Gas Turbine Engine Thermal Management

Abstract

A method for controlling a cooling air system for cooling an aircraft gas turbine engine is provided. The method includes receiving data indicating an environmental condition of the aircraft gas turbine engine, data indicating a degradation parameter of the aircraft gas turbine engine, data indicating an operating condition of the aircraft gas turbine engine, or a combination thereof; and modifying a cooling capacity of the cooling air system in response to the received data indicating the environmental condition of the aircraft gas turbine engine, data indicating a degradation parameter of the aircraft gas turbine engine, data indicating an operating condition of the aircraft gas turbine engine, or a combination thereof.

Description

Technical Field

[0001] The present disclosure relates to a gas turbine engine thermal management system. Background Art

[0002] A gas turbine engine typically includes a turbine and a rotor assembly. Gas turbine engines, such as turbofan engines, can be used for aircraft propulsion. The turbine generally includes, in series flow order, a compressor section with one or more compressors, a combustion section, and a turbine section with one or more turbines. In the case of a turbofan engine, the rotor assembly may be configured as a fan assembly.

[0003] To maintain the temperature of one or more turbines in a turbine section below a maximum operating temperature limit, at least some gas turbine engines bypass the combustion section and provide airflow from the compressor section to one or more components of one or more turbines in the turbine section. However, using airflow from the compressor section reduces the efficiency of the compressor section, thereby reducing the efficiency of the gas turbine engine. Therefore, a gas turbine engine cooling system that reduces the amount of airflow from the compressor section required to cool the turbine section would be useful. Summary of the Invention

[0004] Aspects and advantages of the invention will be set forth in part in the description which follows, or may be obvious from the description, or may be learned through practice of the invention.

[0005] In one exemplary aspect of the present disclosure, a method for controlling a cooling air system for cooling an aircraft gas turbine engine is provided. The method includes receiving data indicating an environmental condition of the aircraft gas turbine engine, data indicating a degradation parameter of the aircraft gas turbine engine, data indicating an operating condition of the aircraft gas turbine engine, or a combination thereof; and modifying a cooling capacity of the cooling air system in response to the received data indicating the environmental condition of the aircraft gas turbine engine, data indicating a degradation parameter of the aircraft gas turbine engine, data indicating an operating condition of the aircraft gas turbine engine, or a combination thereof.

[0006] These and other features, aspects and advantages of the present invention will be better understood with reference to the following description and appended claims.The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the present invention and, together with the description, serve to explain the principles of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0007] This specification sets forth a complete and enabling disclosure of the present invention, including the best mode thereof, to one of ordinary skill in the art with reference to the accompanying drawings, wherein:

[0008] Figure 1is a cross-sectional view of a gas turbine engine according to an exemplary aspect of the present disclosure.

[0009] Figure 2 is a cross-sectional view of a gas turbine engine according to another exemplary aspect of the present disclosure.

[0010] Figure 3 is a close-up schematic diagram of a cooled cooling air (CCA) system according to an exemplary aspect of the present disclosure.

[0011] Figure 4 is a schematic diagram of a thermal bus having a CCA heat exchanger integrated therein according to an exemplary aspect of the present disclosure.

[0012] Figure 5 is a schematic diagram of a first cooling fluid supplier and a first CCA heat exchanger according to an exemplary aspect of the present disclosure.

[0013] Figure 6 is a schematic diagram of a second cooling fluid supplier and a second CCA heat exchanger according to an exemplary aspect of the present disclosure.

[0014] Figure 7 is a schematic diagram of a cooling fluid supplier and a CCA heat exchanger removably coupled within a gas turbine engine according to an exemplary aspect of the present disclosure.

[0015] Figure 8 is a flow chart of a method of operating a CCA system according to an exemplary aspect of the present disclosure. DETAILED DESCRIPTION

[0016] Reference will now be made in detail to present embodiments of the present invention, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. In the drawings and the description, like or similar designations are used to refer to like or similar parts of the invention.

[0017] The word "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any embodiment described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments. Furthermore, all embodiments described herein should be considered exemplary unless otherwise indicated.

[0018] As used herein, the terms "first," "second," and "third" may be used interchangeably to distinguish one component from another, and not for the purpose of indicating the position or importance of each component.

[0019] The terms "forward" and "rearward" refer to relative positions within a gas turbine engine or vehicle and refer to the normal operating attitude of the gas turbine engine or vehicle. For example, in the context of a gas turbine engine, forward refers to a position near the engine's air intake, and rear refers to a position near the engine's nozzle or exhaust.

[0020] The terms "upstream" and "downstream" refer to the direction relative to the flow of a fluid in a fluid path. For example, "upstream" refers to the direction from which the fluid is flowing, while "downstream" refers to the direction to which the fluid is flowing.

[0021] The terms “coupled,” “fixed,” “attached,” and the like refer to both direct coupling, fixing, or attachment as well as indirect coupling, fixing, or attachment through one or more intermediate components or features, unless otherwise specified herein.

[0022] The singular forms "a," "an," and "the" include plural referents unless the context clearly dictates otherwise.

[0023] As used herein throughout the specification and claims, approximate language is applied to modify any quantitative representation that can be permissibly varied without resulting in a change in the basic function to which it is associated. Thus, a value modified by one or more terms, such as "about," "approximately," and "substantially," is not limited to the precise value specified. At least in some cases, approximate language can correspond to the precision of an instrument for measuring the value, or the precision of a method or machine for constructing or manufacturing a component and / or system. For example, approximate language can refer to within a range of 1, 2, 4, 10, 15, or 20%. These approximate ranges can apply to a single value, any one or two endpoints of a defined numerical range, and / or the range between the endpoints.

[0024] Here and throughout the specification and claims, range limitations are combined and interchangeable, and such ranges are identified and include all sub-ranges contained therein unless context or language indicates otherwise. For example, all ranges disclosed herein include the endpoints, and the endpoints are independently combinable with each other.

[0025] The present disclosure generally relates to a gas turbine engine having a cooled cooling air ("CCA") system for providing cooling air from a compressor section of the gas turbine engine to a turbine section of the gas turbine engine to maintain one or more components of the turbine section within maximum operating temperature limits. The gas turbine engine is configured to modify the cooling capacity of the CCA system based on one or more parameters. This can allow the engine to reduce unnecessary cooling resources consumed by the CCA system, thereby making the gas turbine engine more efficient.

[0026] For example, in certain exemplary aspects, the gas turbine engine is configured to receive data indicating one or more environmental conditions of the aircraft gas turbine engine, data indicating a degradation parameter of the aircraft gas turbine engine, data indicating an operating condition of the aircraft gas turbine engine, data indicating a fault condition of the aircraft gas turbine engine, or a combination thereof. The gas turbine engine may be configured to modify the cooling capacity of the CCA system in response to the received data.

[0027] In certain embodiments, modifying the cooling capacity of the CCA system may include modifying the amount of cooling available to the CCA system or modifying the cooling capacity of the CCA airflow provided by the CCA system (eg, modifying the temperature and / or mass flow rate of the CCA airflow).

[0028] For example, the gas turbine engine can modify one or more cooling flows provided to the CCA heat exchanger of the CCA system. Additionally, or if the CCA system is integrated into a thermal bus, the gas turbine engine can modify the amount of cooling fluid provided to the radiator heat exchanger (which can cool the cooling fluid subsequently provided to the CCA heat exchanger), the amount of heating fluid provided to the heat source heat exchanger (reducing the heating of the cooling fluid subsequently provided to the CCA heat exchanger), the amount of heated fluid provided to the CCA heat exchanger via the thermal bus, or a combination thereof.

[0029] Furthermore, in other embodiments, modifying the cooling capacity of the CCA system may include swapping hardware components of the aircraft gas turbine engine. For example, in certain exemplary aspects, modifying the available cooling capacity may include increasing the size of piping in a cooling supply providing a flow of cooling fluid to the CCA system, increasing the size of a heat exchanger in a CCA heat exchanger, and the like.

[0030] For example, it will be appreciated that the amount of cooling required for the CCA system may change based on the ambient conditions of the engine and / or the operating conditions of the engine. Similarly, as the engine degrades and the compressor airflow provided to the CCA system increases, the amount of cooling required for the CCA system may also increase for the same engine power output. In this manner, it will be appreciated that a gas turbine engine incorporating one or more of these systems may operate more efficiently by utilizing only the necessary amount of cooling for the CCA system. Furthermore, when the systems and methods of the present disclosure include swapping hardware, this may result in a lighter aircraft gas turbine engine, for example, in the early portion of the engine's life.

