Gas turbine engine with cooling system

Abstract

A gas turbine engine is provided. In one embodiment, a gas turbine engine includes a turbine including a compressor section, a combustion section defining a compressor discharge chamber, and a turbine section, the compressor section, the combustion section, and the turbine section collectively defining, in part, a working gas flow path, the turbine further including: a reverse bleed system, the reverse deflation system comprises a reverse deflation pipeline and an RBS blower in fluid communication with the reverse deflation pipeline, and the reverse deflation pipeline is in fluid communication with the working gas flow path; and an active clearance control (ACC) system including an inlet, a heat transfer assembly disposed around the turbine of the turbine section, and an ACC conduit assembly extending from the inlet to the heat transfer assembly, the inlet of the ACC system being in fluid communication with the reverse bleed conduit.

Description

TECHNICAL FIELD

[0001] The present disclosure relates to gas turbine engines, and more particularly, to gas turbine engines having one or more cooling systems. BACKGROUND

[0002] Gas turbine engines generally include a turbine and a rotor assembly. Gas turbine engines, such as turbofan engines, can be used for aircraft propulsion. In the case of turbofan engines, the rotor assembly can be configured as a fan assembly. As gas turbine engines are driven to higher efficiency and performance metrics, certain temperatures in the gas turbine engine can be increasing. The materials forming components exposed to these temperatures can strain from these increases. Accordingly, cooling these components is important for modern gas turbine engines. BRIEF DESCRIPTION OF DRAWINGS

[0003] A complete and enabling disclosure of the present disclosure, including the best mode thereof, to one of ordinary skill in the art is set forth in the specification, which is to be taken in conjunction with the appended drawings, wherein:

[0004] Figure 1 is a cross-sectional view of a gas turbine engine according to an exemplary aspect of the present disclosure.

[0005] Figure 2 is a more detailed cross-sectional view of the exemplary gas turbine engine of Figure 1

[0006] Figure 3 is a more detailed cross-sectional view of the exemplary gas turbine engine of Figure 2

[0007] Figure 4 is a schematic view of a CCA system according to an exemplary aspect of the present disclosure.

[0008] Figure 5 is a cross-sectional view of a gas turbine engine according to another exemplary aspect of the present disclosure.

[0009] Figure 6 is a cross-sectional view of a gas turbine engine according to another exemplary aspect of the present disclosure.

[0010] Figure 7 is a cross-sectional view of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0011] Figure 8 is a perspective schematic view of the exemplary gas turbine engine of Figure 7

[0012] Figure 9 ​​​is a close-up view of the inlet of a cold side bleed air assembly of a CCA system according to an exemplary aspect of the present disclosure.

[0013] Figure 10 is a close-up view of the inlet of a cold side bleed air assembly of a CCA system according to another exemplary aspect of the present disclosure.

[0014] Figure 11 is a perspective schematic diagram of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0015] Figure 12 is a cross-sectional view of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0016] Figure 13 is a cross-sectional view of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0017] Figure 14 is a cross-sectional view of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0018] Figure 15 According to an exemplary aspect of the present disclosure Figure 14 Schematic diagram of the CCA system and transient bleed system.

[0019] Figure 16 is a cross-sectional view of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0020] Figure 17 According to an exemplary aspect of the present disclosure Figure 16 Schematic diagram of the CCA system and transient bleed system.

[0021] Figure 18 is a cross-sectional view of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0022] Figure 19 is a cross-sectional view of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0023] Figure 20 is a perspective view of a valve assembly of a CCA system according to an exemplary aspect of the present disclosure.

[0024] Figure 21 yes Figure 20 A close-up view of the valve assembly.

[0025] Figure 22 yes Figure 20 Cross-sectional view of the valve assembly.

[0026] Figure 23is a perspective view of a valve assembly of a CCA system according to another example aspect of the disclosure.

[0027] Figure 24 is a perspective view of a valve assembly of a CCA system according to another example aspect of the disclosure.

[0028] Figure 25 is a method of operating a valve assembly of a gas turbine engine according to an example aspect of the disclosure.

[0029] Figure 26 is a cross-sectional view of a gas turbine engine according to yet another example aspect of the disclosure.

[0030] Figure 27 is a cross-sectional view of a gas turbine engine according to yet another example aspect of the disclosure.

[0031] Figure 28 is a cross-sectional view of a gas turbine engine according to yet another example aspect of the disclosure.

[0032] Figure 29 is Figure 28 a close-up view of an example gas turbine engine according to the disclosure.

[0033] Figure 30 is a cross-sectional view of a gas turbine engine according to yet another example aspect of the disclosure.

[0034] Figure 31 is a cross-sectional view of a gas turbine engine according to yet another example aspect of the disclosure.

[0035] Figure 32 is a cross-sectional view of a gas turbine engine according to yet another example aspect of the disclosure.

[0036] Figure 33 is a cross-sectional view of a gas turbine engine according to yet another example aspect of the disclosure.

[0037] Figure 34 is a flowchart of a method of operating a gas turbine engine according to an example aspect of the disclosure.

[0038] Figure 35 is a cross-sectional view of a gas turbine engine according to yet another example aspect of the disclosure.

[0039] Figure 36 is a cross-sectional view of a gas turbine engine according to yet another example aspect of the disclosure.

[0040] Figure 37 is a cross-sectional view of a gas turbine engine according to yet another example aspect of the disclosure.

[0041] Figure 38 is a cross-sectional view of a gas turbine engine according to yet another example aspect of the disclosure.

[0042] Figure 39 is a schematic view of an accessory system cooling system according to an example aspect of the disclosure.

[0043] Figure 40 is a cross-sectional view of a gas turbine engine according to yet another example aspect of the disclosure.

[0044] Figure 41 is a cross-sectional view of a gas turbine engine according to yet another example aspect of the disclosure. DETAILED DESCRIPTION

[0045] Reference will now be made in detail to present embodiments of the disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations

[0046] The word “exemplary” is used herein to mean “serving as an example, instance, or illustration.” Any implementation described herein as “exemplary” is not necessarily to be construed as preferred or advantageous over other implementations. Furthermore, unless otherwise explicitly provided, none of the disclosure herein is intended to be dependent on the disclosure of any other

[0047] The singular forms “a,” “an,” and “the” include plural referents unless the context clearly dictates otherwise.

[0048] The phrases “from X to Y” and “between X and Y” refer to a range of values including the endpoints (e.g., a range including both X and Y).

[0049] The term “turbomachine” refers to a machine that includes one or more compressors, a heat generating section (e.g., a combustion section), and one or more turbines that together generate a torque output.

[0050] The term “gas turbine engine” refers to an engine having a turbomachine as all or a portion of its power source. Example gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, and the like, as well as hybrid electric versions of one or more of these engines.

[0051] The term “combustion section” refers to any heat addition system for a turbine engine. For example, the term combustion section can refer to a section that includes one or more of a deflagration combustion assembly, a rotary detonation combustion assembly, a pulse detonation combustion assembly, or other suitable heat addition assembly. In certain example embodiments, a combustion section can include an annular combustor, a can combustor, a tubular combustor, a trapped vortex combustor (TVC), or other suitable combustion system, or combinations thereof.

[0052] The terms “low” and “high,” or their respective comparative forms (e.g., more “low” and more “high,” where applicable), when used in conjunction with a compressor, turbine, shaft, or spool component, or the like, refer to relative speeds within an engine, unless otherwise noted. For example, “low turbine” or “low speed turbine” defines a component that is configured to operate at a lower rotational speed (e.g., maximum allowable rotational speed) than a “high turbine” or “high speed turbine” of the engine.

[0053] The terms “forward” and “aft” 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, with respect to a gas turbine engine, forward refers to a position closer to the engine inlet, and aft refers to a position closer to the engine nozzle or exhaust.

[0054] The terms “upstream” and “downstream” refer to relative directions with respect to the flow of fluid in a fluid path. For example, “upstream” refers to the direction from which fluid flows, and “downstream” refers to the direction to which fluid flows.

[0055] The term “rated speed,” as used herein with respect to a gas turbine engine, refers to the maximum rotational speed that the gas turbine engine can achieve during normal operation. For example, during maximum load operation, such as during takeoff operation, the gas turbine engine can operate at rated speed.

[0056] The term “standard day operating conditions” refers to ambient conditions of sea level altitude, 59 degrees Fahrenheit, and 60% relative humidity.

[0057] The term “fan pressure ratio,” as used herein with respect to a plurality of fan blades of a fan, refers to the ratio of the air pressure directly downstream of the fan blades during operation of the fan to the air pressure directly upstream of the fan blades of the fan during operation of the fan.

[0058] The terms “coupled,” “fixed,” “attached to,” and the like, unless stated otherwise, refer to both direct coupling, fixation, or attachment and indirect coupling, fixation, or attachment through one or more intermediary components or features, unless stated otherwise.

[0059] As used herein, the term "cruise conditions" refers to a flight phase in which the aircraft is at a level altitude after the climb phase and before the descent to the approach phase. In various examples, cruise conditions can be conducted at a cruise altitude of up to about 65,000 feet (ft). In certain examples, the cruise altitude is between about 28,000 ft and about 45,000 ft. In other examples, the cruise altitude is expressed in terms of flight levels (FL) based on sea level standard atmospheric pressure, where the cruise conditions are between FL 280 and FL 650. In another example, the cruise conditions are between FL 280 and FL 450. In yet other examples, the cruise altitude is defined based at least on atmospheric pressure, where based on a sea level pressure of about 14.70 psia and a sea level temperature of about 59 degrees Fahrenheit, the cruise altitude is between about 4.85 pounds per square inch absolute (psia) and about 0.82 psia. In another example, the cruise altitude is between about 4.85 psia and about 2.14 psia. It will be appreciated that in certain examples, the cruise altitude range defined by pressure can be adjusted based on different reference sea level pressures and / or sea level temperatures.

[0060] As used herein, the term "cruise conditions" refers to a flight phase in which the aircraft is at a level altitude after the climb phase and before the descent to the approach phase. In various examples, cruise conditions can be conducted at a cruise altitude of up to about 65,000 feet (ft). In certain examples, the cruise altitude is between about 28,000 ft and about 45,000 ft. In other examples, the cruise altitude is expressed in terms of flight levels (FL) based on sea level standard atmospheric pressure, where the cruise conditions are between FL 280 and FL 650. In another example, the cruise conditions are between FL 280 and FL 450. In yet other examples, the cruise altitude is defined based at least on atmospheric pressure, where based on a sea level pressure of about 14.70 psia and a sea level temperature of about 59 degrees Fahrenheit, the cruise altitude is between about 4.85 pounds per square inch absolute (psia) and about 0.82 psia. In another example, the cruise altitude is between about 4.85 psia and about 2.14 psia. It will be appreciated that in certain examples, the cruise altitude range defined by pressure can be adjusted based on different reference sea level pressures and / or sea level temperatures.

[0061] The present disclosure generally relates to a gas turbine engine having a cooling air system for cooling one or more hot components of the gas turbine engine. In particular, as gas turbine engines are being driven to higher efficiency and performance metrics, certain temperatures in the gas turbine engine can be increasing. The materials forming the components exposed to these temperatures can strain from these increases. Accordingly, cooling these components is very important for modern gas turbine engines.

[0062] Certain attempts to cool these components have used pressurized airflow from the compressor section and directed the pressurized airflow to the hot components. However, the inventors of the present disclosure have found that these pressurized airflows are either too hot to prove useful or do not have enough pressure to flow through, for example, a first stage high pressure turbine rotor blade and the ambient pressure is too high.

[0063] The present disclosure addresses these issues via cooling a high pressure airflow from, for example, a compressor discharge cavity of a gas turbine engine with a cooled cooling air heat exchanger utilizing a low pressure airflow from a cold location of the gas turbine engine. After receiving heat from the high pressure airflow, the low pressure airflow can be exhausted through an undercowl region of the gas turbine engine in a manner that avoids damaging components of the turbine, avoids antagonizing other systems of the turbine, etc.

[0064] In particular, the present disclosure provides a gas turbine engine having a turbine including a compressor section having a low pressure compressor and a high pressure compressor, a combustion section defining a compressor discharge cavity, and a turbine section. The turbine defines a working gas flow path, and further includes a cooled cooling air (CCA) system including a cold side bleed assembly defining an inlet in fluid communication with a cold location of the gas turbine engine, a CCA heat exchanger in thermal communication with the cold side bleed assembly downstream of the inlet of the cold side bleed assembly, and a hot side bleed assembly defining an inlet in fluid communication with the working gas flow path at the compressor discharge cavity. The hot side bleed assembly is in thermal communication with the CCA heat exchanger to cool an airflow through the hot side bleed assembly. The hot side bleed assembly is also in thermal communication with hot components of the turbine to cool the hot components of the turbine.

[0065] Such a configuration can allow the gas turbine engine to operate more efficiently and / or achieve improved performance metrics.

