Abstract:
One embodiment of the present disclosure is a unique method for operating a gas turbine engine during flight operation of the gas turbine engine in an aircraft. Another embodiment of the present disclosure is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for 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.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application claims priority to U.S. Provisional Patent Application No. 61/769,625 filed Feb. 26, 2013, the contents of which are hereby incorporated in their entirety. 
     
    
     FIELD OF THE DISCLOSURE 
       [0002]    The present disclosure relates to gas turbine engines, and more particularly, to gas turbine engines and methods for operating gas turbine engines. 
       BACKGROUND 
       [0003]    Controlling the flow of fuel in a gas turbine engine remains 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 
       [0004]    One embodiment of the present disclosure is a unique method for operating a gas turbine engine during flight operation of the gas turbine engine in an aircraft. One embodiment of the present disclosure is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for 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 
         [0005]    The description herein makes reference to the accompanying drawings wherein like reference numerals refer to like parts throughout the several views, and wherein: 
           [0006]      FIG. 1  schematically illustrates some aspects of a non-limiting example of a gas turbine engine in accordance with an embodiment of the present disclosure. 
           [0007]      FIG. 2  schematically illustrates some aspects of a non-limiting example of an orientation of fuel injectors and temperature measurement devices in accordance with an embodiment of the present disclosure. 
       
    
    
     DETAILED DESCRIPTION 
       [0008]    For purposes of promoting an understanding of the principles of the disclosure, 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 disclosure is intended by the illustration and description of certain embodiments of the disclosure. 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 disclosure. Further, any other applications of the principles of the disclosure, as illustrated and/or described herein, as would normally occur to one skilled in the art to which the disclosure pertains, are contemplated as being within the scope of the present disclosure. 
         [0009]    Referring to the drawings, and in particular  FIG. 1 , some aspects of a non-limiting example of a gas turbine engine  10  in accordance with an embodiment of the present disclosure is schematically depicted. In one form, gas turbine engine  10  is an aircraft propulsion power plant installed in an aircraft (not shown), and provides power to the aircraft during ground operations, e.g., taxi, and during flight operations. In one form, gas turbine engine  10  is a multi-spool turbofan engine. In other embodiments, gas turbine engine  10  may take other forms, and may be, for example, a turboshaft engine, a turbojet engine, a turboprop engine, or a combined cycle engine having a single spool or multiple spools. In various embodiments, gas turbine engine  10  may be a land-based engine, e.g., for electrical power generation, pumping or other purposes, a marine engine, or any other type of gas turbine engine. 
         [0010]    As a turbofan engine, gas turbine engine  10  includes a fan system  12 , a bypass duct  14 , a compressor  16 , a diffuser  18 , a combustor  20 , a turbine  22 , a discharge duct  26  and a nozzle system  28 . Bypass duct  14  and compressor  16  are in fluid communication with fan system  12 . Diffuser  18  is in fluid communication with compressor  16 . Combustor  20  is fluidly disposed between compressor  16  and turbine  22 . In one form, combustor  20  includes a combustion liner (not shown) that contains a continuous combustion process. In one form, combustor  20  is an annular combustor. In other embodiments, combustor  20  may be in the form of one or more can combustors. In still other embodiments, combustor  20  may take still other forms, and may be, for example and without limitation, a wave rotor combustion system, a rotary valve combustion system or a slinger combustion system, and may employ deflagration and/or detonation combustion processes. 
         [0011]    Fan system  12  includes a fan rotor system  30 . In various embodiments, fan rotor system  30  includes one or more rotors (not shown) that are powered by turbine  22 . Bypass duct  14  is operative to transmit a bypass flow generated by fan system  12  to nozzle  28 . Compressor  16  includes a compressor rotor system  32 . In various embodiments, compressor rotor system  32  includes one or more rotors (not shown) that are powered by turbine  22 . Each compressor rotor includes a plurality of rows of compressor blades (not shown) that are alternatingly interspersed with rows of compressor vanes (not shown). Turbine  22  includes a turbine rotor system  34 . In various embodiments, turbine rotor system  34  includes one or more rotors (not shown) operative to drive fan rotor system  30  and compressor rotor system  32 . Each turbine rotor includes a plurality of turbine blades (not shown) that are alternatingly interspersed with rows of turbine vanes (not shown). 