[0031] Now refer to Figure 1 , provides a cross-sectional view of an exemplary embodiment of a gas turbine engine that may incorporate one or more inventive aspects of the present disclosure. In particular, Figure 1The exemplary gas turbine engine of FIG. 1 is configured as a single non-ducted rotor engine 10 that defines an axial direction A, a radial direction R, and a circumferential direction C. Figure 1 As shown, the engine 10 takes the form of an open rotor propulsion system and has a rotor assembly 12 that includes an array of airfoils, and more specifically, an array of rotor blades 16, arranged about a central longitudinal axis 14 of the engine 10.

[0032] In addition, as will be explained in detail below, the engine 10 also includes a non-rotating bucket assembly 18, which is positioned rearward of the rotor assembly 12 (i.e., does not rotate relative to the central axis 14), and which includes an array of airfoils also arranged about the central axis 14, and more particularly includes an array of buckets 20 arranged about the central axis 14.

[0033] Rotor blades 16 are typically arranged in an equidistant relationship about centerline 14, with each blade having a root 22 and a tip 24, and a span defined therebetween. Similarly, buckets 20 are also typically arranged in an equidistant relationship about centerline 14, with each bucket having a root 26 and a tip 28, and a span defined therebetween. Rotor assembly 12 also includes a hub 44 located forward of the plurality of rotor blades 16.

[0034] Furthermore, the engine 10 includes a turbine 30 having a core (or high-pressure / high-speed system) and a low-pressure / low-speed system. It will be appreciated that, as used herein, the terms "speed" and "pressure" are used interchangeably with respect to the high-pressure / high-speed system and the low-pressure / low-speed system. Furthermore, it will be appreciated that the terms "high" and "low" are used in the same context to distinguish between the two systems and do not imply any absolute speed and / or pressure values.

[0035] The core generally includes a high-speed compressor 34, a high-speed turbine 36, and a high-speed shaft 38 extending therebetween and connecting the high-speed compressor 34 and the high-speed turbine 36. The high-speed compressor 34, the high-speed turbine 36, and the high-speed shaft 38 may be collectively referred to as the high-speed spool of the engine. Furthermore, a combustion section 40 is located between the high-speed compressor 34 and the high-speed turbine 36. The combustion section 40 may include one or more structures for receiving a mixture of fuel and air and providing a flow of combustion gases through the high-speed turbine 36 to drive the high-speed spool.

[0036] The low-speed system similarly includes a low-speed turbine 42, a low-speed compressor or supercharger 44, and a low-speed shaft 46 extending between and connecting the low-speed compressor 44 and the low-speed turbine 42. The low-speed compressor 44, the low-speed turbine 42, and the low-speed shaft 46 may be collectively referred to as the low-speed spool of the engine.

[0037] Although engine 10 is described as having low-speed compressor 44 positioned forward of high-speed compressor 34, in some embodiments, compressors 34, 44 may be arranged in a staggered arrangement. Furthermore, although engine 10 is described as having high-speed turbine 36 positioned forward of low-speed turbine 42, in some embodiments, turbines 36, 42 may similarly be arranged in a staggered arrangement.

[0038] Still refer to Figure 1 , the turbine 30 is generally enclosed in a cowling 48. Furthermore, it will be appreciated that the cowling 48 at least partially defines an air inlet 50 and an exhaust 52, and includes a turbomachinery flow path 54 extending therebetween. The air inlet 50, in the illustrated embodiment, is an annular or axisymmetric 360-degree air inlet 50 positioned between the rotor blade assembly 12 and the stationary or stationary vane assembly 18, and provides a path for incoming atmospheric air to enter the turbomachinery flow path 54 inboard of the guide vanes 28 (as well as the compressors 44, 34, the combustion section 40, and the turbines 36, 42) along a radial direction R. Such a location may be advantageous for various reasons, including managing icing performance and protecting the air inlet 50 from various objects and materials that may be encountered during operation.

[0039] However, in other embodiments, the air inlet 50 may be positioned at any other suitable location, such as, for example, behind the bucket assembly 18 , arranged in a non-axisymmetric manner, etc.

[0040] In short, it will be appreciated that engine 10 is generally configured as a "high bypass gas turbine engine." In this manner, engine 10 may define a bypass ratio of at least 6:1, such as at least 8:1, such as at least 10:1, such as at least 12:1, such as up to 28:1. As used herein, the term bypass ratio generally refers to the ratio of airflow through rotor assembly 12 over cowling 48 to airflow through air inlet 50. High bypass ratio engines are generally designed for more efficient operation.

[0041] As shown, the rotor assembly 12 is driven by a turbine 30, more specifically, a low speed spool. More specifically, also in Figure 1 In the illustrated embodiment, the engine 10 includes a power gearbox 56, and the rotor assembly 12 is driven by the low-speed spool of the turbine 30 across the power gearbox 56. In this manner, the rotating rotor blades 16 of the rotor assembly 12 can rotate about the axis 14 and generate thrust to propel the engine 10, and thereby propel the aircraft associated therewith, in a forward direction.

[0042] The power gearbox 56 may include a gear set for reducing the rotational speed of the low-speed spool relative to the low-speed turbine 42 so that the rotor assembly 12 can rotate at a slower rotational speed than the low-speed spool.

[0043] As briefly mentioned above, the engine 10 includes a bucket assembly 18. The bucket assembly 18 extends from the fairing 48 and is positioned rearwardly of the rotor assembly 12. The buckets 20 of the bucket assembly 18 may be mounted on a stationary frame or other mounting structure and do not rotate relative to the central axis 14. Figure 1 As shown, the rotor assembly 12 is positioned forward of the turbine 30 in a "pull" configuration, with the exhaust duct 52 positioned behind the guide vanes 28. As can be appreciated, the vanes 20 of the vane assembly 18 can be configured to straighten the airflow from the rotor assembly 12 (e.g., reduce swirl in the airflow) to improve the efficiency of the engine 10. For example, the size, shape, and configuration of the vanes 20 can impart a reactive swirl to the airflow from the rotor blades 16 so that in a downstream direction behind the two rows of airfoils (e.g., blades 16, vanes 20), the airflow has a significantly reduced degree of swirl, which can translate into an increased level of induced efficiency.

[0044] Still refer to Figure 1 It may be desirable for the rotor blades 16, the vanes 20, or both to include a pitch variation mechanism such that the airfoils (e.g., blades 16, vanes 20, etc.) may be rotated independently or in conjunction with one another relative to a pitch axis of rotation. Such pitch variation may be used to vary thrust and / or swirl effects under various operating conditions, including adjusting the magnitude or direction of thrust generated on the rotor blades 16, or providing a thrust reversal feature, which may be useful under certain operating conditions, such as during landing of an aircraft, or to desirably adjust at least in part the acoustic noise generated by the rotor blades 16, vanes 20, or from the aerodynamic interaction of the rotor blades 16 relative to the vanes 20. More specifically, for Figure 1 In the embodiment of the present invention, rotor assembly 12 is described as having a pitch change mechanism 58 for rotating rotor blades 16 about their respective pitch axes 60, and bucket assembly 18 is described as having a pitch change mechanism 62 for rotating buckets 20 about their respective pitch axes 64.

[0045] However, it is understandable that Figure 1 The exemplary single rotor non-ducted engine 10 described in the accompanying drawings is for illustration only, and in other exemplary embodiments, the engine 10 may have any other suitable configuration including, for example, any other suitable number of shafts or spools, turbines, compressors, etc.; fixed pitch blades 16, 20, or both; a direct drive configuration (i.e., the gearbox 56 may not be included); etc.

[0046] In addition, or in other exemplary embodiments, any other suitable gas turbine engine may be provided. For example, in other exemplary embodiments, the gas turbine engine may be a ducted turbofan engine, a turboshaft engine, a turboprop engine, a turbojet engine, a mixed-flow turbofan or turbojet engine, etc. Furthermore, for example, although the engine is described as a single non-ducted rotor engine, in other embodiments, the engine may include a multi-stage open rotor configuration, and various aspects of the disclosure described below may be incorporated therein.

[0047] Furthermore, in still other exemplary embodiments, engine 10 may be configured as a ducted turbofan engine. For example, referring briefly to Figure 2 , depicts an engine 10 according to another exemplary embodiment of the present disclosure. Figure 2 The exemplary embodiment of Figure 1 The exemplary engine 10 is constructed in substantially the same manner, and like or similar reference numerals may refer to like or similar components. However, as will be appreciated, for the illustrated embodiment, the engine 10 further includes a nacelle 80 that at least partially circumferentially surrounds the rotor assembly 12 and the turbine 30, defining a bypass passage 82 therebetween.