[0066] Reference is now made to the drawings, wherein like numerals refer to like elements throughout, Figure 1 is a schematic cross-sectional view of a gas turbine engine in accordance with an example embodiment of the present disclosure. More particularly, for Figure 1 Embodiments of the present disclosure, the gas turbine engine is a high-bypass turbofan jet engine, sometimes referred to simply as a "turbofan engine." As shown, Figure 1 The gas turbine engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 provided for reference), a radial direction R, and a circumferential direction C extending about the longitudinal centerline 12. Generally, the gas turbine engine 10 includes a fan section 14 and a turbine 16 disposed downstream from the fan section 14.

[0067] The depicted exemplary turbine 16 generally includes a substantially tubular outer casing 18 defining an annular inlet 20. The outer casing 18 surrounds, in serial flow relationship: a compressor section including a booster or low pressure (LP) compressor 22 and a high pressure (HP) compressor 24; a combustion section 26; a turbine section including a high pressure (HP) turbine 28 and a low pressure (LP) turbine 30; and an ejection exhaust nozzle section 32. A high pressure (HP) shaft 34, which can additionally or alternatively be a wire shaft, drivingly connects the HP turbine 28 to the HP compressor 24. A low pressure (LP) shaft 36, which can additionally or alternatively be a wire shaft, drivingly connects the LP turbine 30 to the LP compressor 22. The compressor section, the combustion section 26, the turbine section, and the ejection exhaust nozzle section 32 together define a working gas flow path 37.

[0068] For the described embodiment, the fan section 14 includes a fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced apart manner. As depicted, the fan blades 40 generally extend outwardly from the disk 42 in a radial direction R. Each fan blade 40 is rotatable about a pitch axis P relative to the disk 42 by virtue of the fan blades 40 being operably coupled to a suitable pitch mechanism 44 configured to collectively change the pitch of the fan blades 40, e.g., uniformly. The gas turbine engine 10 also includes a power gear box 46, and the fan blades 40, disk 42, and pitch mechanism 44 are rotatable together about the longitudinal centerline 12 across the power gear box 46 by the LP shaft 36. The power gear box 46 includes a plurality of gears for adjusting a rotational speed of the fan 38 relative to a rotational speed of the LP shaft 36 such that the fan 38 can rotate at a more efficient fan speed.

[0069] For example, when the engine is operating under cruise conditions, including the power gear box 46 can allow the fan 38 to define a relatively low fan pressure ratio. For example, when the gas turbine engine 10 is operating under cruise conditions, the fan 38 can define a fan pressure ratio of less than or equal to 1.6.

[0070] Still referring to Figure 1 the exemplary embodiment, the disk 42 is covered by a rotatable forward hub 48 (sometimes also referred to as a "spinner") of the fan section 14. The forward hub 48 is aerodynamically shaped to promote airflow through the plurality of fan blades 40.

[0071] Additionally, the exemplary fan section 14 includes an annular fan casing or outer nacelle 50 that circumferentially surrounds at least a portion of the fan 38 and / or the turbine 16. It should be appreciated that in the depicted embodiment, the nacelle 50 is supported relative to the turbine 16 by a plurality of circumferentially spaced apart outlet guide vanes 52. Moreover, a downstream section 54 of the nacelle 50 extends over an outer portion of the turbine 16 so as to define a bypass passage 56 therebetween.

[0072] During operation of the gas turbine engine 10, an amount of air 58 enters the gas turbine engine 10 through the nacelle 50 and associated inlet 60 of the fan section 14. As the amount of air 58 passes through the fan blades 40, a first portion of air 62 is directed or channeled into the bypass passage 56 and a second portion of air 64 is directed or channeled into the working gas flow path 37, or more specifically, into the LP compressor 22, as indicated by the arrows. The ratio between the first portion of air 62 and the second portion of air 64 is commonly referred to as the bypass ratio. The pressure of the second portion of air 64 is then increased as it is channeled through the HP compressor 24 and into the combustion section 26, where it is mixed with fuel and combusted to provide combustion gases 66. Notably, for the depicted embodiment, the gas turbine engine 10 is configured to define a relatively high overall pressure ratio when operated at a rated speed during standard day operating conditions. In particular, the gas turbine engine 10 is configured to define an overall pressure ratio that is greater than or equal to 50: 1 and less than or equal to 70: 1 when operated at a rated speed during standard day operating conditions.

[0073] The combustion gases 66 are directed through the HP turbine 28, where a portion of thermal and / or kinetic energy from the combustion gases 66 is extracted via sequential stages of HP turbine stator vanes 68 coupled to the outer casing 18 and HP turbine rotor blades 70 coupled to the HP shaft 34, thus rotating the HP shaft 34 in support of operation of the HP compressor 24. The combustion gases 66 are then directed through the LP turbine 30, where a second portion of thermal and kinetic energy is extracted from the combustion gases 66 via sequential stages of LP turbine stator vanes 72 coupled to the outer casing 18 and LP turbine rotor blades 74 coupled to the LP shaft 36, thus rotating the LP shaft 36 in support of operation of the LP compressor 22 and / or rotation of the fan 38.

[0074] The combustion gases 66 are subsequently directed through the ejection exhaust nozzle section 32 of the turbine engine 16 to provide propulsive thrust. Substantially simultaneously, the pressure of the first portion of air 62 is increased as it is directed through the bypass passage 56 prior to being exhausted from the fan nozzle exhaust section 76 of the gas turbine engine 10, also providing propulsive thrust. The HP turbine 28, the LP turbine 30, and the ejection exhaust nozzle section 32 at least partially define a hot gas path 78 for directing the combustion gases 66 through the turbine engine 16.

[0075] For the depicted example embodiment, the gas turbine engine further includes a cooled cooling air system to provide a cooling airflow that can accept heat from a hot gas flow of the gas turbine engine that is subsequently provided to a hot component of the gas turbine engine. Various example embodiments of the cooled cooling air system are described in greater detail below.

[0076] However, it should be understood that Figure 1 The exemplary gas turbine engine 10 depicted is by way of example only, and in other exemplary embodiments, the gas turbine engine 10 can have any other suitable configuration. For example, although the depicted gas turbine engine 10 is configured as a ducted gas turbine engine (e.g., including the outer nacelle 50), in other embodiments, the gas turbine engine 10 can be a non-ducted gas turbine engine (such that the fan 38 is a non-ducted fan, and the outlet guide vanes 52 are cantilevered from the outer casing 18). Additionally or alternatively, although the depicted gas turbine engine 10 is configured as a geared gas turbine engine (e.g., including the power gear box 46) and a variable pitch gas turbine engine (e.g., including the fan 38 configured as a variable pitch fan), in other embodiments, the gas turbine engine 10 can additionally or alternatively be configured as a direct drive gas turbine engine (such that the LP shaft 36 rotates at the same speed as the fan 38), a fixed pitch gas turbine engine (such that the fan 38 includes fan blades 40 that are not rotatable about a pitch axis P), or both. It should also be understood that aspects of the present disclosure can be incorporated into any other suitable gas turbine engine in other exemplary embodiments. For example, in other exemplary embodiments, aspects of the present disclosure can be incorporated into, as the case can be, a turboprop gas turbine engine, a turboshaft gas turbine engine, or a turbojet gas turbine engine.

[0077] Reference is now made to Figure 2 , providing a close-up schematic view of the gas turbine engine 10 according to exemplary embodiments of the present disclosure. Figure 2 The exemplary gas turbine engine 10 of Figure 1 may be configured in a similar manner as the exemplary gas turbine engine 10 described above with reference to

[0078] For example, Figure 2 The exemplary gas turbine engine 10 generally includes a fan section 14 and a turbomachine 16. The turbomachine 16 includes a compressor section having an LP compressor 22 and an HP compressor 24, a combustion section 26, and a turbine section having an HP turbine 28 and an LP turbine 30. The turbomachine 16, and more specifically the compressor section, the combustion section 26, and the turbine section, at least partially define a working gas flow path 37 therethrough. Further, the turbomachine 16 includes a casing, also referred to as an outer casing 18, which encloses the compressor section, the combustion section 26, and the turbine section, and defines an undercowl 104 along a radial direction R between the working gas flow path 37 and the outer casing 18.

[0079] Figure 2The exemplary gas turbine engine 10 also includes a bifurcation extending from the turbine 16. In particular, the bifurcation extends between the turbine 16 and the outer nacelle 50 of the gas turbine engine 10. In certain exemplary embodiments, such as the depicted exemplary embodiment, the bifurcation can be more specifically configured as an upper bifurcation 106 of the gas turbine engine 10.

[0080] It will be appreciated that, as used herein, the term "bifurcation" generally refers to an airfoil extending from the turbine 16 through the airflow above the turbine 16. In certain exemplary embodiments, the bifurcation can be a structural pylon extending from the turbine 16. In certain embodiments, the bifurcation can be the upper bifurcation 106 as described above, or alternatively can be a lower bifurcation.

[0081] Further, Figure 2 The exemplary gas turbine engine 10 also includes a CCA system 100. The CCA system 100 includes a cold side bleed assembly 108, a hot side bleed assembly 110, and a CCA heat exchanger 112.

[0082] The cold side bleed assembly 108 defines an inlet 114 in fluid communication with a cold location of the gas turbine engine 10 to receive a cooling airflow 102 from the cold location of the gas turbine engine 10. As used herein, the term "cold location" refers to any location from which fluid can be extracted at a temperature lower than the airflow through the hot side bleed assembly 110. In particular, for the illustrated embodiment, the cold location is the working gas flowpath 37 through the compressor section at a location through the LP compressor 22, between the LP compressor 22 and the HP compressor 24, or both. In this manner, the inlet 114 of the cold side bleed assembly 108 is in fluid communication with the working gas flowpath 37 at such location.

[0083] More specifically, for the depicted embodiment, the inlet 114 of the cold side bleed assembly 108 is in direct fluid communication with the working gas flowpath 37, and yet more specifically, at a location downstream of the inlet of the LP compressor 22 and upstream of the outlet of the LP compressor 22, with the working gas flowpath 37 through the LP compressor 22. In short, as used herein, the term "direct fluid communication" refers to directly receiving an airflow, but excluding any intervening accessory systems directly supporting such fluid communication (as discussed below in comparison to the embodiment of FIG. 2). Figure 6

[0084] Including the inlet 114 of the cold side bleed assembly 108 at such a location can allow the cold side bleed assembly 108 to receive a relatively cold airflow having sufficient pressure to drive the airflow through the cold side bleed assembly 108, allowing for a relatively high pressure drop within the CCA heat exchanger 112 to enable efficient heat transfer.

[0085] ​As will be discussed in greater detail below, the cold-side bleed assembly 108 generally includes one or more conduits 116 to provide the cooling gas stream 102 from the inlet 114 to the CCA heat exchanger 112, and from the CCA heat exchanger 112. The conduits 116 can be separate pipes (e.g., tubes, conduits, etc.), can be flow passages defined by the structure of the turbine 16, or combinations thereof.

[0086] Briefly, the cold-side bleed assembly 108 also defines an outlet 118 at a location downstream of the CCA heat exchanger 112. In this manner, it will be appreciated that the CCA heat exchanger 112 is in thermal communication with the cold-side bleed assembly 108 downstream of the inlet 114 of the cold-side bleed assembly 108 and upstream of the outlet 118 of the cold-side bleed assembly 108.

[0087] Still referring to the embodiment of FIG. 1, Figure 2 the CCA heat exchanger 112 is located in the undercowl cavity 104 of the turbine 16. In this manner, the amount of piping and tubing required for the cold-side bleed assembly 108 (e.g., the conduits 116) and the hot-side bleed assembly 110 can be relatively small.

[0088] However, it is noted that in other example embodiments, the CCA heat exchanger 112 can instead be positioned at other suitable locations. For example, the upper bifurcation 106 can define a bifurcation cavity 120. In certain example embodiments, the CCA heat exchanger 112 can be located in the bifurcation cavity 120, or in both the bifurcation cavity 20 and the undercowl cavity 104 (as depicted by the dashed lines in FIG. 1). Figure 2

[0089] The hot-side bleed assembly 110 similarly defines an inlet 122 (see Figure 3 ) in fluid communication with the working gas flow path 37 to receive a gas stream from the working gas flow path 37. The hot-side bleed assembly 110 is also in thermal communication with the CCA heat exchanger 112. The CCA heat exchanger 112 is configured to cool the gas stream passing through the hot-side bleed assembly 110, transferring heat from the gas stream to the cooling gas stream 102. In this manner, the gas stream passing through the hot-side bleed assembly 110 can be referred to herein as a cooled gas stream 124.

[0090] The hot-side bleed assembly 110 is also in thermal communication with hot components of the turbine 16 to cool the hot components of the turbine 16 (e.g., configured to provide the cooled gas stream 124 over the hot components, into the hot components, or otherwise to a location in thermal communication with the hot components). The flow of the cooled gas stream 124 to the hot components will be described in greater detail below with reference to Figure 3 .