         [0012]    Turbine rotor system  34  is drivingly coupled to compressor rotor system  32  and fan rotor system  30  via a shafting system  36 . In various embodiments, shafting system  36  includes 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. Turbine  22  is operative to discharge an engine  10  core flow to nozzle  28 . In one form, fan rotor system  30 , compressor rotor system  32 , turbine rotor system  34  and shafting system  36  rotate about an engine centerline  48 . In other embodiments, all or parts of fan rotor system  30 , compressor rotor system  32 , turbine rotor system  34  and shafting system  36  may rotate about one or more other axes of rotation in addition to or in place of engine centerline  48 . Fan rotor system  30  loads, compressor rotor system  32  loads, turbine rotor system  34  loads and shafting system  36  loads are supported and reacted by a plurality of bearing systems, e.g., illustrated as bearing systems  50 ,  52  and  54 . The locations and numbers of bearing systems may vary with the needs of the application—bearing systems  50 ,  52  and  54  are presented for illustrative purposes only. 
         [0013]    Discharge duct  26  extends between a discharge portion  40  of turbine  22  and engine nozzle  28 . Discharge duct  26  is operative to direct bypass flow and core flow from a bypass duct discharge portion  38  and turbine discharge portion  40 , respectively, into nozzle system  28 . In some embodiments, discharge duct  26  may be considered a part of nozzle  28 . Nozzle  28  is in fluid communication with fan system  12  and turbine  22 . Nozzle  28  is operative to receive the bypass flow from fan system  12  via bypass duct  14 , and to receive the core flow from turbine  22 , and to discharge both as an engine exhaust flow, e.g., a thrust-producing flow. In other embodiments, other nozzle arrangements may be employed, including separate nozzles for each of the core flow and the bypass flow. 
         [0014]    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 compressor  16  as core flow, and some of the pressurized air is directed into bypass duct  14  as bypass flow, and is discharged into nozzle  28  via discharge duct  26 . Compressor  16  further pressurizes the portion of the air received therein from fan  12 , which is then discharged into diffuser  18 . Diffuser  18  reduces the velocity of the pressurized air, and directs the diffused core airflow into combustor  20 . Fuel is mixed with the pressurized air in combustor  20 , which is then combusted. The hot gases exiting combustor  20  are directed into turbine  22 , which extracts energy in the form of mechanical shaft power sufficient to drive fan system  12  and compressor  16  via shafting system  36 . The core flow exiting turbine  22  is directed along an engine tail cone  42  and into discharge duct  26 , along with the bypass flow from bypass duct  14 . Discharge duct  26  is configured to receive the bypass flow and the core flow, and to discharge both as an engine exhaust flow, e.g., for providing thrust, such as for aircraft propulsion. 
         [0015]    Fuel is supplied during the operation of engine  10  via a system  70 . System  70  is configured for providing fuel to combustor  20 . System  70  includes a controller  72 ; a plurality of fuel injectors  74  which may also be referred to as fuel nozzles  74 ; a plurality of controllers  76 ; and a plurality of temperature measurement devices  78 . In one form, controller  72  is an engine controller, such as a full authority digital engine controller (FADEC), although controller  72  may take suitable form, and may or may not be an engine controller, depending upon the embodiment. Controller  72  is configured to execute program instructions to control the flow of fuel into combustor  20  via fuel injectors  74 , including balancing the fuel flow delivered by each of the fuel injectors  74  in order to provide a uniform measured gas temperature as measured between each of the temperature measurement devices  78 . 
         [0016]    As an engine controller, controller  72  performs various functions in addition to controlling the flow of fuel into combustor  20 . However, in some embodiments, controller  72  may be a dedicated controller for controlling the flow into combustor  20 , including balancing the fuel flow delivered by each of the fuel injectors  74  in order to provide a uniform measured gas temperature as between each of the temperature measurement devices  78 , or may perform some functions in addition to so controlling the flow into combustor  20 . In one form, controller  72  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 the controller and 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. 
