Patent Document

CROSS REFERENCE TO RELATED APPLICATIONS 
     The present application claims benefit of U.S. Provisional Patent Application No. 61/428,727, filed Dec. 30, 2010, entitled VARIABLE CYCLE GAS TURBINE ENGINE, which is incorporated herein by reference. 
    
    
     FIELD OF THE INVENTION 
     The present invention relates to gas turbine engines, and more particularly, variable cycle gas turbine engines. 
     BACKGROUND 
     Variable cycle gas turbine engines 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 variable cycle gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for variable cycle gas turbine engines. 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 
       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  illustrates some aspects of a non-limiting example of an aircraft having variable cycle gas turbine engines in accordance with an embodiment of the present invention. 
         FIG. 2  schematically illustrates some aspects of a non-limiting example of a variable cycle gas turbine engine in accordance with an embodiment of the present invention. 
         FIG. 3  schematically illustrates some aspects of a non-limiting example of an auxiliary turbine system in accordance with an embodiment of the present invention. 
     
    
    
     DETAILED DESCRIPTION 
     For purposes of promoting an understanding of the principles of the invention, reference will now be made to the embodiments illustrated in the drawings, and specific language will be used to describe the same. It will nonetheless be understood that no limitation of the scope of the invention is intended by the illustration and description of certain embodiments of the invention. In addition, any alterations and/or modifications of the illustrated and/or described embodiment(s) are contemplated as being within the scope of the present invention. Further, any other applications of the principles of the invention, as illustrated and/or described herein, as would normally occur to one skilled in the art to which the invention pertains, are contemplated as being within the scope of the present invention. 
     Referring to  FIG. 1 , there are illustrated some aspects of a non-limiting example of an aircraft  10  in accordance with an embodiment of the present invention. Aircraft  10  includes a fuselage  12 , wings  14 , an empennage  16  and two propulsion systems  18 . In one form, aircraft  10  is a twin engine military turbofan aircraft. In other embodiments, aircraft  10  may be any fixed-wing aircraft, including turbofan aircraft, turbojet aircraft and turboprop aircraft. In still other embodiments, aircraft  10  may be a rotary-wing aircraft or a combination rotary-wing/fixed-wing aircraft. In various embodiments, aircraft  10  may have a single propulsion engine or a plurality of propulsion engines. In addition, in various embodiments, aircraft  10  may employ any number of wings  14 . Empennage  16  may employ a single or multiple flight control surfaces. 
     Referring to  FIG. 2 , there are illustrated some aspects of a non-limiting example of a propulsion system  18  in accordance with an embodiment of the present invention. Propulsion system  18  includes a gas turbine engine  20  as a main engine, i.e., a main propulsion engine, which includes an auxiliary turbine system  22 . Engine  20  is a primary propulsion engine that provides thrust for flight operations of aircraft  10 . In one form, engine  20  is a two-spool engine having a high pressure (HP) spool  24  and a low pressure (LP) spool  26 . In other embodiments, engine  20  may include three or more spools, for example, and may include an intermediate pressure (IP) spool and/or other spools. In one form, engine  20  is a turbofan engine, wherein LP spool  26  is operative to drive a propulsor  28  in the form of a turbofan (fan) system, which may be referred to as a turbofan, a fan or a fan system. In other embodiments, engine  20  may be a turboprop engine, wherein LP spool  26  powers a propulsor  28  in the form of a propeller system (not shown), e.g., via a reduction gearbox (not shown). In still other embodiments, propulsor  28  may take other forms, such as a helicopter rotor or tilt-wing aircraft rotor or a propfan. In one form, two propulsion systems  18  are coupled to fuselage  12  of aircraft  10 . In other embodiments, one or more propulsion systems  18  may be coupled to other portions of aircraft  10 . For example, one or more propulsion systems  18  may be coupled to each wing  14  and/or empennage  16  in addition to or in place of fuselage-mounted propulsion systems  18 . 