[0091] Still referring to the embodiment of FIG. 1, Figure 2 ​The CCA system 100 further includes a first flow control valve 126 in operable communication with the cold side bleed assembly 108, and a second flow control valve 128 in operable communication with the hot side bleed assembly 110. The first and second flow control valves 126, 128 can each regulate the amount of gas flow through the respective system. In particular, in the embodiment depicted, Figure 2 the first flow control valve 126 can regulate the amount of cooling gas flow 102 through the cold side bleed assembly 108, and the second flow control valve 128 can regulate the amount of cooling gas flow 102 through the hot side bleed assembly 110, the first flow control valve 126 being upstream of the CCA heat exchanger 112, and the second flow control valve 128 being similarly upstream of the CCA heat exchanger 112.

[0092] Notably, for the depicted embodiment, the gas turbine engine 10 further includes a controller 130. The controller 130 is in operable communication with the first and second flow control valves 126, 128 to control operation of the first and second flow control valves 126, 128, for example, in response to, for example, received data indicative of operating conditions of the gas turbine engine 10.

[0093] In this manner, the controller 130 can regulate the amount of cooling gas flow 102 through the cold side bleed assembly 108 and the amount of cooling gas flow 124 through the hot side bleed assembly 110 to between 0% and 100% of their respective maximum gas flows, or to one or more intermediate flow levels. Notably, in certain example embodiments, the CCA system 100 can be configured to receive an amount of cooling gas flow 102 (referred to as W25) that is greater than or equal to 0.3% of the core mass flow of air entering the compressor section and less than or equal to 13.5% of the core mass flow of air entering the compressor section during operating conditions of the gas turbine engine 10.

[0094] As noted above, Figure 2 the example controller 130 depicted in the foregoing is configured to receive data sensed from one or more sensors, and may, for example, base control decisions for the CCA system 100 on the received data.

[0095] In one or more example embodiments, Figure 2 the controller 130 depicted in the foregoing can be a standalone controller 130 of the CCA system 100, or alternatively, can be integrated into one or more of a controller of the gas turbine engine 10 with which the CCA system 100 is integrated, a controller of an aircraft including the gas turbine engine 10 with which the CCA system 100 is integrated, etc.

[0096] With particular reference to the operation of the controller 130, in at least certain embodiments, the controller 130 can include one or more computing devices 132. The computing device(s) 132 can include one or more processor(s) 132A and one or more memory device(s) 132B. The one or more processor(s) 132A can include any suitable processing device, such as a microprocessor, microcontroller, integrated circuit, logic device, and / or other suitable processing device. The one or more memory device(s) 132B can include one or more computer-readable media, including, but not limited to, non-transitory computer-readable media, RAM, ROM, hard drives, flash drives, and / or other memory devices.

[0097] The one or more memory device(s) 132B can store information accessible by the one or more processor(s) 132A, including computer-readable instructions 132C that can be executed by the one or more processor(s) 132A. The instructions 132C can be any set of instructions that, when executed by the one or more processor(s) 132A, cause the one or more processor(s) 132A to perform operations. In some embodiments, the instructions 132C can be executed by the one or more processor(s) 132A to cause the one or more processor(s) 132A to perform operations such as any operations and functionalities for which the controller 130 and / or computing device(s) 132 are configured, operations for operating the CCA system 100 as described herein (e.g., the methods 300, 600), and / or any other operations or functionalities of the one or more computing device(s) 132. The instructions 132C can be software written in any suitable programming language or can be implemented in hardware. Additionally, and / or alternatively, the instructions 132C can be executed in logically and / or virtually separate threads on the one or more processor(s) 132A. The one or more memory device(s) 132B can also store data 132D that is accessible by the one or more processor(s) 132A. For example, the data 132D can include data indicative of engine / aircraft operating conditions, and / or any other data and / or information described herein.

[0098] The computing device 132 can also include a network interface 132E for communicating with, for example, other components of the CCA system 100 (e.g., the valves 126, 128), in conjunction with a gas turbine engine of the CCA system 100, an aircraft in conjunction with the gas turbine engine, etc. For example, in the depicted embodiment, as noted above, the gas turbine engine and / or the CCA system 100 include one or more sensors for sensing data indicative of one or more parameters of the gas turbine engine, the CCA system 100, or both. The controller 130 of the CCA system 100 is operably coupled to the one or more sensors by, for example, the network interface, such that the controller 130 can receive data indicative of various operating parameters sensed by the one or more sensors during operation. Further, for the illustrated embodiment, the controller 130 is operably coupled to, for example, the first flow control valve 126 and the second flow control valve 128. In this manner, the controller 130 can be structured to regulate the amount of cooling airflow 102 through the cold side bleed air assembly 108 and the amount of cooled airflow 124 through the hot side bleed air assembly 110 in response to, for example, data sensed by the one or more sensors.

[0099] The network interface 132E can include any suitable components for interfacing with one or more networks, including, for example, transmitters, receivers, ports, controllers, antennas, and / or other suitable components.

[0100] The technology discussed herein makes reference to computer-based systems, actions taken by and information sent to and from computer-based systems. One of ordinary skill in the art will recognize that the inherent flexibility of computer-based systems allows for a great variety of possible configurations, combinations, and divisions of tasks and

[0101] Referring now to Figure 3 , there is provided Figure 2FIG. 1 is a schematic illustration of a gas turbine engine 10 in accordance with at least some example embodiments. As discussed above, the hot-side bleed assembly 110 includes an inlet 122 that is in fluid communication with the working gas flow path 37 of the turbine 16. More specifically, the combustion section 26 includes a combustor 134 that defines a combustion chamber 136, and also defines a compressor discharge plenum 138 downstream of the compressor section (e.g., downstream of the HP compressor 24) and upstream of the combustion chamber 136. The inlet 122 of the hot-side bleed assembly 110 is in fluid communication with the compressor discharge plenum 138 to receive the cooled gas stream 124 from the compressor discharge plenum 138. In this manner, the hot-side bleed assembly 110 can receive a relatively high pressure gas stream, thereby allowing the high pressure gas stream to be provided to various hot components of the turbine 16.

[0102] It will be appreciated, however, that in other example embodiments, the inlet 122 can additionally or alternatively be in fluid communication with the working gas flow path 37 through the compressor section, such as through the HP compressor 24. For example, in other example embodiments, the inlet 122 can additionally or alternatively be in fluid communication with the working gas flow path 37 through a downstream half of the HP compressor 24 (see, e.g., the inlet 242 of the transient bleed assembly 238 discussed below with reference to FIG. 2; the compressor bleed plenum 506 discussed below with reference to FIG. 5). Figure 14 Figure 31 It will be appreciated, however, that in other example embodiments, the inlet 122 can additionally or alternatively be in fluid communication with the working gas flow path 37 through the compressor section, such as through the HP compressor 24. For example, in other example embodiments, the inlet 122 can additionally or alternatively be in fluid communication with the working gas flow path 37 through a downstream half of the HP compressor 24 (see, e.g., the inlet 242 of the transient bleed assembly 238 discussed below with reference to FIG. 2; the compressor bleed plenum 506 discussed below with reference to FIG. 5).

[0103] In at least some example embodiments, the hot component can be a rotor bore of a turbine of the turbine section (such as the HP turbine rotor bore 140 of the HP turbine 28), an airfoil of the turbine (such as the HP turbine rotor blade 142 of the HP turbine 28), a HP compressor rotor bore 144 of the HP compressor 24, a HP compressor rotor blade 146 of the HP compressor 24, an oil sump 148 within the turbine section (depicted in dashed line), an engine frame (e.g., the turbine mid-frame 150 as shown in Figure 3 FIG. 1), the turbine aft-frame 152 (see, e.g., FIG. 1), or a combination thereof. Figure 2

[0104] For example, in the depicted embodiment, the hot component can be the HP turbine 28, such as the HP turbine rotor bore 140, the HP turbine rotor blade 142, or both. Although not depicted, in certain example embodiments, the cooled gas stream 124 can flow through one or more passages in the HP turbine rotor bore 140 to the HP turbine rotor blade 142, where the cooled gas stream 124 flows through the HP turbine rotor blade 142 and exits the HP turbine rotor blade 142 as a film cooling fluid.

[0105] ​​Additionally or alternatively, in the depicted embodiment, the hot components can be a rear stage (such as a last stage or a penultimate stage) of the HP compressor 24 (e.g., one or more of the HP compressor rotor blades 146, HP compressor stator vanes, HP compressor inter-stage seals, etc., as shown).

[0106] Briefly, it will be appreciated that the combustion section 26 also includes a combustor case 154. The hot-side bleed assembly 110 includes a receiving portion 156 and a delivery portion 158. The receiving portion 156 extends from the inlet 122 to the CCA heat exchanger 112, while the delivery portion 158 extends from the CCA heat exchanger 112 toward the hot components. The inlet 122 of the hot-side bleed assembly 110 is located at (e.g., defined in or inboard of) the combustor case 154. The delivery portion 158 of the hot-side bleed assembly 110 extends through the combustor case 154 and through the compressor discharge cavity 138. The delivery portion 158 can define an outlet for providing the cooled gas stream 124 to a duct or other gas stream delivery path inboard of the compressor discharge cavity 138. For example, in the illustrated embodiment, an inner duct 160 is defined between the HP shaft 34 and an inner combustor case 162 of the combustion section 26. The delivery portion 158 can define an outlet opening into the inner duct 160.

[0107] This configuration can allow for delivery of the cooled gas stream 102 at a desired temperature.

[0108] Reference is now made to Figure 4 , which provides a schematic view of a CCA system 100 according to example embodiments of the present disclosure. Figure 4 The example CCA system 100 of Figure 3 may be configured in a similar manner as the example CCA system 100 of Figure 4 . In particular, The example CCA system 100 of

[0109] In particular, the CCA system 100 further includes a second CCA heat exchanger 112B and a third CCA heat exchanger 112C. The cold side bleed assembly 108 includes a main portion 164 and a plurality of branches downstream of the main portion 164 and arranged in a parallel flow arrangement. In particular, in addition to the main portion 164, the cold side bleed assembly 108 further includes a first portion 166 in thermal communication with the first CCA heat exchanger 112A, a second portion 168 in thermal communication with the second CCA heat exchanger 112B, and a third portion 170 in thermal communication with the third CCA heat exchanger 112C. The first portion 166, the second portion 168, and the third portion 170 are arranged in a parallel flow arrangement.

[0110] Further, the hot side bleed assembly 110 is a first hot side bleed assembly 110A, and Figure 4 The CCA system 100 further includes a second hot side bleed assembly 110B and a third hot side bleed assembly 110C. The first hot side bleed assembly 110A is in thermal communication with the first CCA heat exchanger 112A, the second hot side bleed assembly 110B is in thermal communication with the second CCA heat exchanger 112B, and the third hot side bleed assembly 110C is in thermal communication with the third CCA heat exchanger 112C. The first hot side bleed assembly 110A, the second hot side bleed assembly 110B, and the third hot side bleed assembly 110C are each fluidly isolated from one another.

[0111] In certain example embodiments, the first hot side bleed assembly 110A, the second hot side bleed assembly 110B, and the third hot side bleed assembly 110C are operable to deliver the cooled gas stream 124 to one or more hot components of the turbine 16 in conjunction with the CCA system 100. For example, in one example embodiment, the first hot side bleed assembly 110A can be in thermal communication with a first hot component 172A of the turbine 16 to cool the first hot component 172A; the second hot side bleed assembly 110B can be in thermal communication with a second hot component 172B of the turbine 16 to cool the second hot component 172B; and the third hot side bleed assembly 110C can be in thermal communication with a third hot component 172C of the turbine 16 to cool the third hot component 172C.

[0112] For example, in one example embodiment, the first hot side bleed assembly 110A and the second hot side bleed assembly 110B can be in thermal communication with the HP turbine 28 to provide the cooled gas stream 124 to the HP turbine rotor bore 140 of the HP turbine 28, to the plurality of HP turbine rotor blades 142 of the first stage of the HP turbine 28, or both (see Figure 3 ) For example, in one example embodiment, the first hot side bleed assembly 110A and the second hot side bleed assembly 110B can be in thermal communication with the HP turbine 28 to provide the cooled gas stream 124 to the HP turbine rotor bore 140 of the HP turbine 28, to the plurality of HP turbine rotor blades 142 of the first stage of the HP turbine 28, or both (see

[0113] Reference is now made to Figure 5, a gas turbine engine 10 and a CCA system 100 according to another example embodiment of the present disclosure are provided. Figure 5 The example gas turbine engine 10 and CCA system 100 of Figure 2 may be constructed in a similar manner as the example gas turbine engine 10 and CCA system 100 of , discussed above. Thus, like or similar numbers can refer to like or similar parts.