         [0017]    Fuel injectors  74  are configured and operative receive pressurized fuel from a fuel line system, such as a fuel manifold system, and to inject the fuel into combustor  20 . In one form, fuel injectors  74  are circumferentially disposed about engine centerline  48 , e.g., equally spaced from one another circumferentially. In other embodiments, fuel injectors  74  may be arranged in any manner suitable for the particular combustor  20 . The output of each of the fuel injectors  74  is independently variable, and may be varied, for example, to balance the fuel flow delivered by each of the fuel injectors  74  in order to provide a uniform measured gas temperature as between each of the temperature measurement devices  78 . 
         [0018]    In one form, each controller  76  is communicatively coupled to a corresponding fuel injector  74 , and to controller  72  via a communications link  77 . Hence, in one form, the number of controllers  76  corresponds to the number of fuel injectors  74 . In various embodiments, communications link  77  may be a wired, wireless and/or optical link configured to carry analog and/or digital data. Although the present embodiment contemplates a controller  76  corresponding to each fuel injector  74 , in other embodiments, each controller  76  may be associated with or correspond to and be communicatively coupled to more than one fuel injector  76 . Each controller  76  is configured to and operative to control the flow of fuel discharged by its corresponding fuel injector  74 . In one form, each controller  76  is configured to and operative to control the flow of fuel discharged by its corresponding fuel injector  74  under the direction of controller  72 . Other embodiments may not employ controller  72  for controlling controllers  76 , but rather, in some embodiments, controllers  76  may form a self-governing system configured to and operative to control the flow of fuel discharged by individual fuel injectors  74 . Each controller  76  may take a form similar to that described above with respect to controller  72 , or may take any other suitable form. 
         [0019]    Temperature measurement devices  78  are communicatively coupled to controller  72  via a communications link  79 . In various embodiments, communications link  79  may be a wired, wireless and/or optical link configured to carry analog and/or digital data. Temperature measurement devices  78  are configured to and operative to supply measured gas temperature data to the controller  72 . Measured gas temperature is the temperature of the gases flowing past of temperature measurement devices  78 , as measured by temperature measurement devices  78 . In one form, temperature measurement devices  78  are thermocouples. In other embodiments, other temperature measurement devices may be employed in addition to or in place of thermocouples. 
         [0020]    In one form, temperature measurement devices  78  are circumferentially disposed about engine centerline  48 , e.g., equally spaced from one another circumferentially, and are positioned downstream of combustor  20 . In other embodiments, temperature measurement devices  78  may be arranged in any suitable manner. In one form, temperature measurement devices  78  are disposed within the core stream hot gas path downstream of combustor  20 . In one form, temperature measurement devices  78  are disposed within turbine vanes (or nozzles)  80 , which may be any desirable turbine stage, e.g., first stage vanes, second stage vanes, third stage vanes, etc., depending upon the needs of the particular application. In other embodiments, temperature measurement devices  78  may be disposed on an outside portion, e.g., a leading edge of a turbine vane, or may be positioned on or within a strut, or may be positioned at any desirable location, alone or in conjunction with another structure. 
         [0021]    Referring now to  FIG. 2  in conjunction with  FIG. 1 , in a non-limiting example, system  70  includes  16  fuel injectors  74 , designated as fuel injectors  74 A- 74 P, and includes 16 temperature measurement devices  78 , designated as temperature measurement devices  78 A- 78 P, which are disposed downstream of fuel injectors  74 A- 74 P and downstream of combustor  20 . In the depiction of  FIG. 2 , engine centerline  48  is depicted as crosshairs, indicating that engine centerline  48  is perpendicular to the plane of  FIG. 2 . The angular relationship between fuel injectors  74  and temperature measurement devices  78  may vary with the needs of the application. In addition, spacing, e.g., circumferential spacing and/or radial spacing, between fuel injectors  74  themselves, between temperature measurement devices  78  themselves and between fuel injectors  74  and temperature measurement devices  78  may vary with the needs of the application. 