     In one form, engine  20  includes, in addition to auxiliary turbine system  22  and fan system  28 , an accessory gearbox  23 , a bypass duct  30 , a compressor system  32  as part of HP spool  24 , a diffuser  34 , a combustion system  36 , a high pressure (HP) turbine  38  as part of HP spool  24 , a low pressure (LP) turbine  40  as part of LP spool  26 , a nozzle  42 A, and a nozzle  42 B. Accessory gearbox  23  is coupled to HP spool  24  and compressor  32  via conventional means, e.g., a bevel gear set and shafting  25 . In other embodiments, accessory gearbox  23  may be coupled to HP spool  24  and/or LP spool  26  via other mechanical arrangements. In one form, compressor  32  is a variable compressor. In other embodiments, compressor  32  may not be a variable compressor. In one form, compressor  32  is a variable geometry compressor. In other embodiments, compressor  32  may be other types of variable compressors that may or may not employ variable geometry, e.g., including geared compressors that are configured to operate at more than one speed relative to a given shaft input speed. 
     Bypass duct  30  and compressor  32  are in fluid communication with fan system  28 . Nozzle  42 B is in fluid communication with bypass duct  30 . Diffuser  34  is in fluid communication with compressor  32 . Combustion system  36  is fluidly disposed between compressor  32  and HP turbine  38 . LP turbine  40  is fluidly disposed between HP turbine  38  and nozzle  42 B. In one form, combustion system  36  includes a combustion liner (not shown) that contains a continuous combustion process. In other embodiments, combustion system  36  may take other forms, and may be, for example, a wave rotor combustion system, a rotary valve combustion system, a pulse detonation combustion system and/or a slinger combustion system, and may employ deflagration and/or detonation combustion processes. 
     Fan system  28  includes a fan rotor system  48  driven by LP spool  26 . In various embodiments, fan rotor system  48  includes one or more rotors (not shown) that are powered by LP turbine  40 , which may operate at the same or different rotational speeds. Fan system  28  may include one or more stages of vanes (not shown). Bypass duct  30  is operative to transmit a bypass flow generated by fan system  28  around the core of engine  20 . Compressor  32  includes a compressor rotor system  50 . In various embodiments, compressor rotor system  50  includes one or more rotors (not shown) that are powered by HP turbine  38 . HP turbine  38  includes a turbine rotor system  52 . In various embodiments, turbine rotor system  52  includes one or more rotors (not shown) operative to drive compressor rotor system  50 . Turbine rotor system  52  is drivingly coupled to compressor rotor system  50  via a shafting system  54 . LP turbine  40  includes a turbine rotor system  56 . In various embodiments, turbine rotor system  56  includes one or more rotors (not shown) operative to drive fan rotor system  48 . Turbine rotor system  56  is drivingly coupled to fan rotor system  48  via a shafting system  58 . In various embodiments, shafting systems  54  and  58  include a plurality of shafts that may rotate at the same or different speeds and directions. In some embodiments, only a single shaft may be employed in one or both of shafting systems  54  and  58 . LP turbine  40  is operative to discharge the engine  20  core gas flow to nozzle  42 A. 
     During the operation of gas turbine engine  20 , not including the use of auxiliary turbine system  22 , air is drawn into the inlet of fan system  28  and pressurized by fan rotor system  48 . Some of the air pressurized by fan rotor system  48  is directed into compressor  32  as core gas flow, and some of the pressurized air is directed into bypass duct  30  as bypass flow. Compressor  32  further pressurizes the core gas flow received therein from fan system  28 , which is then discharged into diffuser  34 . Diffuser  34  reduces the velocity of the pressurized air, and directs the diffused core gas flow into combustion system  36 . Fuel is mixed with the pressurized air in combustion system  36 , which is then combusted. The core gas flow, in the form of hot gases exiting combustion system  36 , are directed into HP and LP turbines  38  and  40 , e.g., sequentially, which extract energy in the form of mechanical shaft power to drive compressor  32  and fan  28  via respective shafting systems  54  and  58 . In the depicted embodiment, the engine  20  core flow is discharged through nozzle  42 A, and the bypass flow is discharged through nozzle  42 B. In other embodiments, other nozzle arrangements may be employed, e.g., a common nozzle for core and bypass flow; a nozzle for core flow, but no nozzle for bypass flow; or another nozzle arrangement. 