[0048] Now back Figure 1 , it will be further understood that the exemplary engine 10 includes a duct 84 extending from the compressor to a bypass airflow above the cowling 48. In certain exemplary embodiments, the duct 84 can be configured as an exhaust airflow duct for exhausting airflow from the compressor section during certain operations. In addition, or in other words, the duct 84 can be configured as a third flow. With such a configuration, the duct 84 can be an annular duct for providing additional airflow over the cowling 48 to generate additional thrust for the engine 10. In any case, for the embodiment shown, the engine 10 further includes a valve 85 and a heat exchanger 86. The valve 85 can regulate the airflow through the duct 84. The heat exchanger 86 can be thermally connected to the airflow through the duct 84.

[0049] Furthermore, exemplary engine 10 includes a fuel system 88. Fuel system 88 may generally include a fuel source 90 and one or more fuel lines 91. Fuel line(s) 91 are configured to provide a flow of fuel from fuel source 90 to combustion section 40 of engine 10.

[0050] Furthermore, exemplary gas turbine engine 10 is operatively connected to a controller 94. Controller 94 may be an engine controller for engine 10 (e.g., a full authority digital engine control controller), an aircraft controller, a controller dedicated to power system 100, or the like.

[0051] Controller 94 may be configured to receive data indicative of various operating conditions and parameters of engine 10 during operation of engine 10. For example, Figure 1 It will be appreciated that the engine 10 includes one or more sensors 96 configured to sense data indicative of various operating conditions and parameters of the engine 10, such as rotational speed, temperature, pressure, vibration, etc. For example, the one or more sensors 96 may sense data indicative of temperature parameters within the engine 10, such as exhaust temperature, turbine inlet temperature (measured as the temperature of the inlet to the turbine rotor blades of the first stage, also referred to as "T41"), compressor outlet temperature (also referred to as "T3"), etc.

[0052] Additionally, or alternatively, one or more sensors 96 may sense data indicative of the speed of the engine 10 , such as the rotational speed of the low-pressure system, the rotational speed of the high-pressure system, the rotational speed of the rotor section 12 , etc. In this manner, the sensors 96 may sense data indicative of the operating condition of the engine, such as the operating mode (e.g., takeoff operating mode, climb operating mode, cruise operating mode, descent operating mode, etc.).

[0053] Furthermore, for the illustrated embodiment, the one or more sensors 96 include an environmental condition sensor for sensing data indicative of one or more environmental conditions, such as ambient temperature, ambient pressure, ambient altitude, and the like.

[0054] It is understandable that Figure 1 The exemplary sensor 96 described in the foregoing is provided by way of example only. In other exemplary embodiments, the sensor 96 may be positioned at any other suitable location, may include any other suitable number of sensors 96, and may further be configured to sense any other data.

[0055] Furthermore, as will be appreciated from the description herein, the controller 94 may also be configured to receive data from other sources, such as data from an aircraft including the engine, such as data from one or more sensors of the aircraft including the engine.

[0056] With particular reference to the operation of controller 94, in at least some embodiments, controller 94 may include one or more computing devices 98. Computing devices 98 may include one or more processors 98A and one or more memory devices 98B. One or more processors 98A may include any suitable processing device, such as a microprocessor, a microcontroller, an integrated circuit, a logic device, and / or other suitable processing device. One or more memory devices 98B may include one or more computer-readable media, including, but not limited to, non-transitory computer-readable media, RAM, ROM, a hard disk, a flash drive, and / or other memory devices.

[0057] One or more memory devices 98B may store information accessible by one or more processors 98A, including computer-readable instructions 98C executable by one or more processors 98A. Instructions 98C may be any set of instructions that, when executed by one or more processors 98A, cause one or more processors 98A to perform an operation. In some embodiments, instructions 98C may be executed by one or more processors 98A to cause one or more processors 98A to perform operations, such as any operations and functions for which controller 94 and / or computing device 98 are configured, operations for operating gas turbine engine 10 (e.g., method 400), as described herein, and / or any other operations or functions of one or more computing devices 98. Instructions 98C may be software written in any suitable programming language or may be implemented in hardware. Additionally, and / or alternatively, instructions 98C may be executed in logically and / or virtually independent threads of processor 98A. Memory device 98B may further store data 98D accessible by processor 98A. For example, data 98D may include data indicative of power flow, data indicative of engine 10 / aircraft operating conditions, and / or any other data and / or information described herein.

[0058] The computing device 98 may also include a network interface 98E for communicating with, for example, other components of the engine 10, the aircraft incorporating the engine 10, the power system 100, and the like. For example, in the depicted embodiment, as described above, the engine 10 includes one or more sensors 96 for sensing data indicative of one or more parameters of the engine 10 and various accessory systems. The controller 94 is operatively coupled to these components via, for example, the network interface 98E, such that the controller 94 can receive data indicative of various operating parameters sensed by the one or more sensors 96 during operation, various operating conditions of the components, and the like, and can further provide commands to control the amount of current drawn by the power system 100 and other operating parameters of these systems, for example, in response to the data sensed by the one or more sensors 96 and other conditions.

[0059] The network interface 98E may include any suitable components for interfacing with one or more networks, including, for example, transmitters, receivers, ports, controllers, antennas, and / or other suitable components. For example, in the illustrated embodiment, the network interface 98E is configured to communicate wirelessly with these components using a wireless communication network (e.g., a wireless communication network). Figure 1 (shown by the dashed communication line in the middle).

[0060] The techniques discussed herein make reference to computer-based systems and the actions taken by, and information sent to, and from, computer-based systems. Those skilled in the art will recognize that the inherent flexibility of computer-based systems allows for a vast number of possible configurations, combinations, and divisions of tasks and functions between components. For example, the processes discussed herein can be implemented using a single computing device or multiple computing devices working in combination. Databases, memories, instructions, and applications can be implemented on a single system or distributed across multiple systems. Distributed components can operate sequentially or in parallel.

[0061] Still refer to Figure 1 , now also refer to Figure 3 , provides Figure 1 A schematic diagram of an exemplary gas turbine engine 10 is provided, and Figure 1 Referring to further details of accessory systems of the exemplary gas turbine engine 10 , it will be appreciated that the exemplary gas turbine engine 10 further includes a cooled cooling air (“CCA”) system.

[0062] The CCA system 100 generally includes a CCA duct 102 extending from the compressor section to the turbine section and a CCA heat exchanger 104. More specifically, the CCA duct 102 is configured to receive an airflow (CCA airflow) from the high-pressure compressor 34, such as an airflow from a downstream stage of the high-pressure compressor 34, such as an airflow from the outlet of the high-pressure compressor 34. In particular, for the illustrated embodiment, the CCA system 100 further includes a mixing chamber 106, and the CCA duct 102 includes an inlet section 108 and an outlet section 110. The inlet section 108 extends from the compressor section and further includes a first portion 112 extending to the CCA heat exchanger 104 and a second portion 114 extending to the mixing chamber 106. The first portion 112 of the inlet section 108 further extends from the CCA heat exchanger 104 to the mixing chamber 106. In this manner, it can be understood that the first portion 112 and the second portion 114 extend generally in parallel.

[0063] The CCA gas stream from the first portion 112 of the inlet section 108 and the second portion 114 of the inlet section 108 are mixed together in the mixing chamber 106 .

[0064] Briefly, for the illustrated embodiment, the CCA system 100 further includes a gas flow valve 116 in fluid communication with the first portion 112 of the inlet section 108 for controlling the volume of CCA gas flow through the first portion 112 of the inlet section 108 of the CCA conduit 102. In this manner, the gas flow valve 116 can control the volume and / or temperature of the CCA gas flow provided to the turbine section through the CCA conduit 102. The gas flow valve 116 can be operably coupled to the CCA conduit 102 as described above. Figure 1 Controller 94 is described. In this manner, airflow valve 116 can control the cooling capacity of the CCA airflow provided to, for example, a turbine section.

[0065] The CCA duct 102 is further configured to provide the CCA flow to the turbine section. More specifically, for the embodiment shown, the CCA duct 102 further includes an outlet section 110 that extends from the mixing chamber 106 to the turbine section. The outlet section 110 of the CCA duct 102 is configured to deliver the CCA flow to the turbine section. Specifically, Figure 3 As shown, the high pressure turbine generally includes first stage turbine rotor blades and second stage turbine rotor blades. The outlet section 110 of the CCA duct 102 is configured to deliver the CCA airflow to the turbine blades of the first stage 118 of the high pressure turbine 36 and the turbine rotor blades of the second stage 120 of the high pressure turbine 36.

[0066] It is understandable that Figure 3 The exemplary CCA system 100 depicted in FIG is provided by way of example only. For example, although the exemplary CCA system 100 is described as extending from the compressor section to the turbine section at a location outside the combustion section 40 in the radial direction R, in other embodiments, the CCA system 100 may be configured to deliver CCA airflow to the turbine section at a location inside the combustion section 40 in the radial direction R. Furthermore, although the CCA system 100 is described as providing cooling airflow to the turbine blades of the first stage 118, in other embodiments, the CCA airflow may be used to cool any other suitable components, such as any other turbine rotor components (e.g., high-pressure compressor components, low-pressure turbine components), an exhaust system, a structural component (e.g., a frame or casing, such as certain clearance control systems), or an oil cooling system.