         [0022]    System  70  is configured for providing a fuel flow into combustor  20  via the plurality of fuel injectors  74  for combusting the fuel in combustor  20 ; measuring a gas temperature downstream of combustor  20  using the plurality of temperature measurement devices  78 ; and varying the output of at least one of the fuel injectors  74  in a first direction (i.e., an increase or decrease in output) and the output of at least another of the fuel injectors  74  in a second direction (a decrease or an increase—the second direction being opposite of the first direction), while maintaining a power output of gas turbine engine  10  the same as it was prior to the varying. In some embodiments, the output of a single fuel injector  74  is varied in the first direction, and the output of the balance of the fuel injectors  74  are varied in the second or opposite direction. The varying of the output of the fuel injectors identifies which of the temperature measurement devices  78  is affected or impacted by the particular at least one of the fuel injectors  74  (a particular fuel injector  74  or a group of fuel injectors  74 ). 
         [0023]    By identifying which temperature measurement device(s)  78  is/are impacted by the output of a particular fuel injector  74  (or group of fuel injectors in some embodiments), it is known, e.g., by controller  72 , that the temperature data supplied by those impacted temperature measurement devices is pertinent to the output of the particular fuel injector  74  (or group of fuel injectors  74 ), and hence is used, e.g., by controller  72 , to trim or adjust the output of the particular fuel injector  74  (or group of fuel injectors  74 ). By repeating the process of varying the output of each fuel injector  74  in the first direction and the output of at least another fuel injector  74  (e.g., the balance of fuel injectors  74 ) in the second or opposite direction, it becomes known, e.g., within controller  72 ), which temperature measurement device(s)  78  is/are impacted by the fuel output of which fuel injector  74  (or group of fuel injectors). 
         [0024]    Thus, by identifying which temperature measurement device(s)  78  is impacted by the fuel output of which fuel injector  74 , the fuel flow output of each of the fuel injectors  74  may be trimmed based on temperature data provided from the impacted temperature measurement device(s)  78  to yield a balanced (or equalized) fuel flow as between the fuel injectors  74  in accordance with embodiments of the present disclosure, yielding a uniform measured gas temperature as measured between the plurality of temperature measurement devices. The amount that the fuel flow is varied in the second direction for the at least another fuel injector  74  (e.g., the balance of the fuel injectors) is determined to be sufficient to compensate for the for the amount that the fuel flow is varied in the first direction for the at least one of the fuel injectors  74 , such that the total fuel flow and hence the output of engine  10  remains the same, which is particularly useful during flight operations of an engine  10  installed in an aircraft, whereby the fuel injectors  74  may be trimmed during flight operations, which may yield an improvement in turbine blade life in some embodiments, without adversely affecting engine thrust output. Similar benefits would be obtained by land-based engine, wherein the fuel injectors may be trimmed during operation of the engine, but without adversely affecting, for example, power generation, pumping, etc. 
         [0025]    In one non-limiting example, when a more uniform temperature field at the combustor  20  exit plane is achieved, a longer life for turbine components may result. The fuel flow rate for each fuel injector  74  is determined and updated, e.g., each flight, by a fuel flow variation technique, such as that mentioned above, a non-limiting example of which is herein described. In some embodiments, the flow variations are performed during a phase of the flight when the overall fuel flow is held constant and when the turbine temperatures are moderate enough to be able to withstand modest fuel flow changes without significantly impacting turbine life, such as during the initial cruise portion of a flight. In other embodiments, the fuel flow variations may be performed under other conditions, for both flight engines and non-flight engines. The fuel flow variation and learned temperature impacts would be used to trim flows to individual injectors so as to minimize temperature variability at the turbine inlet. The trimmed fuel flows would then be applied at all flight conditions, including high-temperature conditions such as takeoff, thereby extending the life of the turbine in some embodiments. 