     It is desirable that engine  20  product peak thrust output during aircraft  10  takeoff, and during some aircraft  10  maneuvering operations. In addition, it is desirable that engine  20  operate at high efficiency during cruise conditions, including supercruise conditions, i.e., supersonic cruise without the use of thrust augmentation (e.g., afterburners). Conventionally, a fixed geometry gas turbine engine sized for takeoff thrust conditions yields a greater than ideal specific fuel consumption during cruise conditions because the engine is running at an “off-design” point during cruise conditions. On the other hand, a fixed geometry gas turbine sized for peak efficiency during cruise conditions may have insufficient thrust for desirable takeoff and maneuver performance. In order to maximize thrust at high power, e.g., takeoff and maneuver conditions, engine  20  is configured as a variable cycle gas turbine engine. In particular, engine  20  employs auxiliary turbine system  22  for selectively expanding and contracting the turbine flow capacity of engine  20 . In some embodiments, compressor  32  may be variable, e.g., a variable geometry compressor, which in conjunction with auxiliary turbine system  22  further enhances the cycle variability of engine  20 . 
     Referring to  FIG. 3 , some aspects of a non-limiting example of auxiliary turbine system  22  in accordance with an embodiment of the present invention are schematically depicted. In one form, auxiliary turbine system  22  includes an auxiliary turbine  60 , a valve  62 , a controller  64 , inlet ducting  66  and  68 , and exhaust ducting  70 . In one form, auxiliary turbine  60  is mechanically coupled to accessory gearbox  23  via a shaft  72 , and is coupled to compressor  32  via accessory gearbox  23 . In other embodiments, auxiliary turbine  60  may be coupled to compressor  32  or one or more other HP spool  24  components, e.g., shafting system  54 , via other mechanical arrangements. 
     In one form, Inlet ducting  66  is coupled to a plenum  74  at one end, and is coupled to valve  62  at the other end. Plenum  74  is disposed between the outlet of HP turbine  38  and the inlet of LP turbine  40 , and is operative to receive a portion of the core gas flow exiting HP turbine  38  for use by auxiliary turbine  60 . In other embodiments, the portion of core gas flow for use by auxiliary turbine  60  may be obtained from one or more other turbine stages, in addition to or in place of the HP turbine  38  outlet. In addition, in other embodiments, the portion of core gas flow for use by auxiliary turbine  60  may be obtained via other arrangements, which may or may not employ the use of a plenum for the extraction of the portion of the core gas flow for auxiliary turbine  60 , depending upon the needs of the particular application. 
     Inlet ducting  68  is coupled to valve  62  at one end, and to the inlet of auxiliary turbine  60  at the other end. In one form, portions of inlet ducting  68  that pass through bypass duct  30  are disposed within an aerodynamic strut  76  in order to minimize losses. In other embodiments, other arrangements may be employed. Exhaust ducting  70  is coupled to the outlet of auxiliary turbine  60  at one end, and is configured to direct the exhaust from auxiliary turbine  60  into bypass duct  30  at the other end for conversion to thrust, e.g., via nozzle  42 B. In other embodiments, auxiliary turbine system  22  may be configured to discharge the auxiliary turbine  60  exhaust flow to other locations, for example and without limitation, into nozzle  42 A or overboard engine  20 . Although the depicted embodiment envisions the use of inlet ducting  66 , inlet ducting  68  and exhaust ducting  70  as set forth herein, other embodiments may employ other arrangements to channel flow to and from valve  62  and auxiliary turbine  60 . In addition, although depicted embodiment envisions auxiliary turbine  60  being disposed outside of bypass duct  30 , in other embodiments, auxiliary turbine  60  may be disposed in other locations, including radially inward of bypass duct  30 , inside bypass duct  30 , or upstream or downstream of bypass duct  30 . 