[0067] Still refer to Figure 3, the exemplary gas turbine engine 10 further includes a cooling fluid supplier 122 for providing a cooling fluid flow to the CCA heat exchanger 104. Specifically, for the embodiment shown, the cooling fluid supplier 122 includes a cooling fluid source 124, a cooling fluid conduit 126, and a cooling fluid valve 128. The cooling fluid source 124 can provide a cooling fluid flow to the CCA heat exchanger 104 through the cooling fluid conduit 126. Within the CCA heat exchanger 104, the cooling fluid can accept heat from the CCA airflow and reduce the temperature of the CCA airflow. The cooling fluid valve 128 can control the volume of the cooling fluid flow from the cooling fluid source 124 through the cooling fluid conduit 126 to the CCA heat exchanger 104. In this way, the cooling fluid valve 128 can control the cooling capacity of the CCA system 100, and more specifically, can control the amount of cooling available for the CCA system 100. The cooling fluid valve 128 can also be used with reference to the above. Figure 1 A controller 94 is depicted as being operatively connected.

[0068] In particular, as will be understood from the description herein, the gas turbine engine 10 may be configured to receive data indicative of an environmental condition of the gas turbine engine 10, data indicative of a degradation parameter of the gas turbine engine 10, data indicative of an operating condition of the gas turbine engine 10, or a combination thereof. In response, the gas turbine engine 10 may be configured to modify the cooling capacity of the CCA system 100 in response to the received data. In certain exemplary embodiments, the engine 10 may modify the amount of cooling available to the CCA system 100 by opening or closing the cooling fluid valve 128 using the controller 94 and / or may actuate the valve 116 to vary the cooling capacity of the CCA airflow.

[0069] In this manner, the gas turbine engine 10 may conserve energy by providing only the amount of cooling required by the CCA system 100 based on the received data, thereby making the gas turbine engine 10 more efficient.

[0070] For example, in response to receiving data indicating that the ambient condition is one or more of a hot day or high altitude, the gas turbine engine 10 can increase the amount of cooling available to the CCA system 100 so that the CCA system 100 adequately cools the turbine section, and in particular maintains the temperature of one or more components of the turbine section within their respective maximum operating temperature ranges.

[0071] Additionally, or alternatively, in response to receiving data indicative of a degradation parameter of the gas turbine engine 10, the degradation parameter indicating that the gas turbine engine 10 has degraded beyond a design condition, the gas turbine engine 10 may increase the amount of cooling available to the CCA system 100 to allow the CCA system 100 to adequately cool the turbine section. The increase in the amount of cooling available to the CCA system 100 may be proportional to the degree of degradation indicated by the degradation parameter. In particular, it will be appreciated that as the gas turbine engine 10 degrades, the CCA airflow provided from the compressor section may begin at a higher temperature (compared to an engine operating at design) as a result of, for example, reduced efficiency within the compressor section of the gas turbine engine 10. Consequently, in order to provide the desired amount of cooling to one or more components of the turbine section (e.g., the turbine rotor blades of the first stage of the high-pressure turbine), it may be necessary to provide additional cooling to the CCA heat exchanger 104.

[0072] Furthermore, in response to receiving data indicative of the operating condition of the gas turbine engine 10, the gas turbine engine 10 may increase or decrease the amount of cooling available to the CCA system 100. For example, in response to receiving data indicative of the operating condition being a low-power operating mode, the gas turbine engine 10 may decrease the amount of cooling available to the CCA system 100. In contrast, in response to receiving data indicative of the operating condition being a high-power operating mode, the gas turbine engine 10 may increase the amount of cooling available to the CCA system 100.

[0073] Still further, the gas turbine engine 10 may receive data indicating a fault condition of the gas turbine engine 10 , such as one or more of a ruptured duct or pipe, a failed or degraded seal, a cracked blade, a flooded circuit, etc. In response, the gas turbine engine 10 may modify the amount of cooling available to the CCA system 100 , again in response to the data, so that the CCA system 100 effectively cools the turbine section. For example, the gas turbine engine 10 may increase the amount of cooling available to the CCA system 100 , in response to the received data.

[0074] In each of the above examples, the gas turbine engine 10 may increase the cooling capacity of the CCA gas flow by actuating the valve 116 in addition to or as an alternative to increasing the cooling available to the CCA system 100, and / or may decrease the cooling capacity of the CCA gas flow by actuating the valve 116 in addition to or as an alternative to decreasing the cooling available to the CCA system 100. Figure 8 Other exemplary aspects of these control steps of the present disclosure are described in more detail.

[0075] In at least some exemplary aspects, the cooling fluid flow may be a bypass air flow (e.g., air flow from a bypass air flow path, such as air flow from bypass passage 82), a fuel flow (e.g., fuel flow from a fuel system of gas turbine engine 10), a hot fluid flow from a thermal bus, an ambient air flow, a flow exhaust gas flow (e.g., air flow from or through conduit 84), or a combination thereof. Generally speaking, utilizing more of such cooling fluid flow to increase the cooling available for CCA system 100 in response to one or more of the above data may result in a decrease in the efficiency of gas turbine engine 10.

[0076] For example, in at least some exemplary embodiments, the CCA system 100 can be integrated with a thermal bus (also referred to as a thermal transport bus 200). Figure 4 , a heat transfer bus 200 is provided according to an exemplary aspect of the present disclosure, into which the CCA system 100 is integrated.

[0077] The heat transfer bus 200 includes an intermediate heat exchange fluid flowing therethrough and may be formed by one or more suitable fluid conduits. The heat exchange fluid may be an incompressible fluid having a high temperature operating range. Alternatively, the heat exchange fluid may be a single-phase fluid, or alternatively, a phase-change fluid. In certain exemplary embodiments, the heat exchange fluid may be a supercritical fluid, such as supercritical carbon dioxide.

[0078] The pump 202 is in fluid communication with the heat exchange fluid in the heat transfer bus 200 to generate a flow of the heat exchange fluid in / through the heat transfer bus 200. Figure 2 As shown, pump 202 may generate a flow of heat exchange fluid in a generally clockwise direction through heat transfer bus 200 .

[0079] In addition, the exemplary thermal management system 100 includes one or more heat source exchangers 204 in thermal communication with the heat exchange fluid in the heat transfer bus 200. Specifically, the depicted thermal management system 100 includes a plurality of heat source exchangers 204. The plurality of heat source exchangers 204 are configured to transfer heat from one or more accessory systems of the turbofan engine 10 (or in operable communication with the engine 10) to the heat exchange fluid in the heat transfer bus 200. For example, in certain exemplary embodiments, the plurality of heat source exchangers 204 may include one or more of the following: a main lubrication system heat exchanger for transferring heat from the main lubrication system; an ACC system heat source exchanger for transferring heat from the ACC system; a generator lubrication system heat source exchanger for transferring heat from the generator lubrication system; an ECS heat exchanger for transferring heat from the ECS; an electronics cooling system heat exchanger for transferring heat from the electronics cooling system; a vapor compression system heat exchanger; an air circulation system heat exchanger; and an auxiliary system heat source exchanger.

[0080] For the depicted embodiment, the CCA system 100 is integrated into a thermal bus 200 , wherein the CCA heat exchanger 104 is configured as a heat source exchanger 204 .

[0081] Furthermore, for the depicted embodiment, there are three heat source exchangers 204 (including the CCA heat exchanger 104). The heat source exchangers 204 are each arranged in a serial flow arrangement along the heat transfer bus 200. However, in other exemplary embodiments, any other suitable number of heat source exchangers 204 may be included, and one or more heat source exchangers 204 may be arranged in a parallel flow arrangement along the heat transfer bus 200 (in addition to or in lieu of the depicted serial flow arrangement). For example, in other embodiments, there may be a single heat source exchanger 204 in thermal communication with the heat exchange fluid in the heat transfer bus, or there may be at least two heat source exchangers 204, at least four heat source exchangers 204, at least five heat source exchangers 204, or at least six heat source exchangers 204 in thermal communication with the heat exchange fluid in the heat transfer bus 200.