[0114] Figure 5 For example, The example CCA system 100 of

[0115] includes a cold side bleed assembly 108, a hot side bleed assembly 110, and a CCA heat exchanger 112 in thermal communication with the cold side bleed assembly 108 and the hot side bleed assembly 110 to cool a flow of gas (e.g., a cooled flow of gas 124) through the hot side bleed assembly 110. Further, the CCA system 100 includes a first flow control valve 126 in operable communication with the cold side bleed assembly 108 and a second flow control valve 128 in operable communication with the hot side bleed assembly 110. Notably, however, for the illustrated embodiment, the hot side bleed assembly 110 further includes an upstream portion 174 defining the inlet 122, a first portion 176 in thermal communication with the CCA heat exchanger 112, and a bypass portion 178 that bypasses the CCA heat exchanger 112. The second flow control valve 128 is located at a junction between the upstream portion 174, the first portion 176, and the bypass portion 178. In this manner, the second flow control valve 128 can regulate the amount of cooled flow of gas 124 through the hot side bleed assembly 110 that bypasses the CCA heat exchanger 112. Figure 5 Figure 2 In at least certain example embodiments, Figure 4 the bypass arrangement of may be incorporated into other embodiments of the present disclosure, such as one or more of the hot side bleed assemblies 110 described above with reference to the embodiments of

[0116] , Figure 6 , a gas turbine engine 10 and a CCA system 100 according to another example embodiment of the present disclosure are provided. Figure 6 The example gas turbine engine 10 and CCA system 100 of Figure 2 may be constructed in a similar manner as the example gas turbine engine 10 and CCA system 100 of , discussed above. Thus, like or similar numbers can refer to like or similar parts.

[0117] Figure 6The exemplary CCA system 100 includes a cold side bleed assembly 108, a hot side bleed assembly 110, and a CCA heat exchanger 112 in thermal communication with both the cold side bleed assembly 108 and the hot side bleed assembly 110 to cool the gas stream passing through the hot side bleed assembly 110. The cold side bleed assembly 108 defines an inlet 114 in fluid communication with the working gas flow path 37 passing through the compressor section at a location between, between, or both the LP compressor 22 and the HP compressor 24. In particular, for embodiments of the exemplary CCA system 100, Figure 6 The turbine 16 includes an operable bleed assembly 180 having an operable bleed conduit 182 extending between an inlet 184 in fluid communication with the working gas flow path 37 at a location between the LP compressor 22 and the HP compressor 24 and an outlet 186 in fluid communication with the bypass passage 56 of the gas turbine engine 10. In particular, the inlet 184 of the operable bleed conduit 182 is in fluid communication with the working gas flow path 37 at a location between the LP compressor 22 and the HP compressor 24. The inlet 114 of the cold side bleed assembly 108 is in fluid communication with the operable bleed conduit 182. In this manner, the inlet 114 of the cold side bleed assembly 108 is indirectly in fluid communication with the working gas flow path 37 at a location between the LP compressor 22 and the HP compressor 24.

[0118] Notably, the operable bleed assembly 180 further includes a variable bleed valve 188 in communication with the operable bleed conduit 182 at the inlet 184 of the operable bleed conduit 182. The variable bleed valve 188 is located upstream of the location where the inlet 114 of the cold side bleed assembly 108 meets the operable bleed conduit 182. In this manner, the variable bleed valve 188 can control the cooling gas stream 102 passing through the cold side bleed assembly 108 of the CCA system 100. The variable bleed valve 188 can be in operable communication with a controller, such as the exemplary controller 130 of Figure 2

[0119] Reference is now made to Figure 7 and Figure 8 depicting a gas turbine engine 10 according to another exemplary embodiment. Figure 7 A schematic cross-sectional view of an exemplary gas turbine engine 10 is provided, and Figure 8 A perspective view of an exemplary gas turbine engine 10 with certain components removed for clarity is provided. Figure 7 and Figure 8 The exemplary gas turbine engine 10 of Figures 1 to 6 may be constructed in substantially the same manner as one or more of the exemplary gas turbine engines 10 described above with reference to ​

[0120] For example, Figure 7 and Figure 8 The exemplary gas turbine engine 10 includes a turbine 16 having a shroud 18 (shown in FIG. 1 for clarity), an outer nacelle 50, and an extension. Figure 8 , outer nacelle 50 at least partially surrounds turbine 16 (for clarity, Figure 8 16 and extending outwardly therefrom. More specifically, for the embodiment depicted, the extension is a bifurcated portion extending between the turbine 16 and the outer nacelle 50. Still more specifically, for the embodiment shown, the bifurcated portion is Figure 7 and Figure 8 The upper fork 106 in the embodiment of FIG. Figure 8 (a large portion of a wall is removed).

[0121] The exemplary gas turbine engine 10 also includes a CCA system 100. The exemplary CCA system 100 includes a cold side bleed air assembly 108, a hot side bleed air assembly 110, and a CCA heat exchanger 112 in thermal communication with both the cold side bleed air assembly 108 and the hot side bleed air assembly 110 to cool the airflow passing through the hot side bleed air assembly 110. More specifically, for the illustrated embodiment, the CCA system 100 includes a plurality of CCA heat exchangers 112 and a corresponding plurality of hot side bleed air assemblies 110 in thermal communication with a corresponding one of the plurality of CCA heat exchangers 112 (see FIG. Figure 8 ). This construction can be similar to the above reference Figure 4 The exemplary arrangement described is notable. Figure 7 Only the first CCA heat exchanger 112A and the first hot side bleed air assembly 110A of the CCA system 100 are depicted.

[0122] for Figure 7 and Figure 8 , the cold side bleed assembly 108 defines an inlet 114. However, for the illustrated embodiment, the inlet 114 is positioned to be in fluid communication with the airflow above the extension portion, and more specifically, above the bifurcated portion, and even more specifically, above the upper bifurcated portion 106.

[0123] More specifically, the upper bifurcation 106 includes a flowpath surface 190. In the context of the upper bifurcation 106, the term "flowpath surface" refers to a surface that is exposed to the airflow through the bypass passage 56 during operation of the gas turbine engine 10. In the depicted embodiment, the inlet 114 is positioned on the flowpath surface 190 of the upper bifurcation 106. In this manner, the inlet 114 is configured to receive the airflow through the bypass passage 56 that flows over the upper bifurcation 106. This configuration can result in the provision of cooler air to the CCA system 100, particularly to one or more of the CCA heat exchangers 112 of the CCA system 100.

[0124] Notably, in the depicted embodiment, the upper bifurcation 106 includes a leading edge 192 and a trailing edge 194, as well as a radially inner end 196 and a radially outer end 198. In the depicted embodiment, the inlet 114 is located on the flowpath surface 190 at the leading edge 192 of the upper bifurcation 106 and proximate the radially outer end 198 of the upper bifurcation 106 (e.g., closer to the radially outer end 198 than to the radially inner end 196). This configuration can maximize the pressure of the cooling airflow 102 received by the cold side bleed assembly 108. In particular, this configuration can allow the cold side bleed assembly 108 to receive a substantially free-stream airflow as compared to the boundary layer airflow at other locations.

[0125] It will be appreciated, however, that in other example embodiments, the inlet 114 can be located on the flowpath surface 190 of the bifurcation, or more specifically, the flowpath surface 190 of the upper bifurcation 106, at other suitable locations. For example, as depicted by the dashed lines in Figure 7 In other example embodiments, the inlet 114 can be positioned proximate the radially inner end 196 and between the leading edge 192 and the trailing edge 194. This configuration can reduce the amount of ducting required by the cold side bleed assembly 108 while still providing the cold side bleed assembly 108 with airflow at a reduced boundary layer airflow as compared to other locations exposed to the bypass passage 56.

[0126] Referring now to Figure 9 , a close-up, cross-sectional, top-down view of the inlet 114 of the cold side bleed assembly 108 of the CCA system 100 according to example embodiments of the present disclosure is provided. Similar to the example embodiments depicted in Figure 7 and Figure 8 , the inlet 114 depicted in Figure 9 may be positioned on the flowpath surface 190 of the upper bifurcation 106 of the gas turbine engine 10.

[0127] As shown, the airflow 200 above the bifurcation defines an airflow direction, and the inlet 114 defines an upstream end 202 and a downstream end 204. In the depicted embodiment, the cold side bleed assembly 108 includes a scoop 206 at the inlet 114 to increase the amount of airflow 200 received through the inlet 114 and into the cold side bleed assembly 108 as the cooling airflow 102. In particular, in the illustrated embodiment, the scoop 206 is positioned at the downstream end 204 of the inlet 114 facing the airflow direction. Such a configuration can allow for receiving a desired amount of airflow as the cooling airflow 102.

[0128] It will be appreciated, however, that other suitable arrangements can be provided in other example embodiments. For example, referring now to Figure 10 , another close-up, cross-sectional, top-down view of the inlet 114 of the cold side bleed assembly 108 of the CCA system 100 according to example embodiments of the present disclosure is provided. Figure 10 The example inlet 114 of Figure 9 may be configured in substantially the same manner as the example inlet 114 of Figure 10 . For the example embodiment of Figure 10 , however, the cold side bleed assembly 108 also includes a variable geometry component 208 at the inlet 114. In particular, for the illustrated embodiment, the variable geometry component 208 includes a door 210, an actuator 212, and an extension 214. The actuator 212 is a linear actuator that is operable with the door 210 through the extension 214. The door 210 is movable by the actuator 212 between an open position, a closed position (depicted in dashed lines), and one or more intermediate positions (shown) in the airflow direction of the airflow 200. In this manner, it will be appreciated that the variable geometry component 208 can control the mass flow rate of the cooling airflow 102 through the cold side bleed assembly 108. Thus, the variable geometry component can be referred to as a flow control valve of the cold side bleed assembly 108 (see, e.g., the valve 126 in Figure 2 .

[0129] Although not depicted, in certain example embodiments, the actuator 212 of the variable geometry component 208 can be operably coupled to a controller, such as the example controller 130 of Figure 2 .

[0130] It will be further appreciated that any other suitable variable geometry component can be provided in other example embodiments. For example, in other example embodiments, the variable geometry component 208 can include a plurality of sequentially arranged vanes, slats, louvers, or doors (e.g., one or more hinged doors) that are each rotatable about a respective pivot between an open position, a closed position, and one or more intermediate positions.

[0131] Referring now to Figure 11FIG. 1 provides a partial perspective view of a gas turbine engine 10 and a CCA system 100 according to another example embodiment of the present disclosure. Figure 11 The example gas turbine engine 10 and CCA system 100 can be constructed in substantially the same manner as the example gas turbine engine 10 and CCA system 100 described above with reference to Figure 7 and Figure 8 .

[0132] For example, Figure 11 The example CCA system 100 includes a cold side bleed assembly 108 defining an inlet 114 positioned in fluid communication with the airflow over the split of the gas turbine engine 10. Notably, however, for the example embodiment, the inlet 114 is not positioned on the flowpath surface 190 of the split. Rather, the inlet 114 is positioned to indirectly receive a portion of the airflow over the split. Figure 11

[0133] More specifically, the example gas turbine engine 10 further includes an accessory system cooling system 216 having a cooling system inlet 218 located on the split and a duct 220 in fluid communication with the cooling system inlet 218. The cooling system inlet 218 of the accessory system cooling system 216 is positioned on the flowpath surface 190 of the split. The inlet 114 of the cold side bleed assembly 108 is in fluid communication with the duct 220 of the accessory system cooling system 216 at a location downstream of the cooling system inlet 114.

[0134] In certain example embodiments, the accessory system cooling system 216 can be, for example, an environmental control system for an aircraft including the gas turbine engine 10. Notably, for the example embodiment, the accessory system cooling system 216 further includes a dedicated heat exchanger 222, such as an environmental control system precooler, configured to cool the airflow through the duct 220 of the accessory system cooling system 216.

[0135] Further, for the example embodiment, the CCA system 100 includes a flow control valve 224 located at the inlet 114 of the cold side bleed assembly 108. For the example embodiment, the flow control valve 224 is also operable with the airflow through the duct 220 of the accessory system cooling system 216 to control the shunting of airflow between the duct 220 of the accessory system cooling system 216 and the cold side bleed assembly 108 of the CCA system 100. The flow control valve can be operably coupled to a controller, such as the example controller 130 described above with reference to Figure 2 .

[0136] ​In this manner, the number of individual openings on the flow path surface 190 of the split of the gas turbine engine 10 can be reduced, and the amount of plumbing of the gas turbine engine 10 overall can be reduced.

[0137] It will be appreciated, however, that in other example embodiments, the cold side bleed assembly 108 can define an inlet 114 positioned in fluid communication with the airflow over another extension. For example, in other embodiments, the extension can additionally or alternatively be configured as a pylon, a lower split, an outlet guide vane, an embedded structure of the gas turbine engine, etc.