     Valve  62  is configured to regulate the portion of the core gas flow that is received by auxiliary turbine  60 . In one form, valve  62  is configured to modulate the portion of the core gas flow received by auxiliary turbine  60  between a minimum flow amount and a maximum flow amount in accordance with the needs of the particular application. Valve  62  is also configured to close to prevent flow to auxiliary turbine  60 . Valve  62  is controlled by controller  64  to selectively allow or disallow flow through valve  62 . Valve  62  may take any suitable form, and may be, for example and without limitation, a butterfly valve, a gate valve, a poppet valve or any other suitable valve type. Valve  62  is actuated by an actuation mechanism (not shown) under the direction of controller  64 . 
     Controller  64  is communicatively coupled to valve  62  via a communications link  78 . Communications link  78  may take any suitable form, and may be, for example, a wired and/or wireless and/or optical link capable of transmitting control signals to valve  62 . In some embodiments, valve  62  may provide feedback information to controller  64  indicative of valve position, in which case communications link  78  is also configured to transmit feedback signals to controller  64  from valve  62 . In some embodiments, communications link  78  may also be configured to provide electrical power for actuating valve  62 . 
     Controller  64  is configured to execute program instructions to control valve  62  to selectively prevent or allow flow to auxiliary turbine  60 , and to regulate the flow rate to a desired level during engine  20  operations where such flow is desired. The flow regulation, including starting and stopping flow to auxiliary turbine  60  may be based on, for example and without limitation, one or more lookup tables and/or rate schedules, and/or may be based on, for example and without limitation, sensed and/or calculated engine  20  parameters, engine  20  inlet conditions, aircraft  10  speed and/or power lever angle. 
     In one form, controller  64  is microprocessor based and the program instructions are in the form of software stored in a memory (not shown). However, it is alternatively contemplated that controller  64  and the program instructions may be in the form of any combination of software, firmware and hardware, including state machines, and may reflect the output of discreet devices and/or integrated circuits, which may be co-located at a particular location or distributed across more than one location, including any digital and/or analog devices configured to achieve the same or similar results as a processor-based controller executing software or firmware based instructions. In other embodiments, controller  64  may not be configured with the level of functionality associated with a processor-based controller, but rather may be a simple controller configuration. In one form, controller  64  is a gas turbine engine controller, such as a full authority digital electronic control (FADEC) unit. In other embodiments, controller  64  may take any suitable form, and in some embodiments may be a dedicated controller for operating valve  62 . 
     During aircraft  10  takeoff, the power (thrust) output of engine  20  is enhanced by employing auxiliary turbine system  22  to expand turbine flow capacity by opening valve  62  to allow a portion of core gas flow to flow through auxiliary turbine  60 . Auxiliary turbine  60  extracts power from the gas flow and transmits the power via accessory gearbox  23  to compressor  32 , thereby increasing the output of compressor  32 , and hence engine  20 . The exhaust gas from auxiliary turbine  60  is directed into bypass duct  30 , from where it will contribute to the thrust output of engine  20 . During some operating conditions, such as aircraft  10  takeoff, controller  64  may command valve  62  to open fully, thereby providing a maximum flow to auxiliary turbine  20 , yielding a higher takeoff power output by engine  20  than a similar engine not equipped with auxiliary turbine system  22 . The amount by which valve  62  opens may vary with conditions, for example and without limitation, ambient/inlet conditions. During other operating conditions of aircraft  10  that require high thrust levels, controller  64  may command valve  62  to open partially or fully, e.g., depending operating conditions and/or pilot input. During cruise conditions, including supercruise flight, controller  64  may command valve  62  to close fully, thereby contracting the turbine flow capacity of engine  20 , which may result in increased fuel efficiency, as engine  20  is effectively operating closer to design point at the cruise power condition. 