[0082] also, Figure 4 The exemplary thermal management system 100 further includes one or more heat sink exchangers 206 in permanent or selective thermal communication with the heat exchange fluid in the heat transfer bus 200. The one or more heat sink exchangers 206 are located downstream of the plurality of heat source exchangers 204 and are configured to transfer heat from the heat exchange fluid in the heat transfer bus 200 to, for example, the atmosphere, fuel, fan flow, etc. For example, in certain embodiments, the one or more heat sink exchangers 206 may include at least one of a RAM heat exchanger, a fuel heat exchanger, a fan flow heat exchanger, an exhaust air heat exchanger, an engine intercooler, or a cold air output of an air circulation system. The RAM heat exchanger may be configured as an "air to heat exchange fluid" heat exchanger that is integrated into one or both of the gas turbine engine 10 or an aircraft including the turbofan engine 10. During operation, the RAM heat exchanger may remove heat from any heat exchange fluid therein by passing a volume of RAM air through the RAM heat exchanger. Furthermore, the fuel heat exchanger is a "fluid-to-heat exchange fluid" heat exchanger, wherein heat from the heat exchange fluid is transferred to the fluid fuel stream for the gas turbine engine 10. Furthermore, the fan flow heat exchanger is typically an "air-to-heat exchange fluid" heat exchanger, which flows, for example, bypass air over the heat exchange fluid to remove heat from the heat exchange fluid. Furthermore, the exhaust air heat exchanger is typically an "air-to-heat exchange fluid" heat exchanger, which flows, for example, exhaust air from an LP compressor over the heat exchange fluid to remove heat from the heat exchange fluid.

[0083] for Figure 4In the embodiment described, the one or more heat sink exchangers 206 of the thermal management system 100 include a plurality of individual heat sink exchangers 206. More particularly, for Figure 4 In an embodiment, the one or more radiator exchangers 206 include three radiator exchangers 206 arranged in series. The three radiator exchangers 206 are configured as a RAM heat exchanger, a fuel heat exchanger, and a fan flow heat exchanger. However, in other exemplary embodiments, the one or more radiator exchangers 206 may include any other suitable number of radiator exchangers 206. For example, in other exemplary embodiments, a single radiator exchanger 206 may be provided, at least two radiator exchangers 206 may be provided, at least four radiator exchangers 206 may be provided, or at least five radiator exchangers 206 may be provided. Furthermore, in other exemplary embodiments, two or more of the one or more radiator exchangers 206 may alternatively be arranged to flow in parallel with each other.

[0084] Still refer to Figure 4 In the exemplary embodiment depicted in FIG, one or more of the plurality of heat sink exchangers 206 and one or more of the plurality of heat source exchangers 204 are selectively in thermal communication with a heat exchange fluid in the heat transfer bus 200. More particularly, the depicted thermal management system 100 includes a plurality of bypass lines 208 for selectively bypassing each heat source exchanger 204 and each heat sink exchanger 206 in the plurality of heat sink exchangers 206. Each bypass line 208 extends between an upstream connection point 210 upstream of the corresponding heat source exchanger 204 or heat sink exchanger 206 and a downstream connection point 212 downstream of the corresponding heat source exchanger 204 or heat sink exchanger 206.

[0085] In addition, each bypass line 208 is connected to the heat transfer bus 200 at a corresponding upstream connection point 210 via a three-way valve 214. The three-way valves 214 each include an inlet fluidly connected to the heat transfer bus 200, a first outlet fluidly connected to the heat transfer bus 200, and a second outlet fluidly connected to the bypass line 208. The three-way valves 214 can be variable throughput three-way valves, such that the three-way valves 214 can vary the throughput from the inlet to the first outlet and / or the second outlet. For example, the three-way valves 214 can be configured to provide between zero percent (0%) and one hundred percent (100%) of the heat exchange fluid from the inlet to the first outlet. Similarly, the three-way valves 214 can be configured to provide between zero percent (0%) and one hundred percent (100%) of the heat exchange fluid from the inlet to the second outlet.

[0086] Notably, three-way valve 214 may be in operable communication with a controller (eg, controller 94 ) of gas turbine engine 10 and / or an aircraft including turbofan engine 10 .

[0087] In addition, each bypass line 208 also connects to the heat transfer bus 200 at a corresponding downstream connection point 212. Between each heat source exchanger 204 or heat sink exchanger 206 and the downstream connection point 212, the heat transfer bus 200 includes a check valve 216 to ensure the correct flow direction of the heat exchange fluid. More specifically, the check valve 216 prevents the heat exchange fluid from flowing from the downstream connection point 212 to the corresponding heat source exchanger 204 or heat sink exchanger 206.

[0088] It is worth noting that for the embodiment shown, as described above, the CCA system 100 is integrated into Figure 4 In the thermal management system described in . Specifically, for the illustrated embodiment, the CCA heat exchanger 104 is configured as a heat source exchanger 204-1. In this configuration, the valve 128 for modifying the amount of cooling available to the CCA heat exchanger 104 in response to, for example, received data is a three-way valve 214-1. Furthermore, in this configuration, the cooling fluid supplier 122 includes an upstream portion of the thermal bus 200, including the heat sink exchanger 206.

[0089] However, it will be appreciated that in other exemplary embodiments, the CCA heat exchanger 104 may be configured to receive the cooling fluid 302 in any other suitable manner. For example, referring now to Figure 5 , provides a schematic diagram of a CCA heat exchanger 104 and a cooling fluid supplier 122 according to another exemplary embodiment of the present disclosure. The CCA heat exchanger 104 and the cooling fluid supplier 122 may be incorporated into Figure 1 or Figure 2 engine 10, or any other suitable engine 10.

[0090] For the illustrated embodiment, the cooling fluid supplier 122 is an airflow conduit 300 having a flow of cooling fluid 302 flowing therethrough. The CCA heat exchanger 104 is positioned within the airflow conduit 300 and is in thermal communication with the flow of cooling fluid 302 flowing therethrough.

[0091] exist Figure 5 In the embodiment depicted in FIG, the amount of cooling available to the CCA heat exchanger 104 may be relatively fixed based on the configuration of the cooling fluid supplier 122, the size of the CCA heat exchanger 104, or both. For example, Figure 5The airflow duct 300 depicted in FIG. 1 defines a fixed cross-sectional area, indicated by the cross-sectional measurement 304 . In this manner, the airflow duct 300 may only allow a certain amount of cooling fluid 302 to reach the CCA heat exchanger 104 .

[0092] As described above, over the life of the engine, the CCA system 100 may require more cooling to ensure that it can provide the required amount of cooling to the turbine section of the engine. In at least some previous gas turbine engines, the CCA system 100 may have been designed to provide the necessary amount of cooling to the turbine section not only when the engine is brand new and operating at "as designed" specifications, but also when the engine is nearing the end of its life and operating under severe environmental conditions (e.g., hot weather, high altitude).

[0093] In contrast, for at least some of the depicted exemplary embodiments, the CCA system 100 and the cooling fluid supply 122 are designed to provide the required amount of cooling for the CCA system 100 during a first portion of the life of the engine 10, such that the CCA system 100 and the cooling fluid supply 122 can provide the required amount of cooling for the turbine section until a certain threshold of degradation of the engine 10 occurs. For example, Figure 5 In the described embodiment, it will be appreciated that the compressor section of the gas turbine engine 10, including the CCA system 100 and the depicted supplied cooling fluid 302, defines an "as designed" compressor temperature during a first steady-state operating condition, an actual compressor temperature during the first steady-state operating condition, and a compressor temperature at the expected end of life during the first steady-state operating condition. Furthermore, the turbine section of the gas turbine engine 10 may define a maximum turbine temperature limit.

[0094] The first steady-state operating condition may be a takeoff operating condition or a climb operating condition. The designed compressor temperature may be a designed compressor outlet temperature during the takeoff operating condition or the climb operating condition, and the actual compressor temperature may be an actual compressor outlet temperature during the takeoff operating condition or the climb operating condition. The maximum turbine temperature limit may be a temperature limit of the turbine rotor blades of the first stage of the high-pressure turbine 36.

[0095] It will be understood that the term "as designed" refers to the value of a particular parameter of the engine under new conditions during standard daily operating conditions, e.g., less than 100 hours of operation. Similarly, it will be understood that the term "end of life" with respect to compressor temperature at expected end of life refers to the expected compressor temperature of the gas turbine engine under operating conditions when the engine is in a condition requiring a strip and overhaul (e.g., requiring repair or replacement of one or more blades or vanes, repair or replacement of one or more seals within the engine's turbomachinery flow path, or the like). An engine condition requiring a strip and overhaul may be a condition where the exhaust gas temperature reaches or exceeds an exhaust gas temperature limit during high power operation of the engine (e.g., takeoff or climb).