[0138] The present disclosure relates generally to gas turbine engines having a cooling air system for cooling of one or more hot components of the gas turbine engine. In particular, as gas turbine engines are driven to higher efficiency and performance metrics, certain temperatures in the gas turbine engine can be increasing. The materials forming the components exposed to these temperatures can strain from these increases. Accordingly, cooling of these components is of great importance to modern gas turbine engines.

[0139] Certain attempts to cool these components have used pressurized airflow from the compressor section and directed the pressurized airflow to the hot components. However, the inventors of the present disclosure have found that these pressurized airflows are either too hot to prove useful or do not have sufficient pressure to flow through, for example, a first stage high pressure turbine rotor blade and the ambient pressure is too high.

[0140] The present disclosure addresses these issues via cooling of a high pressure airflow from, for example, a compressor discharge cavity of a gas turbine engine with a low pressure airflow from a cold location of the gas turbine engine through a cooled cooling air heat exchanger. After accepting heat from the high pressure airflow, the low pressure airflow can be exhausted through an undercowl region of the gas turbine engine in a manner that avoids damaging components of the turbomachinery, avoids antagonizing other systems of the turbomachinery, etc.

[0141] In particular, certain aspects of the present disclosure relate to a gas turbine engine including a turbomachinery having a compressor section, a combustion section defining a compressor discharge cavity, and a turbine section. The turbomachinery defines a working gas flow path and further includes a CCA system. The CCA system includes a hot side bleed assembly, a CCA heat exchanger in thermal communication with the hot side bleed assembly; and a cold side bleed assembly. The cold side bleed assembly defines an inlet and an outlet, the inlet in fluid communication with a cold location of the gas turbine engine. The outlet of the cold side bleed assembly is in fluid communication with the working gas flow path at a location downstream of the turbine section, in fluid communication with a core cowl vent of the turbomachinery, in fluid communication with a tunnel over the turbomachinery, or a combination thereof.

[0142] Reference is now made to the following drawings Figure 12, a gas turbine engine 10 according to another example embodiment of the present disclosure is provided. Figure 12 The gas turbine engine 10 can be constructed in a similar manner as one or more of the example gas turbine engines 10 described above with reference to Figures 1 to 11

[0143] For example, the example gas turbine engine 10 generally includes a turbomachine 16 having a compressor section, a combustion section 26, and a turbine section that collectively at least partially define a working gas flow path 37. The turbomachine 16 also includes a CCA system 100. The CCA system 100 includes a hot-side bleed assembly 110, a cold-side bleed assembly 108, and a CCA heat exchanger 112 in thermal communication with the hot-side bleed assembly 110 and the cold-side bleed assembly 108. The cold-side bleed assembly 108 defines an inlet 114 and an outlet 118, the inlet 114 being in fluid communication with a cold location of the gas turbine engine 10. The cold location can be one or more of the locations discussed above with reference to the embodiments of Figures 2 to 11 Figure 26

[0144] Further, for the illustrated embodiment, the outlet 118 of the cold-side bleed assembly 108 is in fluid communication with the working gas flow path 37 at a location downstream of the turbine section, in fluid communication with a core cowl vent 226 of the turbomachine 16, in fluid communication with a passage above the turbomachine 16, or a combination thereof.

[0145] In particular, the depicted example turbomachine 16 includes an outer casing 18 surrounding the compressor section, the combustion section 26, and the turbine section. The outer casing 18 includes a core cowl vent 226 at an aft end. In particular, the core cowl vent 226 is located aft of a downstream end 227 of the outer nacelle 50 and radially outward of the working gas flow path 37 through the ejection exhaust nozzle section 32. The outlet 118 of the cold-side bleed assembly 108 is in fluid communication with the core cowl vent 226 of the turbomachine 16.

[0146] ​​​In particular, the cold side bleed assembly 108 includes one or more conduits 116 that extend from the CCA heat exchanger 112 toward the core shroud vent 226 and to a location aft of the LP turbine 30 of the turbine section. In this manner, the cooling airflow 102 through the cold side bleed assembly 108 (which can be heated by virtue of its thermal communication with the cooling airflow 124 through the hot side bleed assembly 110) can not contact the outer casing surrounding the LP turbine 30 or the outer casing surrounding the HP turbine 28 of the turbine section, such that, for example, the clearance within the LP turbine 30 is not affected (and the operation of the active clearance control system is not affected; as discussed below). In particular, for the depicted embodiment, the one or more conduits 116 of the cold side bleed assembly 108 extend to the core shroud vent 226 such that the cooling airflow 102 through the cold side bleed assembly 108 is exhausted directly through the core shroud vent 226.

[0147] It will be appreciated, however, that in other example embodiments, any other suitable configuration can be provided. For example, referring now to Figure 13 , a gas turbine engine 10 and CCA system 100 are provided in accordance with another example embodiment of the present disclosure. Figure 13 The example CCA system 100 of Figure 12 may be configured in a similar manner as the example CCA system 100 of Figure 13 . For example, The CCA system 100 of

[0148] In particular, it will be appreciated that the turbine engine 16 further includes a HP turbine active clearance control (“ACC”) system 228 and a LP turbine ACC system 230. The HP turbine ACC system 228 is located about the HP turbine 28 of the turbine section, while the LP turbine ACC system 230 is located about the LP turbine 30 of the turbine section. The depicted example turbine engine 16 further defines an under-shroud flow path 232 for delivering the cooling airflow 102 through the outlet 118 of the cold side bleed assembly 108 to the core shroud vent 226. The under-shroud flow path 232 extends at least from the outlet 118 to the core shroud vent 226 and over at least a portion of the LP turbine 30.

[0149] To avoid the temperature of the cooling airflow 102 exiting the outlet 118 of the cold side bleed assembly 108 from affecting the operation of, for example, the LP turbine ACC system 230, the turbine 16 further includes a thermal shield 234 positioned between the shroud downstream flowpath 232 and the LP turbine 30 of the turbine section. More specifically, the thermal shield 234 extends along the axial direction A of the gas turbine engine 10 between at least the outlet 118 of the cold side bleed assembly 108 and a downstream end of the LP turbine 30.

[0150] Briefly stated, it will be appreciated that the thermal shield 234 can additionally or alternatively prevent the temperature of the cooling airflow 102 exiting the outlet 118 of the cold side bleed assembly 108 from affecting the operation of other systems, including but not limited to pumps, valves, controllers, fuel system components, environmental control systems, electrical systems such as generators or power converters, and lubrication system components.

[0151] Further, the cold side bleed assembly 108 further includes a diffuser 236 at the outlet 118. The diffuser 236 is configured to temper the cooling airflow 102 passing through the outlet 118 of the cold side bleed assembly 108 to avoid undesirable impingement heating on one or more components within the turbine 16. The diffuser 236 can be a simple area change diffuser that defines an area ratio between a downstream end and an upstream end that is greater than or equal to 1.5: 1 and less than or equal to 10: 1. Alternatively, more complex diffusers can be used.

[0152] Further, in other example embodiments, other suitable configurations can be utilized. For example, referring now to Figure 14 A gas turbine engine 10 and a CCA system 100 according to another example embodiment of the present disclosure are provided. Figure 14 The example CCA system 100 of Figure 12 may be configured in a similar manner as the example CCA system 100 of Figure 14 The CCA system 100 of

[0153] In particular, the depicted exemplary turbine 16 of gas turbine engine 10 further includes a transient bleed assembly 238 including a transient bleed conduit 240 extending between an inlet 242 and an outlet 244. Inlet 242 of the transient bleed conduit is in fluid communication with working gas flow path 37 at HP compressor 24 of the compressor section, with compressor discharge plenum 138 of combustion section 26 of turbine 16, or both. Outlet 244 of the transient bleed conduit is in fluid communication with working gas flow path 37 downstream of LP turbine 30.

[0154] For the depicted embodiment, outlet 118 of cold side bleed assembly 108 is in fluid communication with transient bleed conduit 240 at a location between inlet 242 of transient bleed conduit 240 and outlet 244 of transient bleed conduit 240. In this manner, the cooling gas flow 102 through cold side bleed assembly 108 can be bled into transient bleed conduit 240 and merged with the gas flow through transient bleed conduit 240 before being provided into working gas flow path 37 at a location downstream of LP turbine 30 through outlet 244.

[0155] Referring now to Figure 15 , a perspective view of transient bleed assembly 238 of Figure 14 and CCA system 100 of Figure 14 is provided. As will be appreciated, transient bleed conduit 240 includes a primary conduit 246 extending from inlet 114 and a plurality of branches 248. Outlet 244 of transient bleed conduit 240 is one of a plurality of outlets 244, and the plurality of branches 248 extend between primary conduit 246 and one of the respective plurality of outlets 244. Further, in the depicted embodiment, CCA heat exchanger 112 is a first CCA heat exchanger 112A, and CCA system 100 further includes a second CCA heat exchanger 112B. Further, cold side bleed assembly 108 includes a first portion 250 and a second portion 252, with first portion 250 being in thermal communication with first CCA heat exchanger 112A and second portion 252 being in thermal communication with second CCA heat exchanger 112B. First portion 250 of cold side bleed assembly 108 and second portion 252 of cold side bleed assembly 108 each define one or more outlets 118 that open into a respective one of the plurality of branches 248 of transient bleed conduit 240.

[0156] Notably, exemplary transient bleed assembly 238 further includes a transient bleed valve 254 that is operable with transient bleed conduit 240, and more particularly, with primary conduit 246 of transient bleed conduit 240. Transient bleed valve 254 can regulate the amount of bleed gas flow through transient bleed conduit 240. Transient bleed valve 254 can be in communication with a controller, such as controller 220, to regulate the amount of bleed gas flow through transient bleed conduit 240.Figure 2 The exemplary controller 130 of the example CCA system 100 of

[0157] Although not depicted, the cold side bleed assembly 108 can include one or more flow control valves that regulate the flow of the cooling gas stream 102 through the cold side bleed assembly 108, upstream of the CCA heat exchanger 112 or downstream of the CCA heat exchanger 112.

[0158] Figure 14 And Figure 15 The configuration of the gas turbine engine 10 and CCA system 100 of

[0159] Briefly, with reference now to Figure 16 And Figure 17 a gas turbine engine 10 and CCA system 100 according to another example embodiment of the present disclosure are provided. Figure 16 And Figure 17 The gas turbine engine 10 and CCA system 100 of Figure 14 And Figure 15 may be configured in a similar manner as the example gas turbine engine 10 and CCA system of Figure 16 And Figure 17 The CCA system 100 of and the example turbine 16 of the depicted gas turbine engine 10 further includes a transient bleed assembly 238 that includes a transient bleed conduit 240 extending between an inlet 242 and an outlet 244. The outlet 244 of the transient bleed conduit is in fluid communication with the working gas flow path 37 downstream of the LP turbine 30. However, for the depicted embodiment, the inlet 242 of the transient bleed conduit is in fluid communication with the hot side bleed assembly 110, and more particularly, the hot side bleed assembly 110 at a location upstream of the CCA heat exchanger 112. In this manner, the transient bleed assembly 238 does not need to bleed off separately from the working gas flow path 37, the compressor discharge plenum 138, or both. Rather, the hot side bleed assembly 110 can extract enough gas flow to achieve the goals of both the CCA system 100 and the transient bleed assembly 238. As will be appreciated, the CCA system 100 and the transient bleed assembly 238 generally operate under different operating conditions of the gas turbine engine 10, such that the inlet of the hot side bleed assembly 110 can not need to be increased to be able to accept an amount of air equal to (and in some cases can not be able to accept) the maximum bleed requirement of the hot side bleed assembly 110 and the maximum bleed requirement of the transient bleed assembly 238.

[0160] Briefly referring to Figure 17 In the depicted embodiment, the example transient bleed assembly 238 also includes a transient bleed valve 254 operable with the transient bleed conduit 240. More specifically, the transient bleed conduit 240 includes a first transient bleed conduit 240A in fluid communication with the first hot-side bleed assembly 110A, and the transient bleed conduit 240 also includes a second transient bleed conduit 240B in fluid communication with the second hot-side bleed assembly 110B (the first and second transient bleed conduits 240A, 240B each at a location upstream of the respective CCA heat exchangers 112A, 112B). For this example embodiment, the transient bleed valve 254 is a first transient bleed valve 254A operable with the first transient bleed conduit 240A, and the example transient bleed assembly 238 also includes a second transient bleed valve 254B operable with the second transient bleed conduit 240B. The first and second transient bleed valves 254A, 254B can be in operable communication with a controller (see, e.g., the controller 130 in Figure 2 to control the amount of air flow through the transient bleed assembly 238.

[0161] Notably, as schematically depicted, the CCA system 100 can also include a check valve 255 in the cold-side bleed assembly 108 downstream of the CCA heat exchanger 112 to prevent air from the transient bleed assembly 238 from traveling through the cold-side bleed assembly 108 from the transient bleed conduit 240 toward the CCA heat exchanger 12.