         [0026]    In the example illustrated in  FIG. 2 , the control logic, e.g., program instructions of controller  72 , would initially establish uniform fuel flows to all fuel injectors  74  via a signal, e.g., an electronic signal, sent to each controller  76 . Although the intent may be to flow the same amount of fuel out of each injector, this may not happen for one of a variety of reasons, including, for example and without limitation, manufacturing tolerances in fuel injectors  74 . Then, at any time the overall fuel flow is held constant, maintaining a constant output of engine  10 , e.g., a constant thrust output and/or other form of power output, such as a constant shaft power output, a fuel flow dithering technique is employed to identify which temperature measurement devices  78  are affected by which fuel injectors  74 . Data pertaining to the identification of which temperature measurement device(s)  78  are affected or impacted by which fuel injector(s)  74  is stored in a controller, e.g., controller  72 , or alternatively one or more controllers  76 , for subsequent use in balancing the fuel flow output of the fuel injectors  74  by trimming one or more of the fuel injectors  74  to yield a uniform measured gas temperature as measured between temperature measurement devices  78 , such that the temperature measurement devices  78  each indicate approximately the same temperature. By use of the term, “approximately,” it will be understood that an acceptable tolerance may be established, allowing some acceptable deviation from perfect agreement as between each of the temperature measurement devices  78 . 
         [0027]    For example and without limitation, a fuel flow dithering technique may include reducing the output of a fuel injector  74 , e.g., fuel injector  74 A, by 5% for a brief period, e.g., 10 seconds, while at the same time increasing the output of injectors  74 B- 74 P by 0.33% for the same amount of time, to maintain the same fuel output (and hence, the same thrust or power output of engine  10 ) as prior to the initiation of the dithering technique, (in some embodiments, a plurality of fuel injectors  74  may be adjusted in one direction, and the balance of the fuel injectors adjusted in the other direction). The output of temperature measurement devices  78  is then read, e.g., by controller  72 , which determines which of the temperature measurement devices are impacted by the reduction in fuel flow output from fuel injector  74 A, e.g., based on a decrease in measured gas temperature, since those temperature measurement devices  78  impacted by the particular fuel injector will be exposed to a lower-temperature gas stream than those temperature measurement devices  78  not impacted by the reduction in fuel flow output from fuel injector  74 A. The impacted temperature measurement devices  78  are not necessarily those that are circumferentially aligned with and directly downstream of the particular fuel injector  74 , e.g.,  74 A. For example, for combustors with high dome swirl, e.g., often employed in modern military and lean-burn civil aircraft engines, the impacted temperature measurements devices may not only be the in-line temperature measurement devices  78 A and  78 B, but may also be temperature measurement devices  78 O,  78 P,  78 C and  78 D. 
         [0028]    Once it is determined which temperature measurement devices  78  are impacted, for the first fuel injector  74 A, the process is repeated, e.g., wherein the fuel flow output of the circumferentially adjacent fuel injector  74  is reduced by 5%, and the fuel flow output of the balance of the fuel injectors  74  is increased by 0.33%, and the determination which temperature measurement devices  78  are impacted is made. In one form, the process is repeated, sequentially, e.g., from fuel injector  74 A to fuel injector  74 B, to fuel injector  74 C, etc. to fuel injector  74 P, e.g., in the circumferential direction, until each of the fuel injectors  74 A- 74 P have been subjected to the  5 % reduction in fuel flow output (and the 0.33% increase in fuel flow output). In other embodiments, the dithering process may take place using patterns other than circumferentially sequential. Once the dithering process is completed, controller  72  will have data as to which of the temperature measurement devices  78  are impacted by which of the fuel injectors  74 . This data is used to trim the fuel flow output of one or more fuel injectors  74  based on the output of the known impacted temperature measurement devices  78 . The trimming is performed by controller  72  sending a signal to the controller(s)  76  associated with the fuel injector(s)  74  to be trimmed, in response to which the requisite controller(s)  76  adjust the output of the corresponding fuel injector(s)  74 . The output of the fuel injectors  74  is trimmed to achieve a uniform temperature field as measured by temperature measurement devices  78 , which yields a balanced fuel flow as between fuel injectors  74 . It will be understood that the 5% reduction and 0.33% increase described herein is a non-limiting example, and that other values may be employed in accordance with the needs of the particular application. It will also be understood that instead of a 5% reduction and 0.33% increase, as described above, a 5% increase and 0.33% reduction may be employed or any other values suitable for the particular application. 
         [0029]    While the disclosure 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 disclosure 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 disclosure, 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.