[0096] For the illustrated embodiment, the cooling fluid supplier 122 is configured to provide a flow of cooling fluid 302 to the CCA heat exchanger 104 to maintain the temperature of one or more components of the turbine section within a maximum turbine temperature limit during a takeoff operating condition or a climb operating condition, as long as the actual compressor temperature during the operating condition is at most 1.15 times the designed compressor temperature, such as at most 1.1 times the designed compressor temperature, such as at most 1.08 times the designed compressor temperature, such as at most 1.06 times the designed compressor temperature, and at least until the actual compressor temperature during the operating condition is at most 1.03 times the designed compressor temperature. For example, it will be appreciated that the engine defines a lifetime temperature rise between the compressor temperature at the expected end of life under operating conditions and the designed compressor temperature, as well as an actual temperature rise between the actual compressor temperature under operating conditions and the designed compressor temperature. For the illustrated embodiment, the cooling fluid supplier 122 is configured to provide a flow of cooling fluid 302 to the CCA heat exchanger 104 to maintain the temperature of one or more components of the turbine section within a maximum turbine temperature limit during a takeoff operating condition or a climb operating condition so long as the actual temperature rise during the operating condition is at most seventy percent (70%) of the lifetime temperature rise, for example, at most sixty percent (60%), and at least until the actual temperature rise during the operating condition is thirty percent (30%) of the lifetime temperature rise.

[0097] This limitation applies to operation of the engine under normal ambient conditions, as well as operation under adverse ambient conditions (e.g., hot weather, high altitude). It will be appreciated that such a configuration may allow the CCA heat exchanger 104 and cooling fluid supplier 122 to adequately cool the turbine section of the gas turbine engine 10 during a first portion 112 of the engine's life (e.g., at least approximately 10% of the engine's life, such as at least approximately 25% of the engine's life, such as at least approximately 50% of the engine's life), but not during a later portion or end of the engine's life.

[0098] It is noteworthy that although the above examples are described in conjunction with a CCA system 100 for providing CCA airflow to the turbine section of an engine, in other embodiments, the CCA system 100 may be additionally or alternatively configured to provide CCA airflow to any other suitable engine component, such that the maximum turbine temperature limit may instead be a maximum component temperature limit (e.g., a maximum temperature limit of any other turbine rotor component (e.g., a high-pressure compressor component, a low-pressure turbine component), an exhaust system, a structural component (e.g., a frame or, for example, a casing, such as certain clearance control systems), or an oil cooling system).

[0099] As can be appreciated, the requirements for the CCA system 100 may increase significantly over the operational life of the gas turbine engine 10. For example, in one exemplary configuration, the CCA system 100 may require a thermal capacity to provide a temperature reduction of 175 degrees Fahrenheit for a first mass of CCA airflow at a first operating condition at a design state, and may further require a thermal capacity to provide a temperature reduction of 850 degrees Fahrenheit for a second mass of CCA airflow at a first operating condition at an end-of-life state. The second mass of CCA airflow may be at least 5% greater than the first mass of CCA airflow, for example, at least 10% greater than the first mass of CCA airflow. This increase in the thermal capacity required of the CCA system 100 may be further exacerbated by environmental conditions.

[0100] Once it is determined that the engine has deteriorated to Figure 5 To the extent that the CCA heat exchanger 104 and cooling fluid supply 122 described in the accompanying drawings may not be sufficient to adequately cool the turbine section of the gas turbine engine 10 under all expected ambient conditions, the CCA heat exchanger 104, the cooling fluid supply 122, or both may be modified, replaced, and / or supplemented so that the CCA heat exchanger 104 and the cooling fluid supply 122 can provide additional cooling capacity to adequately cool the turbine section of the gas turbine engine 10 under all expected ambient conditions. This may occur during an overhaul of the engine 10, during maintenance operations on the engine 10, and the like.

[0101] For example, now refer to Figure 6 In at least some embodiments, the cooling duct 300 of the cooling fluid supplier 122 can be replaced with a larger cooling duct 300, defining a larger cross-sectional area, to allow for greater airflow through the cooling duct 300. Additionally, the CCA heat exchanger 104 can alternatively be replaced with a larger CCA heat exchanger 104 to increase cooling of the CCA airflow.

[0102] For example, in certain exemplary embodiments, the cross-sectional area of the cooling duct 300 can be increased by at least about 10%, such as at least 20%, such as at least 30%, such as at least 40%, such as up to 200%. Similarly, in at least certain embodiments, the maximum capacity of the CCA heat exchanger 104 can be increased by at least about 10%, such as at least 20%, such as at least 30%, such as at least 40%, such as up to 200%. In this discussion, the maximum capacity of the CCA heat exchanger 104 is measured in British thermal units per hour (Btu / hr).

[0103] Now refer to Figure 7 , provides a cross-sectional view of a CCA heat exchanger 104 and a cooling fluid supply 122 according to an exemplary embodiment of the present disclosure. It will be appreciated that in order to interchange the hardware of the CCA heat exchanger 104, the cooling fluid supply 122, or both, it may be desirable to removably couple these components within the gas turbine engine 10. Specifically, for the illustrated embodiment, a cooling duct 300 of the cooling fluid supply 122 is removably coupled within the gas turbine engine 10. More specifically, for the illustrated embodiment, a plurality of fasteners 304 are used to removably couple the cooling duct 300 of the cooling fluid supply 122 within the gas turbine engine 10. For the illustrated embodiment, the plurality of fasteners 304 couple the cooling duct to the cowling 48 of the engine 10; however, in other embodiments, the cooling duct 302 of the cooling fluid supply 122 may be removably coupled to any other suitable location.

[0104] Similarly, for the illustrated embodiment, the CCA heat exchanger 104 is removably coupled within the gas turbine engine 10. More specifically, for the illustrated embodiment, the CCA heat exchanger 104 is also removably coupled to the cooling duct 300 of the cooling fluid supplier 122 using a plurality of fasteners 304. However, it will be appreciated that in other embodiments, the CCA heat exchanger 104 may alternatively be removably coupled to any other suitable location within the gas turbine engine 10.

[0105] Furthermore, it will be appreciated that in other exemplary embodiments, other hardware modifications may additionally or alternatively be made to increase the cooling capacity of the CCA system 100. For example, in certain exemplary embodiments, the hardware modifications may include adding an airflow splitter to increase the airflow through the duct 300, adding an airflow cooler to cool the airflow through the duct 300, etc. Additionally, alternatively, still, when the CCA heat exchanger 104 is integrated into a thermal bus (e.g., Figure 4 When the heat bus 200 is added to the heat bus 200, the hardware modification may include swapping one or more heat source exchangers or heat sink exchangers, removing one or more heat source exchangers, adding heat sink exchangers, etc. For example, a heat sink heat exchanger may be added along with an air cycle machine, the cooling capacity of the heat bus (and the CCA heat exchanger 104) may be increased by adding a steam compressor system, etc.

[0106] Now refer to Figure 8 , a flow chart of a method 400 for controlling a CCA system of an aircraft gas turbine engine is provided. This exemplary method 400 may be used in conjunction with the above reference Figures 1 to 6 However, in other embodiments, the method 400 may be used with any other suitable gas turbine engine.

[0107] Method 400 includes, at (402), receiving data indicating an environmental condition of an aircraft gas turbine engine, data indicating a degradation parameter of the aircraft gas turbine engine, data indicating an operating condition of the aircraft gas turbine engine, or a combination thereof; and, at (404), modifying an amount of cooling capacity of a CCA system in response to the data received at (402).

[0108] More specifically, for the exemplary aspects described, receiving data at (402) includes receiving data indicative of a degradation parameter of the aircraft gas turbine engine at (406), and modifying the amount of cooling capacity of the CCA system at (404) includes modifying the amount of capacity of the CCA system at (408) in response to the data indicative of the degradation parameter of the aircraft gas turbine engine received at (406). The data indicative of the degradation parameter may include data indicative of turbine blade inlet temperature, exhaust temperature, or both. Furthermore, alternatively, the degradation parameter may be a parameter based on engine age (e.g., accumulated hours, etc.), any turbine temperature parameter, or a calculated parameter based on one or more of these temperatures and / or any other suitable parameter for tracking degradation of the gas turbine engine.

[0109] The degradation parameter generally characterizes the degree of degradation of the engine. Degradation generally occurs through wear of seals, damage to components, etc., which results in reduced operating efficiency of the engine. As the engine degrades, the compressor outlet temperature may increase for the same engine output, and similarly, T3 and T41 may also increase for the same engine output. Therefore, compared to an engine operating as designed, an older engine will require more cooling for the CCA system to allow the CCA system to maintain the temperature of the turbine section below the maximum turbine temperature limit (e.g., maintain the temperature of the first stage turbine rotor blades below the temperature limit of the first stage turbine rotor blades).

[0110] Further to the described exemplary aspects, receiving data at (402) includes receiving data indicative of an ambient condition of the aircraft gas turbine engine at (410), and modifying the amount of cooling capacity of the CCA system at (404) includes modifying the amount of cooling capacity of the CCA system at (412) in response to the data indicative of the ambient condition of the aircraft gas turbine engine received at (410). The data indicative of the ambient condition may include an ambient temperature, an ambient altitude, or both.