[0162] Although Figure 16 two transient bleed conduits 240A, 240B are shown in

[0163] It will be appreciated, however, that in other example embodiments, the CCA system 100 and gas turbine engine 10 can have other suitable configurations. For example, now referring to Figure 18 , a gas turbine engine 10 and CCA system 100 according to another example embodiment of the present disclosure are provided. Figure 18 The example CCA system 100 of Figure 12 may be configured in a similar manner as the example CCA system 100 of Figure 18 For example, The CCA system 100 of

[0164] In particular, for the depicted embodiment, the outlet 118 of the cold-side bleed assembly 108 is in fluid communication with a bypass passage 56 of the gas turbine engine 10, the bypass passage 56 being defined between an outer nacelle 50 of the gas turbine engine 10 and a turbine 16 of the gas turbine engine 10. In particular, the depicted gas turbine engine 10 includes a bifurcation connected to the turbine 16 and extending between the turbine 16 and the outer nacelle 50. The outlet 118 of the cold-side bleed assembly 108 is in fluid communication with an opening in a flowpath surface 190 of the bifurcation. More particularly, the bifurcation extends between a leading edge 192 and a trailing edge 194, and the opening in the flowpath surface 190 is positioned at the trailing edge 194 of the bifurcation. Still more particularly, in accordance with the depicted embodiment, the CCA system 100 includes a diffusion device 256 integrated into the bifurcation at the trailing edge 194 of the bifurcation. The diffusion device 256 defines a plurality of openings (not labeled), and the outlet 118 of the cold-side bleed assembly 108 is in fluid communication with the plurality of openings of the diffusion device 256 to allow the cooling airflow 102 through the cold-side bleed assembly 108 to be effectively merged with the airflow through the bypass passage 56.

[0165] It will be appreciated that such a configuration can limit the amount of plumbing required for the CCA system 100, while also allowing the heated cooling airflow 102 through the cold-side bleed assembly 108 to be reintroduced into the flowstream of the gas turbine engine 10, which can generate useful work for the gas turbine engine.

[0166] As noted above, aspects of the present disclosure relate to a gas turbine engine having a cooling air system for cooling one or more hot components of the gas turbine engine. In particular, the CCA system includes a cold-side bleed assembly defining an inlet in fluid communication with a cold location of the gas turbine engine, a hot-side bleed assembly defining an inlet in fluid communication with a working gas flowpath at a compressor discharge cavity (or, for example, through a compressor section), and a CCA heat exchanger in thermal communication with the cold-side bleed assembly and the hot-side bleed assembly to transfer heat from a cooling airflow through the hot-side bleed assembly to a cooling airflow through the cold-side bleed assembly. The hot-side bleed assembly is also in thermal communication with a hot component of the turbine to cool the hot component of the turbine.

[0167] As will be further appreciated, certain embodiments of the CCA system include a first flow control valve that controls the flow rate of the cooling gas stream through the cold side bleed assembly and a second flow control valve that controls the flow rate of the cooled gas stream through the hot side bleed assembly. Aspects of the present disclosure discussed below provide a system for regulating the cooling gas stream and the cooled gas stream in tandem, allowing for more efficient use of the cooling gas stream and the cooled gas stream (e.g., ensuring that the amount of cooling gas stream is extracted only as needed to cool the amount of cooled gas stream), while also reducing the complexity and cost of the system by combining features of the first and second flow control valves.

[0168] More specifically, referring now to Figure 19 , a schematic cross-sectional view of a gas turbine engine 10 having a CCA system 100 according to another example embodiment of the present disclosure is provided. Figure 19 The example gas turbine engine 10 and CCA system 100 can be constructed in a similar manner as one or more of the example gas turbine engines 10 and CCA systems 100 described above with reference to Figures 1 to 18

[0169] For example, Figure 19 The example gas turbine engine 10 generally includes a fan section 14 and a turbine machine 16, with the turbine machine 16 including a CCA system 100. The CCA system 100 includes a cold side bleed assembly 108 defining an inlet 114 and an outlet 118, with the inlet 114 being in fluid communication with a cold location of the gas turbine engine 10. A hot side bleed assembly 110 defines an inlet 122 that is in fluid communication with the working gas flow path 37 through the turbine machine 16, and the hot side bleed assembly 110 is also in thermal communication with hot components of the turbine machine 16 to cool the hot components of the turbine machine 16. During operation, a cooling gas stream 102 is provided through the cold side bleed assembly 108, and a cooled gas stream 124 is provided through the hot side bleed assembly 110.

[0170] However, for Figure 19 The gas turbine engine 10, and more specifically the turbine machine 16 of the gas turbine engine 10, includes a valve assembly 260 that is operable with the cold side bleed assembly 108 and the hot side bleed assembly 110. The valve assembly 260 is structured to regulate the gas streams through the cold side bleed assembly 108 and the hot side bleed assembly 110 in tandem. In particular, the valve assembly 260 is structured to regulate the cooling gas stream 102 through the cold side bleed assembly 108 and the cooled gas stream 124 through the hot side bleed assembly 110 in tandem.

[0171] ​As used herein, the term“in coordination” with respect to the regulation of two gas flows means an arrangement in which the regulation of one gas flow necessarily regulates the other gas flow. The relative regulation can be the same (e.g., a 10% increase in one gas flow means a 10% increase in the other gas flow), or the relative regulation can be different (e.g., a 10% increase in one gas flow means a 20% increase in the other gas flow).

[0172] In particular, referring now to Figure 20 , there is provided Figure 19 a perspective view of a portion of the CCA system 100 of FIG. 1, showing an example valve assembly 260. As will be appreciated, the valve assembly 260 includes a cold-side valve section 262 operable with the cold-side off-gas assembly 108 and a hot-side valve section 264 operable with the hot-side off-gas assembly 110. Additionally, the valve assembly 260 includes an actuator 266 coupled to both the cold-side valve section 262 and the hot-side valve section 264. The cold-side valve section 262 and the hot-side valve section 264 are each operable to change the cooling gas flow 102 and the cooled gas flow 124, respectively, in response to movement by the actuator 266.

[0173] More particularly, referring now to Figure 21 , there is provided Figure 20 a close-up view of the valve assembly 260 of FIG. 2, the valve assembly 260 further including a shaft 268 movable by the actuator 266. The actuator 266 is coupled to the cold-side valve section 262 and the hot-side valve section 264 by the shaft 268. In the depicted embodiment, the actuator 266 is configured to rotate the shaft 268 in a circumferential direction CI about an axis of the shaft 268. In this manner, it will be appreciated that the cold-side valve section 262 and the hot-side valve section 264 are each configured to regulate the cooling gas flow 102 and the cooled gas flow 124, respectively, in response to circumferential movement of the shaft 268 by the actuator 266.

[0174] Still more particularly, referring now to Figure 22 , there is provided Figure 21 a schematic cross-sectional view of the valve assembly 260 of FIG. 3. It will be appreciated that, for the depicted embodiment, the cold-side valve section 262 includes a first ball valve portion 270, while the hot-side valve section 264 includes a second ball valve portion 272. The first ball valve portion 270 is positioned within the cold-side off-gas assembly 108, and more particularly, within a duct of the cold-side off-gas assembly 108, while the second ball valve portion 272 is positioned within the hot-side off-gas assembly 110, and more particularly, within a duct of the hot-side off-gas assembly 110. The first ball valve portion 270 and the second ball valve portion 272 are each coupled to the shaft 268 such that rotation of the shaft 268 correspondingly rotates the first ball valve portion 270 and the second ball valve portion 272 in unison, thereby regulating gas flow therethrough in coordination.

[0175] In this manner, it will be appreciated that the valve assembly 260 is configured to cooperatively regulate the cooling airflow 102 through the cold side bleed assembly 108 and the cooled airflow 124 through the hot side bleed assembly 110 such that the amount of cooling airflow 102 provided through the CCA heat exchanger 112 is matched to the amount of cooled airflow 124 provided through the CCA heat exchanger 112 to the required amount of cooling.

[0176] Notably, in the embodiment depicted above with reference to Figure 20 and Figure 21 , the cold side valve segment 262 is positioned upstream of the CCA heat exchanger 112 (and downstream of the inlet 114 of the cold side bleed assembly 108), while the hot side valve segment 264 is positioned downstream of the CCA heat exchanger 112. In this manner, the hot side valve segment 264 can be exposed to lower temperatures during operation.

[0177] It will also be appreciated that, as depicted by the dashed lines in Figure 20 , the actuator 266 of the valve assembly 260 can be operably coupled to the controller 130 of the gas turbine engine 10, such that the actuator 266 can be controlled in response to received data indicative of operating conditions or other suitable parameters of the gas turbine engine 10. The actuator 266 can be wirelessly coupled, as indicated by the dashed lines in Figure 20 , or alternatively can be operably coupled by a wired connection.

[0178] In other example embodiments of the present disclosure, other suitable configurations can be provided. For example, in other embodiments, any other suitable flow control valve can be utilized as the cold side valve segment 262, the hot side valve segment 264, or both. For example, in other example embodiments, the valve assembly 260 can utilize one or more butterfly valves, gate valves, ball valves, or the like.

[0179] Further, although the actuator 266 of the valve assembly 260 described above is a rotary actuator 266, in other example embodiments, the actuator 266 can be a linear actuator 266 operable to move the shaft 268 along the axis of the shaft 268. For example, referring now to Figure 23 and Figure 24 , two additional perspective views of valve assemblies according to example embodiments of the present disclosure are provided.

[0180] Figure 23 and Figure 24 , each generally include a cold side valve segment 262 operable with the cold side bleed assembly 108 and a hot side valve segment 264 operable with the hot side bleed assembly 110. Additionally, the valve assembly 260 includes an actuator 266 coupled to the respective cold side valve segment 262 and hot side valve segment 264 by a respective shaft 268 of the valve assembly 260.

[0181] With particular reference to Figure 23 , the valve assembly 260 further includes a first pull rod 274 and a second pull rod 276. The actuator 266 is coupled to a shaft 268, and the first pull rod 274 and the second pull rod 276 are each rotatably coupled to an opposite end of the shaft 268 at a first connection point 278. The first pull rod 274 is coupled to the cold side valve segment 262, and the second pull rod 276 is coupled to the hot side valve segment 264. Moving the shaft 268 linearly along an axis of the shaft 268 by the actuator 266 moves the first pull rod 274 and the second pull rod 276, which in turn rotates the first rod 280 of the cold side valve segment 262 and the second rod 282 of the hot side valve segment 264 in unison to actuate the cold side valve segment 262 and the hot side valve segment 264.

[0182] Similarly, now referring to Figure 24 , the shaft 268 of the valve assembly 260 is coupled to a connecting rod 284 of the valve assembly 260, which in turn is pivotably coupled to the first rod 280 of the cold side valve segment 262 and the second rod 282 of the hot side valve segment 264. Moving the connecting rod 284 by the shaft 268 and the actuator 266 of the valve assembly 260 in Figure 24 correspondingly actuates the cold side valve segment 262 and the hot side valve segment 264 in unison.

[0183] Now referring to Figure 25 , a method 300 of operating a valve assembly of a gas turbine engine in accordance with example aspects of the present disclosure is provided. The method 300 can be used with one or more of the example embodiments discussed above. Figures 19 to 22

[0184] The method 300 generally includes receiving data indicative of an operating condition of the gas turbine engine at (310), and actuating a valve assembly of a cooled cooling air system in response to receiving the data indicative of the operating condition of the gas turbine engine to regulate a first airflow through a cold side bleed assembly of the cooled cooling air system and a second airflow through a hot side bleed assembly of the cooled cooling air system at (312).

[0185] In certain example aspects, the data indicative of the operating condition can be data indicative of a power level of the gas turbine engine. For example, the data can be indicative of a rotational speed of the gas turbine engine, a fuel flow to the gas turbine engine, a power lever position, etc.

[0186] ​In certain example aspects, the valve assembly can include a cold-side valve section operable with the cold-side bleed assembly, a hot-side valve section operable with the hot-side bleed assembly, and an actuator coupled to both the hot-side valve section and the cold-side valve section. In such example aspects, actuating the valve assembly at (312) includes moving a shaft of the valve assembly with the actuator at (314), the shaft being coupled to both the cold-side valve section and the hot-side valve section.

[0187] Further, for the depicted example aspects, the method 300 further includes receiving data indicative of a degradation parameter of the gas turbine engine at (320), and actuating the valve assembly in response to receiving the data indicative of the degradation parameter at (322). For example, the degradation parameter can be any data indicative of a level of degradation of the gas turbine engine. For example, the data indicative of the level of degradation can be data indicative of an operational life of the gas turbine engine, a fuel flow rate of the gas turbine engine required to achieve a particular power level, etc.

[0188] For such example aspects, the method 300 can allow the cooling airflow and the amount of the cooled airflow to vary based on a degree of degradation of the engine to compensate for such degradation, while further allowing for higher efficiency for engines with lower degrees of degradation.