     Embodiments of the present invention include a variable cycle gas turbine engine, comprising: a compressor configured to compress a core gas flow; a combustor in fluid communication with the compressor and configured to combust the core gas flow; a primary turbine drivingly coupled to the compressor and configured to receive the core gas flow, wherein the primary turbine is configured to drive the compressor; an auxiliary turbine drivingly coupled to the compressor; and a valve configured to selectively direct a portion of the core gas flow to the auxiliary turbine, wherein the auxiliary turbine is configured to extract power from the portion of the core gas flow and supply the power to the compressor when the valve is open. 
     In a refinement, the valve is in fluid communication with the primary turbine and operative to receive the portion of the core gas flow from the primary turbine. 
     In another refinement, the valve is configured to modulate the portion of the core gas flow between a minimum flow amount and a maximum flow amount. 
     In yet another refinement, the valve is configured to close to prevent flow to the auxiliary turbine. 
     In still another refinement, the valve is operative to open during a takeoff power condition of the engine. 
     In yet still another refinement, the valve is operative to close during a cruise power condition of the engine. 
     In a further refinement, the compressor is a variable compressor. 
     In a yet further refinement, the variable cycle gas turbine engine further comprises a fan and a fan bypass duct in fluid communication with the fan, wherein variable cycle gas turbine engine is configured to direct the exhaust of the auxiliary turbine into the fan bypass duct. 
     In a still further refinement, the variable cycle gas turbine engine further comprises an accessory gearbox coupled to the compressor, wherein the auxiliary turbine is drivingly coupled to the compressor via the accessory gearbox. 
     Embodiments of the present invention include a variable cycle gas turbine engine, comprising: a compressor configured to compress a core gas flow; a combustor in fluid communication with the compressor and configured to combust the core gas flow; a primary turbine drivingly coupled to the compressor and configured to receive the core gas flow, wherein the primary turbine is configured to drive the compressor; and an auxiliary turbine system having an auxiliary turbine drivingly coupled to the compressor, wherein the auxiliary turbine system is configured to selectively receive a portion of the core gas flow; generate shaft power using the portion of the core gas flow; and supply the shaft power to the compressor. 
     In a refinement, the auxiliary turbine system includes a valve configured to selectively direct a portion of the core gas flow to the auxiliary turbine. 
     In another refinement, the primary turbine is a high pressure turbine; and wherein the valve is in fluid communication with the discharge of the high pressure turbine, and is operative to receive the portion of the core gas flow from the discharge of the high pressure turbine. 
     In yet another refinement, the valve is configured to modulate the portion of the core gas flow between a minimum flow amount and a maximum flow amount. 
     In still another refinement, the variable cycle gas turbine engine further comprises a low pressure turbine; a fan driven by the low pressure turbine; and a fan bypass duct in fluid communication with the fan, wherein variable cycle gas turbine engine is configured to direct the exhaust of the auxiliary turbine into the fan bypass duct. 
     In yet still another refinement, the variable cycle gas turbine engine is configured wherein the portion of the core gas flow is received by the auxiliary turbine from upstream of the low pressure turbine. 
     In a further refinement, the variable cycle gas turbine engine further comprises an accessory gearbox coupled to the compressor, wherein the auxiliary turbine is drivingly coupled to the compressor via the accessory gearbox. 
     In a yet further refinement, the compressor is a variable geometry compressor. 
     Embodiments of the present invention include a variable cycle gas turbine engine, comprising: a compressor configured to compress a core gas flow; a combustor in fluid communication with the compressor and configured to combust the core gas flow; a turbine drivingly coupled to the compressor and configured to receive the core gas flow, wherein the turbine is configured to drive the compressor; and means for selectively expanding and contracting a turbine flow capacity. 
     In a refinement, the means for selectively expanding and contracting the turbine flow capacity includes an auxiliary turbine system having an auxiliary turbine drivingly coupled to the compressor, wherein the auxiliary turbine system is configured to selectively receive a portion of the core gas flow; generate shaft power using the portion of the core gas flow; and supply the shaft power to the compressor. 
     In another refinement, the means for selectively expanding and contracting the turbine flow capacity includes a valve configured to selectively direct a portion of the core gas flow to the auxiliary turbine. 
     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.

Technology Category: 2