pressurized air flow provided by compressor stage  32  via heat exchanger  30 , which increases the SFC benefit to engine  10  from the use of heat exchanger  30 , relative to the use of lower temperature-capable fuels. Additionally, by using a higher temperature-capable fuel, purging of heat exchanger  30  after engine  10  shutdown may not be required to avoid fuel coking at high temperatures, e.g., high operating temperatures and hot soak-back conditions. Hence, such embodiments may not require a purge system, reducing the cost and weight of engine  10  relative to systems that do require a purge system. 
         [0029]    In still other embodiments, the fuel used by engine  10  is a conventional gas turbine engine fuel, e.g., JP-8, and combustor  24 , fuel injectors  26  and air/fuel heat exchanger  30  are configured for use with the conventional fuel. In one form, conventional fuel has a temperature limit of approximately 450° F. 
         [0030]    Embodiments of the present invention include an aircraft propulsion gas turbine engine, comprising: a first compressor stage configured to produce a pressurized air flow; a second compressor stage disposed downstream of the first compressor stage; a primary annular flowpath fluidly coupling the first compressor stage and the second compressor stage, wherein the primary annular flowpath is disposed within the aircraft propulsion gas turbine engine; a combustor disposed downstream of the second compressor stage; a fuel injector configured to inject a fuel into the combustor, wherein the combustor is configured to combust the fuel injected therein by the fuel injector; and an air/fuel heat exchanger disposed in the primary annular flowpath, wherein the air/fuel heat exchanger is in fluid communication with the fuel injector, the first compressor stage and the second compressor stage; and wherein the air/fuel heat exchanger is configured to receive the pressurized air flow from the first compressor stage, to discharge the pressurized air flow to the second compressor stage, to heat the fuel by heat exchange with the pressurized air flow prior to delivery of the fuel to the fuel injector, and to cool the pressurized air flow by heat exchange with the fuel. 
         [0031]    In a refinement, the air/fuel heat exchanger is an annular heat exchanger. 
         [0032]    In another refinement, the annular heat exchanger includes a plurality of individual heat exchanger modules arranged annularly within the primary annular flowpath to form the annular heat exchanger. 
         [0033]    In yet another refinement, the air/fuel heat exchanger includes a plate-and-fin heat exchanger. 
         [0034]    In still another refinement, the fuel is a deox fuel; and the combustor, the fuel injector and the air/fuel heat exchanger are configured for use with the deox fuel. 
         [0035]    In yet still another refinement, the fuel is an endothermic fuel; and the combustor, the fuel injector and the air/fuel heat exchanger are configured for use with the endothermic fuel. 
         [0036]    In a further refinement, the aircraft propulsion gas turbine engine further comprises an engine case, wherein a maximum radial extent of the primary annular flowpath is less than a maximum radial extent of the engine case. 
         [0037]    In a yet further refinement, the engine case is one of a compressor case, a combustor case and a turbine case. 
         [0038]    In a still further refinement, the engine case is an HP compressor case. 
         [0039]    Embodiments of the present invention include a gas turbine engine, comprising: a first compressor stage configured to produce a pressurized air flow; a second compressor stage disposed downstream of the first compressor stage; a combustor disposed downstream of the second compressor stage; a fuel injector configured to inject a fuel into the combustor, wherein the combustor is configured to combust the fuel injected therein by the fuel injector; and an air/fuel heat exchanger fluidly disposed between the first compressor stage and the second compressor stage, wherein the air/fuel heat exchanger is in fluid communication with the fuel injector, the first compressor stage and the second compressor stage; and wherein the air/fuel heat exchanger is configured to receive the pressurized air flow from the first compressor stage, to discharge the pressurized air flow to the second compressor stage, to heat the fuel by heat exchange with the pressurized air flow prior to delivery of the fuel to the fuel injector, and to cool the pressurized air flow by heat exchange with the fuel prior to delivery of the pressurized air flow to the second compressor stage, wherein air/fuel heat exchanger is disposed within the gas turbine engine. 
         [0040]    In a refinement, the gas turbine engine further comprises a primary annular flowpath fluidly coupling the first compressor stage and the second compressor stage, wherein the air/fuel heat exchanger is disposed within the primary annular flowpath. 
         [0041]    In another refinement, the primary annular flowpath includes a diffuser portion upstream of the air/fuel heat exchanger and a converging portion downstream of the air/fuel heat exchanger. 
         [0042]    In yet another refinement, the gas turbine engine further comprises a flow splitter disposed in the diffuser portion proximate to the air/fuel heat exchanger, wherein the flow splitter is configured to prevent or reduce flow separation in the diffuser portion. 
         [0043]    In still another refinement, the gas turbine engine further comprises a seal disposed between the flow splitter and the air/fuel heat exchanger. 
         [0044]    In yet another refinement, the gas turbine engine further comprises an engine case, wherein a maximum radial extent of the air/fuel heat exchanger is less than a maximum radial extent of the engine case. 
         [0045]    In yet still another refinement, the engine case is one of a compressor case, a combustor case and a turbine case. 
         [0046]    In a further refinement, the engine case is an HP compressor case. 
         [0047]    In a yet further refinement, the gas turbine engine is configured as an aircraft propulsion gas turbine engine. 
         [0048]    Embodiments of the present invention include a gas turbine engine, comprising: a first compressor stage configured to produce a pressurized air flow; a second compressor stage disposed downstream of the first compressor stage; a combustor disposed downstream of the second compressor stage; a fuel injector configured to inject a fuel into the combustor, wherein the combustor is configured to combust the fuel injected therein by the fuel injector; and means for cooling the pressurized air flow prior to delivery of the pressurized air flow to the second compressor stage and for heating the fuel prior to delivery of the fuel to the fuel injector. 
         [0049]    In a refinement, the gas turbine engine further comprising an engine case, wherein a maximum radial extent of the means for cooling and for heating is less than a maximum radial extent of the engine case; and wherein the gas turbine engine is configured as an aircraft propulsion gas turbine engine. 
         [0050]    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.