[0189] As discussed above, aspects of the present disclosure relate to a gas turbine engine having a cooled cooling air system for cooling one or more hot components of the gas turbine engine. In particular, the CCA system includes a cold-side bleed assembly defining an inlet in fluid communication with a cold location of the gas turbine engine, a hot-side bleed assembly defining an inlet in fluid communication with a working gas flow path at a compressor discharge cavity (or, for example, through a compressor section), and a CCA heat exchanger in thermal communication with both the cold-side bleed assembly and the hot-side bleed assembly to transfer heat from a cooled airflow through the hot-side bleed assembly to a cooling airflow through the cold-side bleed assembly. The hot-side bleed assembly is also in thermal communication with a hot component of a turbine of the gas turbine engine to cool the hot component of the turbine.

[0190] As will be further appreciated, it is useful to include an active clearance control (ACC) system within a turbine section of a gas turbine engine to maintain a desired clearance between turbine rotor blades and a casing of a turbine of the turbine section. The ACC system can operate by providing an airflow onto the casing of the turbine to control thermal growth of the casing relative to the turbine rotor blades.

[0191] In view of the relatively low temperature of the airflow, certain ACC systems use a low pressure airflow, such as a bypass airflow. However, depending on, for example, a complexity of the ACC system, a higher pressure airflow can be required to drive the airflow through a manifold of the ACC system. Accordingly, a blower (e.g., pump) can be provided to push the airflow through the ACC system.

[0192] Surprisingly, the inventors of the present disclosure have found that, by pressurizing the low pressure air of the ACC system, the same air flow (or more precisely, a portion of the same air flow) can also be used as the air flow source for the CCA system. In particular, the inventors have found that, by including such a blower with the CCA system, a sufficient amount of pressurized air can be generated for the CCA system and the ACC system from the low pressure air source.

[0193] Such an arrangement can allow for improved overall engine efficiency by reducing or eliminating extraction of air flow from the working gas flow path of the engine downstream of one or more compression stages.

[0194] In particular, reference is now made to Figure 26 , a schematic cross-sectional view of a gas turbine engine 10 having a CCA system 100 according to another example embodiment of the present disclosure is provided. Figures 1 to 25 The example gas turbine engine 10 and CCA system 100 can be constructed in a similar manner as one or more of the example gas turbine engines 10 and CCA systems 100 described above with reference to Figure 26

[0195] For example, Figure 26 The example gas turbine engine 10 generally includes a fan section 14 and a turbine machine 16, where the turbine machine 16 includes a compressor section, a combustion section 26, and a turbine section that collectively at least partially define a working gas flow path 37. The turbine section includes a first turbine and a second turbine, and more specifically, an LP turbine 30 and an HP turbine 28. The turbine machine 16 also includes a CCA system 100. The CCA system 100 includes a cold side bleed assembly 108 defining an inlet 114 and an outlet 118, the inlet 114 being in fluid communication with a cold location of the gas turbine engine 10. A hot side bleed assembly 110 defines an inlet 122, the inlet 122 being in fluid communication with the working gas flow path 37 through the turbine machine 16, and the hot side bleed assembly 110 is also in thermal communication with hot components of the turbine machine 16 to cool the hot components of the turbine machine 16. During operation, a cooling air flow 102 is provided through the cold side bleed assembly 108, and a cooled air flow 124 is provided through the hot side bleed assembly 110. Further, the CCA system 100 includes a CCA heat exchanger 112, the CCA heat exchanger 112 being in thermal communication with the cold side bleed assembly 108 and the hot side bleed assembly 110, and more specifically, with the cooling air flow 102 through the cold side bleed assembly 108 and the cooled air flow 124 through the hot side bleed assembly 110.

[0196] Further, for Figure 26 ​embodiments, a gas turbine engine 10, and more specifically, a turbine 16 of the gas turbine engine 10, includes an active clearance control (ACC) system 400. The ACC system 400 includes an inlet 402, a heat transfer assembly arranged about a turbine section of the turbine, and an ACC piping assembly 404 extending from the inlet 402 to the heat transfer assembly.

[0197] In particular, for the depicted embodiment, the turbine is a first turbine of the turbine section, or more specifically, a HP turbine 28 of the turbine section, and the turbine section further includes a second turbine, or more specifically, an LP turbine 30. For this example embodiment, the heat transfer assembly is a first heat transfer assembly 406, and the ACC system 400 further includes a second heat transfer assembly 408 arranged about the LP turbine 30. The ACC piping assembly 404 extends to both the first heat transfer assembly 406 and the second heat transfer assembly 408. In particular, for the illustrated embodiment, the ACC piping assembly 404 includes a first portion 410 extending to the first heat transfer assembly 406, a second portion 412 extending to the second heat transfer assembly 408, and an upstream portion 414 extending from the inlet 402 of the ACC system 400 to both the first portion 410 and the second portion 412.

[0198] Still referring to Figure 26 embodiments, it will be appreciated that the inlet 402 of the ACC piping assembly 404 is in fluid communication with the bypass passage 56 defined above the turbine 16. In particular, Figure 26 The example gas turbine engine 10 includes an outer nacelle 50, and the bypass passage 56 is defined between the outer nacelle 50 and the outer casing 18 of the turbine 16. The inlet 402 of the ACC system 400 is defined in the outer casing 18 such that the ACC system 400, and more specifically, the ACC piping assembly 404, receives a flow of gas from the bypass passage 56 through the inlet 402 to the ACC system 400 during operation.

[0199] Notably, in other example embodiments, the inlet 402 of the ACC system 400 can be at any other suitable location to receive a flow of gas from, for example, a cold location of the gas turbine engine 10. For example, in other example embodiments, the inlet 402 can be configured to receive a flow of gas through the working gas flow path 37 of the gas turbine engine 10 at a location through the LP compressor 22 of the compressor section, between the LP compressor 22 and the HP compressor 24 of the compressor section, or both. Additionally or alternatively, the inlet 402 can be positioned on the flow path surface 190 of the diverging portion of the gas turbine engine 10.

[0200] Further, for Figure 2In accordance with the depicted example embodiment, the inlet 114 of the cold side bleed assembly 108 of the CCA system 100 is in fluid communication with the ACC duct assembly 404. In this manner, the CCA system 100 can be integrated with the ACC system 400, thereby reducing the ducting and complexity of the gas turbine engine 10.

[0201] For the depicted embodiment, the CCA system 100 includes a flow control valve 415. More specifically, the CCA system 100 includes the flow control valve 415 at the inlet 114 of the cold side bleed assembly 108, the flow control valve 415 being positioned within the ACC duct assembly 404 such that the inlet 114 of the cold side bleed assembly 108 is in fluid communication with the upstream portion 414 of the ACC duct assembly 404 through the flow control valve 415. The flow control valve 415 is more specifically a variable control valve (e.g., operable to vary the flow rate to the various outlets) of the CCA system 100.

[0202] Briefly, it will be appreciated that, for the depicted embodiment, the ACC system 400 includes one or more flow control valves operable with the first portion 410 of the ACC duct assembly 404, the second portion 412 of the ACC duct assembly 404, or both. More specifically, the ACC system 400 includes a first flow control valve 416 operable with the first portion 410 of the ACC duct assembly 404, and a second flow control valve 418 operable with the second portion 412 of the ACC duct assembly 404. The first flow control valve 416, the second flow control valve 418, and the flow control valve 415 can each be operably connected to a controller (such as the example controller 130) for controlling operation of the respective valve and gas flow through the respective duct. Figure 26

[0203] Further, still referring to Figure 27 , the example CCA system 100 additionally includes a blower 420 in fluid communication with the cold side bleed assembly 108 for pushing the cooling gas stream 102 through the cold side bleed assembly 108. As will be appreciated, the cooling gas stream 102 received from the ACC system 400 (which in turn is received from the bypass passage 56) can be a relatively low pressure gas stream. The blower 420 can increase the pressure of the cooling gas stream 102 to allow for desired operation of the CCA system 100.

[0204] However, it will be appreciated that other suitable configurations can be provided in other example embodiments. For example, referring now to Figure 28 and Figure 27 schematic diagrams of the gas turbine engine 10 are provided in accordance with additional example embodiments of the present disclosure. Figure 28 and Figure 26 The example gas turbine engine 10 of Figure 27 ​in a similar manner to the exemplary gas turbine engines 10.

[0205] For example, the exemplary gas turbine engines 10 each generally include a turbine 16 having an ACC system 400 and a CCA system 100. The ACC system 400 of each gas turbine engine 10 generally includes an inlet 402, a first heat transfer assembly 406 disposed about a turbine section of the HP turbine 28 of the respective turbine 16, a second heat transfer assembly 408 disposed about a turbine section of the LP turbine 30 of the respective turbine 16, and an ACC duct assembly 404 extending from the inlet 402 to the first and second heat transfer assemblies. The ACC duct assembly 404 of each ACC system 400 includes an upstream portion 414, a first portion 410, and a second portion 412.

[0206] Further, the CCA system 100 of each turbine 16 generally includes a cold side bleed assembly 108, a hot side bleed assembly 110, and a CCA heat exchanger 112 in thermal communication with the hot side bleed assembly 110 and the cold side bleed assembly 108. The cold side bleed assembly 108 defines an inlet 114 in fluid communication with the ACC duct assembly 404.

[0207] With particular reference to Figure 26 For the depicted exemplary turbine 16, the CCA system 100 does not include a separate blower in fluid communication with the cold side bleed assembly 108 (see Figure 28 the blower 420 in FIG. 4). Rather, for the depicted embodiment, the ACC system 400 includes a blower 420 in fluid communication with the ACC duct assembly 404 for increasing the pressure through the ACC duct assembly 404. For this configuration, the inlet 114 of the cold side bleed assembly 108 is in fluid communication with the ACC duct assembly 404 at a location downstream of the blower 420. More specifically, for the depicted embodiment, the blower 420 of the ACC system 400 is in fluid communication with the upstream portion 414 of the ACC duct assembly 404, and the inlet 114 is also in fluid communication with the upstream portion 414 of the ACC duct assembly 404. In this manner, the blower 420 can increase the pressure of the airflow provided to the first heat transfer assembly 406 through the first portion 410 of the ACC duct assembly 404, to the second heat transfer assembly 408 through the second portion 412 of the ACC duct assembly 404, and to the CCA system 100 / through the CCA system 100.

[0208] With particular reference to Figure 28, likewise the CCA system 100 does not include a blower, and instead the ACC system 400 includes a blower 420. For the depicted embodiment, however, the blower 420 of the ACC system 400 is located in the second portion 412 of the ACC duct assembly 404. For this configuration, the inlet 114 of the cold side bleed assembly 108 of the CCA system 100 is in fluid communication with the second portion 412 of the ACC duct assembly 404 at a location downstream of the blower 420. In this manner, the blower 420 can increase the pressure of the airflow provided through the second portion 412 of the ACC system 400 to the second heat transfer assembly 408 (e.g., push the airflow through the second portion 412), as well as the pressure of the airflow provided by the CCA system 100 as part of the cooling airflow 102 (received from the second portion 412 of the ACC duct assembly 404; not shown in FIG. 4A). Figure 29

[0209] Briefly, now referring to Figure 28 , there is provided Figure 30 a close-up view of the first heat transfer assembly 406 and the second heat transfer assembly of the ACC system 400. With specific reference to the first heat transfer assembly 406, the first heat transfer assembly 406 includes a first manifold 422 and a first plurality of circumferential airflow extensions 424. Notably, the HP turbine 28 includes a casing 426 that encloses the multi-stage HP turbine rotor blades of the HP turbine 28, and more specifically, the two-stage HP turbine rotor blades of the HP turbine 28. The first plurality of circumferential airflow extensions 424 includes two circumferential airflow extensions 424 that extend circumferentially around the casing 426 of the HP turbine 28. The two circumferential airflow extensions 424 are configured to blow airflow onto the casing 426 to control thermal growth of the casing 426 and maintain a gap between the multi-stage HP turbine rotor blades and the casing 426.

[0210] With specific reference to the second heat transfer assembly 408, the second heat transfer assembly 408 includes a second manifold 428 and a second plurality of circumferential airflow extensions 430. The LP turbine 30 also includes a casing 432 that encloses the multi-stage LP turbine rotor blades of the LP turbine 30, and more specifically, the four-stage LP turbine rotor blades. The second plurality of circumferential airflow extensions 430 includes four circumferential airflow extensions 430 that extend circumferentially around the casing 432 of the LP turbine 30. Each of the second plurality of circumferential airflow extensions 430 is configured to blow airflow onto the casing 432 to control thermal growth of the casing 432 and maintain a gap between the multi-stage LP turbine rotor blades and the casing 432.