[0111] Furthermore, still for the described exemplary aspects, receiving data at (402) includes receiving data indicating an operating condition of the aircraft gas turbine engine at (414), and modifying the amount of cooling capacity of the CCA system at (404) includes modifying the amount of cooling capacity of the CCA system at (416) in response to the data indicating an operating condition of the aircraft gas turbine engine received at (414). The operating condition may be an operating mode of the gas turbine engine, such as a takeoff operating mode, a climb operating mode, a cruise operating mode, etc.

[0112] Furthermore, for the described exemplary aspects, receiving data at (402) may further include receiving data indicating a fault condition of the aircraft gas turbine engine at (418). In such exemplary aspects, modifying the amount of cooling capacity of the CCA system at (404) may further include modifying the amount of cooling capacity of the CCA system at (420) in response to the data indicating a fault condition of the gas turbine engine received at (418).

[0113] In certain exemplary aspects, modifying the amount of cooling capacity of the CCA system at (404) includes modifying the amount of cooling of the cooling air system available for cooling at (421). In such a configuration, method 400 can adjust the capacity of the CCA system to provide cooling to the CCA airflow.

[0114] For example, in the described exemplary aspect, modifying the amount of cooling available to the cooling air system for cooling at (421) includes modifying a cooling fluid flow to a CCA heat exchanger of the CCA system at (422). The cooling fluid flow can be a bypass air flow, a fuel flow, a hot fluid flow from a thermal bus, an ambient air flow, a flow exhaust air flow, or a combination thereof.

[0115] For example, modifying the flow of cooling fluid to the CCA heat exchanger of the CCA system at (422) may include: actuating a valve in flow communication with a pipe or conduit for providing a flow of cooling fluid to the CCA heat exchanger; actuating a variable geometry component in flow communication with the pipe for providing a flow of cooling fluid to the CCA heat exchanger (e.g., actuating a baffle, door, scoop, etc. constructed to increase or decrease the flow of fluid through the pipe); regulating some other flow modifying device in flow communication with the pipe or conduit for providing a flow of cooling fluid to the CCA heat exchanger; and the like.

[0116] Additionally, alternatively, the CCA heat exchanger can be integrated into the thermal bus. The thermal bus can include a radiator exchanger. With such an exemplary embodiment, modifying the amount of cooling available at (404) can further include modifying the flow of hot fluid within the thermal bus to the CCA heat exchanger at (424), and / or can further include modifying the flow of hot fluid within the thermal bus to the radiator exchanger at (426). The radiator exchanger can be a fuel radiator exchanger, an air-cooled radiator exchanger, or both. In this manner, the method 400 can directly or indirectly modify the amount of cooling available to the CCA system.

[0117] Additionally, alternatively, still in certain exemplary aspects, modifying the amount of cooling available at (404) can further include modifying the hardware configuration of the engine at (425). For example, at (425), method 400 can swap one or more hardware components on the engine to change the cooling capacity of the CCA system, such as installing a different cooling fluid supplier with an increased capacity to provide cooling fluid, installing a different CCA heat exchanger or an additional CCA heat exchanger to increase the cooling capacity of the CCA system, changing the configuration of a thermal bus that provides cooling to the CCA heat exchanger (e.g., adding a cooling heat exchanger, removing a heat source heat exchanger, swapping an existing heat exchanger with a different heat exchanger of different capacity, etc.), or a combination thereof.

[0118] Furthermore, in addition to or as an alternative to modifying the amount of cooling available to the cooled cooling air system at (421), modifying the amount of cooling capacity of the CCA system at (404) may further include, at (428), modifying the thermal capacity of the CCA airflow provided to the turbine section of the aircraft gas turbine engine in response to the data received at (402). Modifying the thermal capacity at (428) may include modifying the temperature of the airflow, modifying the volume of the airflow, or both.

[0119] It will be appreciated that although the method 400 is described as providing CCA airflow to a turbine section of an engine (e.g., a high-pressure turbine), in other exemplary aspects, the method 400 may additionally or alternatively provide CCA airflow to any other suitable location of the engine (e.g., a high-pressure compressor, a low-pressure turbine, any other rotating component, structure, housing, etc.).

[0120] This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.

[0121] Further aspects are provided by the subject matter of the following clauses:

[0122] A method for controlling a cooling air system for cooling an aircraft gas turbine engine, the method comprising: receiving data indicating an environmental condition of the aircraft gas turbine engine, data indicating a degradation parameter of the aircraft gas turbine engine, data indicating an operating condition of the aircraft gas turbine engine, or a combination thereof; and modifying a cooling capacity of the cooling air system in response to the received data indicating the environmental condition of the aircraft gas turbine engine, data indicating the degradation parameter of the aircraft gas turbine engine, data indicating an operating condition of the aircraft gas turbine engine, or a combination thereof.

[0123] The method of one or more of these clauses, wherein modifying the cooling capacity of the cooled cooling air system comprises modifying an amount of cooling available for the cooled cooling air system.

[0124] A method according to one or more of these clauses, wherein the step of receiving comprises receiving data indicative of the degradation parameter of the aircraft gas turbine engine, and wherein the step of modifying comprises modifying the amount of the cooling capacity of the cooled cooling air system in response to the received data indicative of the degradation parameter of the aircraft gas turbine engine.

[0125] A method according to one or more of these clauses, wherein the data indicative of the degradation parameter comprises data indicative of turbine blade inlet temperature, exhaust temperature, or both.

[0126] A method according to one or more of these clauses, wherein the step of receiving comprises receiving data indicative of the environmental conditions of the aviation gas turbine engine, and wherein the step of modifying comprises modifying the amount of the cooling capacity of the cooled cooling air system in response to the received data indicative of the environmental conditions of the aviation gas turbine engine.

[0127] A method according to one or more of these clauses, wherein the data indicative of the environmental condition comprises ambient temperature, ambient altitude, or both.

[0128] A method according to one or more of these clauses, wherein the step of receiving comprises receiving data indicative of the operating condition of the aviation gas turbine engine, and wherein the step of modifying comprises modifying the amount of the cooling capacity of the cooled cooling air system in response to the received data indicative of the operating condition of the aviation gas turbine engine.

[0129] The method according to one or more of these clauses further comprises receiving data indicative of a fault condition of the aviation gas turbine engine, and wherein the step of modifying further comprises modifying the amount of the cooling capacity of the cooled cooling air system in response to the received data indicative of the fault condition of the aviation gas turbine engine.

[0130] The method according to one or more of these clauses, wherein the step of modifying comprises modifying the flow of cooling fluid to a CCA heat exchanger of the cooled cooling air system.

[0131] The method of one or more of these clauses, wherein the cooling fluid flow is a bypass air flow, a fuel flow, a hot fluid flow from a thermal bus, an ambient air flow, a flow exhaust gas flow, or a combination thereof.

[0132] A method according to one or more of these clauses, wherein the cooled cooling air system includes a CCA heat exchanger, wherein the CCA heat exchanger is integrated into a thermal bus, wherein the thermal bus further includes a radiator heat exchanger, and wherein the step of modifying further includes modifying the flow of thermal fluid within the thermal bus to the radiator heat exchanger, the temperature of the thermal fluid flow within the thermal bus, or both.

[0133] The method of one or more of these clauses, wherein modifying the cooling capacity of the cooled cooling air system comprises modifying a hardware configuration of the engine to change the cooling capacity of the cooled cooling air system.

[0134] A method according to one or more of these clauses, wherein modifying the cooling capacity of the cooled cooling air system includes modifying the amount of CCA airflow provided to the turbine section of the aviation gas turbine engine in response to received data indicating the ambient condition of the aviation gas turbine engine, data indicating the degradation parameter of the aviation gas turbine engine, data indicating the operating condition of the aviation gas turbine engine, or a combination thereof.

[0135] An aviation gas turbine engine comprises: a compressor section, a combustor section and a turbine section arranged in a serial flow order; a cooled cooling air system, the cooled cooling air system comprising a CCA heat exchanger in airflow communication with the compressor section and the turbine section, the cooled cooling air system being configured to receive airflow from the compressor section and provide CCA airflow to the turbine section; and a controller operably connected to the cooled cooling air system, the controller being configured to: receive data indicating an environmental condition of the aviation gas turbine engine, data indicating a degradation parameter of the aviation gas turbine engine, data indicating an operating condition of the aviation gas turbine engine, or a combination thereof; and modify the amount of cooling capacity of the CCA heat exchanger in response to the received data indicating the environmental condition of the aviation gas turbine engine, the received data indicating the degradation parameter of the aviation gas turbine engine, the received data indicating the operating condition of the aviation gas turbine engine, or a combination thereof.

[0136] An aviation gas turbine engine comprises: a compressor section, a combustor section, and a turbine section arranged in a serial flow order, the compressor section defining a designed compressor temperature during a first steady-state operating condition and an actual compressor temperature during the first steady-state operating condition, the gas turbine engine defining a maximum component temperature limit during the first steady-state operating condition; a cooled cooling air system, the cooled cooling air system comprising a CCA heat exchanger in airflow communication with the compressor section and the turbine section, the cooled cooling air system being configured to receive airflow from the compressor section and to provide a CCA airflow to the turbine section; and a cooling fluid supplier in fluid communication with the CCA heat exchanger, wherein the cooling fluid supplier is configured to provide a cooling fluid flow to the CCA heat exchanger to maintain a component temperature within the maximum component temperature limit during the first steady-state operating condition when the actual compressor temperature during the first steady-state operating condition is at most 1.15 times the designed compressor temperature.

[0137] An aviation gas turbine engine according to one or more of these clauses, wherein the designed compressor temperature is the designed compressor outlet temperature during the first steady-state operating condition, and wherein the actual compressor temperature is the actual compressor outlet temperature during the first steady-state operating condition.

[0138] Aviation gas turbine engine according to one or more of these clauses, wherein the first steady-state operating condition is a takeoff operating condition or a climb operating condition.

[0139] An aviation gas turbine engine according to one or more of these clauses, wherein the turbine section includes a high-pressure turbine having a first stage of turbine rotor blades, wherein the CCA heat exchanger is in airflow communication with the turbine rotor blades of the first stage of the high-pressure turbine, and wherein the maximum component temperature limit is the maximum allowable operating temperature of the turbine rotor blades of the first stage.

[0140] An aviation gas turbine engine according to one or more of these clauses, wherein the cooled cooling air system defines a maximum cooling capacity limited by the size of the cooling fluid supplier, the size of the CCA heat exchanger, or both.

[0141] An aviation gas turbine engine according to one or more of these clauses, wherein the cooling fluid supplier is constructed to provide the cooling fluid flow to the CCA heat exchanger to maintain the turbine temperature within the maximum component temperature limit during the first steady-state operating condition when the actual compressor temperature during the first steady-state operating condition is at most 1.08 times the designed compressor temperature.

Claims

1. A method for controlling a cooling air system for cooling an aircraft gas turbine engine, characterized in that: The method comprises: receiving data indicative of an environmental condition of the aircraft gas turbine engine, data indicative of a degradation parameter of the aircraft gas turbine engine, data indicative of an operating condition of the aircraft gas turbine engine, or a combination thereof; and modifying a cooling capacity of the cooled cooling air system in response to received data indicative of the ambient condition of the aircraft gas turbine engine, data indicative of the degradation parameter of the aircraft gas turbine engine, data indicative of an operating condition of the aircraft gas turbine engine, or a combination thereof; wherein modifying the cooling capacity of the cooled cooling air system comprises modifying the cooling capacity of a cooling fluid provided to a CCA heat exchanger of the cooled cooling air system; and Wherein modifying the cooling capacity of the cooling fluid provided to the CCA heat exchanger of the cooled cooling air system includes modifying aspects other than a temperature of a cooling fluid flow provided to the CCA heat exchanger of the cooled cooling air system.

2. The method according to claim 1, characterized in that Wherein modifying the cooling capacity of the cooled cooling air system comprises modifying an amount of cooling available for the cooled cooling air system.

3. The method according to claim 1, characterized in that wherein the step of receiving comprises receiving data indicative of the degradation parameter of the aircraft gas turbine engine, and wherein the step of modifying comprises modifying the amount of the cooling capacity of the cooled cooling air system in response to the received data indicative of the degradation parameter of the aircraft gas turbine engine.

4. The method according to claim 3, characterized in that The data indicative of the degradation parameter includes data indicative of turbine blade inlet temperature, exhaust temperature, or both.

5. The method according to claim 1, wherein Wherein the step of receiving includes receiving data indicative of the ambient condition of the aircraft gas turbine engine, and wherein the step of modifying includes modifying the amount of the cooling capacity of the cooled cooling air system in response to the received data indicative of the ambient condition of the aircraft gas turbine engine.

6. The method according to claim 5, characterized in that The data indicating the environmental conditions include environmental temperature, environmental altitude, or both.

7. The method according to claim 1, characterized in that wherein the step of receiving comprises receiving data indicative of the operating condition of the aircraft gas turbine engine, and wherein the step of modifying comprises modifying the amount of the cooling capacity of the cooled cooling air system in response to the received data indicative of the operating condition of the aircraft gas turbine engine.

8. The method according to claim 1, characterized in that Further comprising receiving data indicative of a fault condition of the aircraft gas turbine engine, and wherein the step of modifying further comprises modifying the amount of the cooling capacity of the cooled cooling air system in response to the received data indicative of the fault condition of the aircraft gas turbine engine.

9. The method according to claim 1, characterized in that The cooling fluid flow is a bypass air flow, a fuel flow, a hot fluid flow from a thermal bus, an ambient air flow, a flow exhaust gas flow, or a combination thereof.

10. The method according to claim 1, characterized in that wherein the CCA heat exchanger is integrated into a thermal bus, wherein the thermal bus further comprises a radiator heat exchanger, and wherein the step of modifying further comprises modifying a flow of thermal fluid within the thermal bus to the radiator heat exchanger, a temperature of the thermal fluid flow within the thermal bus, or both.

11. The method according to claim 1, wherein Wherein modifying the cooling capacity of the cooled cooling air system includes modifying the hardware configuration of the engine to change the cooling capacity of the cooled cooling air system, wherein modifying the hardware configuration includes replacing one or more hardware on the engine to change the cooling capacity of the cooled cooling air system.

12. The method according to claim 1, characterized in that Wherein modifying the cooling capacity of the cooled cooling air system includes modifying the amount of CCA airflow provided to the turbine section of the aviation gas turbine engine in response to received data indicating the environmental condition of the aviation gas turbine engine, data indicating the degradation parameter of the aviation gas turbine engine, data indicating the operating condition of the aviation gas turbine engine, or a combination thereof.

13. An aviation gas turbine engine, characterized in that: include: a compressor section, a combustor section, and a turbine section arranged in serial flow order; a cooled cooling air system including a CCA heat exchanger in airflow communication with the compressor section and the turbine section, the cooled cooling air system being configured to receive airflow from the compressor section and provide CCA airflow to the turbine section; and a controller operatively connected to the cooled cooling air system, the controller being configured to: receiving data indicative of an environmental condition of the aircraft gas turbine engine, data indicative of a degradation parameter of the aircraft gas turbine engine, data indicative of an operating condition of the aircraft gas turbine engine, or a combination thereof; as well as modifying an amount of cooling capacity of the CCA heat exchanger in response to received data indicative of the ambient condition of the aircraft gas turbine engine, data indicative of the degradation parameter of the aircraft gas turbine engine, data indicative of an operating condition of the aircraft gas turbine engine, or a combination thereof; wherein modifying the cooling capacity of the cooled cooling air system comprises modifying the cooling capacity of a cooling fluid provided to a CCA heat exchanger of the cooled cooling air system; and Wherein modifying the cooling capacity of the cooling fluid provided to the CCA heat exchanger of the cooled cooling air system includes modifying aspects other than a temperature of a cooling fluid flow provided to the CCA heat exchanger of the cooled cooling air system.

14. An aviation gas turbine engine, characterized in that: include: a compressor section, a combustor section, and a turbine section arranged in serial flow order; a cooled cooling air system including a CCA heat exchanger in airflow communication with the compressor section and the turbine section, the cooled cooling air system being configured to receive airflow from the compressor section and provide CCA airflow to the turbine section; and a controller operatively connected to the cooled cooling air system, the controller being configured to: receiving data indicative of a degradation parameter of the aircraft gas turbine engine; as well as modifying an amount of cooling capacity of the CCA heat exchanger in response to receiving data indicative of the degradation parameter of the aircraft gas turbine engine; Wherein modifying the cooling capacity of the cooled cooling air system includes modifying the cooling capacity of a cooling fluid provided to a CCA heat exchanger of the cooled cooling air system.

15. The aviation gas turbine engine according to claim 14, characterized in that: Further including: A thermal bus, wherein the CCA heat exchanger is integrated into the thermal bus, wherein the thermal bus further includes a radiator heat exchanger, and wherein the step of modifying further includes modifying a flow of thermal fluid within the thermal bus to the radiator heat exchanger, a temperature of the thermal fluid flow within the thermal bus, or both.

16. The aviation gas turbine engine according to claim 14, characterized in that: Wherein modifying the cooling capacity of the cooling fluid provided to the CCA heat exchanger of the cooled cooling air system includes modifying a flow of cooling fluid provided to the CCA heat exchanger of the cooled cooling air system.

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