[0211] ​Notably, given the increase in the number of stages within LP turbine 30 as compared to HP turbine 28, and the corresponding increase in the number of circumferential airflow extensions 424, 430 in second heat transfer assembly 408 as compared to first heat transfer assembly 406, the desired operation within second heat transfer assembly 408 can require an increase in airflow and / or an increase in the pressure of the airflow as compared to first heat transfer assembly 406. Moreover, the effectiveness of ACC system 400 is dependent on the use of high pressure air to impinge on enclosures 426, 432 to achieve maximum heat transfer rates. First heat transfer assembly 406 for HP turbine 28 can generally include larger diameter tubes (e.g., circumferential airflow extensions 424, first manifold 422), thus there is little loss in flow pressure when delivering cooling air to impingement baffle / extensions 424. In contrast, second heat transfer assembly 408 for LP turbine 30 can generally include smaller diameter tubes because it is desirable to distribute flow to more spatial locations and reduce engine core cowl size, which reduces the cooling air pressure at impingement baffle / extensions 430. Thus, increasing the pressure of second heat transfer assembly 408 of LP turbine 30 can improve its cooling efficiency.

[0212] In this manner, it will be appreciated that including a blower 420 in fluid communication with the second portion 412 of ACC duct assembly 404 (rather than the first portion 410 of ACC duct assembly 404) can facilitate the desired pressure through the second portion 412 of ACC duct assembly 404 to achieve the desired operation.

[0213] Referring now to Figure 30 , there is provided a gas turbine engine 10 according to yet another example embodiment of the present disclosure. Figures 26 to 29 The example gas turbine engine 10 of Figure 30 may be constructed in a similar manner to one or more of the example gas turbine engines 10 described above with reference to Figure 29The exemplary gas turbine engine 10 includes an ACC system 400 and a CCA system 100. The ACC system 400 generally includes an inlet 402, a first heat transfer assembly 406, a second heat transfer assembly 408, and an ACC duct assembly 404 extending from the inlet 402 to the first heat transfer assembly 406 and the second heat transfer assembly 408. In particular, the ACC duct assembly 404 includes an upstream portion 414, a first portion 410 extending from the upstream portion 414 to the first heat transfer assembly 406, and a second portion 412 extending from the upstream portion 414 to the second heat transfer assembly 408. The first portion 410, the second portion 412, and the upstream portion 414 of the ACC duct assembly 404 meet at a junction 434. For the illustrated embodiment, the CCA system 100 includes a hot side bleed assembly 110, a cold side bleed assembly 108, and a CCA heat exchanger 112 in thermal communication with the hot side bleed assembly 110 and the cold side bleed assembly 108. The cold side bleed assembly 108 defines an inlet 114 in fluid communication with the ACC duct assembly 404. In particular, for the illustrated embodiment, the inlet 114 of the cold side bleed assembly 108 is in fluid communication with the ACC duct assembly 404 at the junction 434. The cold side bleed assembly 108 also defines an outlet 118. For the illustrated embodiment, the outlet 118 is selectively in fluid communication with the ACC duct assembly 404 at a location upstream of the first heat transfer assembly 406, a location upstream of the second heat transfer assembly 408, or both. In particular, for the illustrated embodiment, the outlet 118 of the cold side bleed assembly 108 is selectively in fluid communication with the first portion 410 of the ACC duct assembly 404 at a location downstream of the junction 434 and upstream of the first heat transfer assembly 406, and is selectively in fluid communication with the second portion 412 of the ACC duct assembly 404 at a location downstream of the junction 434 and upstream of the second heat transfer assembly 408.

[0214] It will be appreciated that, for this configuration, the CCA system 100, and more particularly the cold side bleed assembly 108 of the CCA system 100, can provide a relatively hot airflow to the first heat transfer assembly 406, the second heat transfer assembly, or both. This arrangement can allow the ACC system 400 to slow the thermal growth around the outer casing 432 (see Figure 29 ) of the LP turbine 30, around the outer casing 426 (see Figure 31 ) of the HP turbine 28, or both, or alternatively can increase the thermal growth of the outer casings 432, 426 by providing hot air from the cold side bleed assembly 108 of the CCA system 100. This configuration can provide greater flexibility in the control of the gas turbine engine 10.

[0215] As discussed above, aspects of the present disclosure relate to a gas turbine engine having a cooling air system for cooling one or more hot components of the gas turbine engine. In particular, the CCA system includes a cold side bleed assembly defining an inlet in fluid communication with a cold location of the gas turbine engine, a hot side bleed assembly defining an inlet in fluid communication with a working gas flow path at a compressor discharge plenum (or, for example, through a compressor section), and a CCA heat exchanger in thermal communication with both the cold side bleed assembly and the hot side bleed assembly to transfer heat from a cooled gas stream through the hot side bleed assembly to a cooled gas stream through the cold side bleed assembly. The hot side bleed assembly is also in thermal communication with a hot component of the turbine to cool the hot component of the turbine.

[0216] As will be further appreciated, during engine operation, both fuel and compressed air flow through the fuel nozzles of the combustion section of the gas turbine engine, and the fuel nozzles are bathed in an external flow of relatively cool compressor discharge air. These flows carry heat away from the fuel nozzles, keeping the fuel temperature relatively low. However, when engine operation ceases, an amount of fuel will continue to reside in the fuel nozzles and can be heated to a coking temperature. A small amount of coke interfering with the flow of fuel through these orifices can make a large difference in fuel nozzle performance. Additionally, when engine operation ceases, heat within the combustion section, and, for example, the downstream end of the compressor section and the upstream end of the turbine section, tends to rise, creating a thermal mismatch along the vertical direction. This can create a bow in one or more rotors of the gas turbine engine (a rotor bow condition).

[0217] An air cooling system, referred to as a reverse bleed system (also referred to herein as an RBS), can be configured to reduce or minimize coke formation within, for example, the fuel nozzles of a gas turbine engine, and can also be configured to reduce or minimize rotor bow conditions within the engine. The reverse bleed system generally includes a reverse bleed duct extending from a cold location to a hot location, such as a compressor discharge plenum, and a blower operable during a shutdown operating condition of the gas turbine engine to circulate a gas stream through a working gas flow path of the gas turbine engine and reduce, for example, rotor bow conditions.

[0218] The inventors of the present disclosure have surprisingly discovered that the blower of the reverse bleed system can provide sufficient pressurization of a low pressure gas stream to use the low pressure gas stream for a CCA system during non-shutdown operating conditions of the gas turbine engine. Such an arrangement can allow for improved overall engine efficiency by reducing or eliminating extraction of a gas stream from the working gas flow path of the engine downstream of one or more compression stages, and by utilizing components that are not used during non-shutdown operating conditions of the gas turbine engine.

[0219] In particular, reference is now made to Figure 31FIG. 1 provides a schematic cross-sectional view of a gas turbine engine 10 having a CCA system 100 in accordance with another example embodiment of the present disclosure. Figures 1 to 30 The example gas turbine engine 10 and CCA system 100 can be constructed in a similar manner as one or more of the example gas turbine engines 10 and CCA systems 100 described above with reference to Figure 31

[0220] For example, Figure 31 The example gas turbine engine 10 generally includes a fan section 14 and a turbomachine 16, where the turbomachine 16 includes a compressor section, a combustion section 26, and a turbine section that collectively at least partially define a working gas flow path 37. The CCA system 100 includes a cold side bleed assembly 108 that defines an inlet 114 and an outlet 118. A hot side bleed assembly 110 defines an inlet 122 that is in fluid communication with the working gas flow path 37 through the turbomachine 16, and the hot side bleed assembly 110 is also in thermal communication with hot components of the turbomachine 16 to cool the hot components of the turbomachine 16. During operation, a cooling gas stream 102 is provided through the cold side bleed assembly 108, and a cooled gas stream 124 is provided through the hot side bleed assembly 110. In addition, the CCA system 100 includes a CCA heat exchanger 112 that is in thermal communication with the cold side bleed assembly 108 and the hot side bleed assembly 110, and more particularly, in thermal communication with the cooling gas stream 102 through the cold side bleed assembly 108 and the cooled gas stream 124 through the hot side bleed assembly 110.

[0221] In addition, for embodiments of the gas turbine engine 10, and more particularly, the turbomachine 16 of the gas turbine engine 10, include a reverse bleed system 500 for providing a gas stream into the working gas flow path 37 of the turbomachine 16 during certain operating conditions, such as during a shutdown operating condition. The reverse bleed system 500 generally includes a reverse bleed duct 502 and a reverse bleed blower 504 (also referred to herein as an RBS blower) that is in fluid communication with the reverse bleed duct 502. In addition, the reverse bleed duct 502 is in fluid communication with the working gas flow path 37. Figure 31

[0222] ​​More specifically, the compressor section, the combustion section 26, or both define a compressor bleed plenum 506 that is located radially outward of the working gas flowpath 37 along the radial direction R of the gas turbine engine 10 proximate a downstream end of the HP compressor 24 of the compressor section. For the depicted embodiment, the reverse bleed duct 502 is in fluid communication with the working gas flowpath 37 through the compressor bleed plenum 506. For example, the compressor bleed plenum 506 can be fluidly connected to the working gas flowpath 37 through one or more bleed ports (not shown). In certain example embodiments, the reverse bleed duct 502 can be fluidly coupled with the working gas flowpath 37 through a compressor bleed duct, such as the transient bleed duct 240 as discussed above.

[0223] Further, during operation of the reverse bleed system 500, the reverse bleed duct 502 is in fluid communication with a cold location of the gas turbine engine 10, as in the airflow source or the reverse bleed system 500. In particular, for the depicted embodiment, it will be appreciated that the gas turbine engine 10 defines a bypass passage 56 above the turbine 16, and the cold location (and the airflow source of the reverse bleed system 500) is the bypass passage 56, the working gas flowpath 37 through the LP compressor 22 of the compressor section, the working gas flowpath 37 between the LP compressor 22 and the HP compressor 24, or a combination thereof. Still more specifically, for the depicted embodiment, the turbine 16 includes an operability bleed assembly 180 having an operability bleed duct 182 extending between an inlet 184 in fluid communication with the working gas flowpath 37 and an outlet 186 in fluid communication with the bypass passage 56. The inlet 184 is in fluid communication with the working gas flowpath 37 at a location between the LP compressor 22 and the HP compressor 24. For the depicted embodiment, the reverse bleed duct 502 defines an inlet 508 in fluid communication with the operability bleed duct 182. For this configuration, the cold location is the working gas flowpath 37 between the LP compressor 22 and the HP compressor 24 when the gas turbine engine 10 is in the non-shutdown operating condition, and the cold location is the bypass passage 56 when the gas turbine engine 10 is in the shutdown operating condition.

[0224] The reverse bleed air system 500 can be configured to operate during a shutdown operating condition of the gas turbine engine 10. Specifically, following flight operation of the gas turbine engine 10 (a non-shutdown operating condition), the gas turbine engine 10 can be shut down (e.g., a shutdown operating condition) and airflow through the working gas flow path 37 can be stopped. During a shutdown operating condition, when airflow over certain components (e.g., the downstream end of the HP compressor 24, the combustion section 26, the upstream end of the HP turbine 28, or a combination thereof) is stopped, heat within these components can rise substantially vertically. This can at least result in a circumferential temperature mismatch within the gas turbine engine 10, thereby creating bowing or bending in certain components. However, during a shutdown operating condition, the reverse bleed air system 500 can generate airflow through the working gas flow path 37 to counteract this effect.

[0225] For example, in response to determining that the gas turbine engine 10 is in a shutdown operating condition, the gas turbine engine 10 may operate the reverse bleed air blower 504 to generate airflow from an inlet 508 (in fluid communication with a cold location of the gas turbine engine 10, such as the bypass passage 56), through the reverse bleed air duct 502, through an outlet of the reverse bleed air duct 502, and into the working gas flow path 37 (e.g., through the compressor bleed air plenum 506), through the downstream end of the HP compressor 24, the combustion section 26, or both. This configuration may minimize the risk of these components bowing or bending during shutdown operating conditions.

[0226] Notably, the reverse bleed air duct 502 includes a check valve 512 downstream of the reverse bleed air blower 504 and upstream of an outlet (not labeled) to ensure that during non-shutdown operating conditions of the gas turbine engine 10 (e.g., where pressurized air flow is provided through the working gas flow path 37), air flow from the working gas flow path 37 does not flow backward through the reverse bleed air duct 502.

[0227] It should also be understood that in certain embodiments, the Figure 31 The reverse bleed system 500 schematically depicted in FIG. 1 may be constructed in a manner similar to that of U.S. patent application Ser. No. 16 / 800,677, filed on February 25, 2020 (issued as U.S. Patent No. 11,047,306; referred to herein as the '306 patent), which is incorporated herein by reference in its entirety for all purposes. In particular, the reverse bleed system 500 of the present disclosure may be constructed in a manner similar to the reverse bleed device of the '306 patent (e.g., may have an inlet fluidly coupled to an operable bleed conduit / bleed conduit and an outlet fluidly connected to a working gas flow path through the bleed conduit, the compressor bleed port, and the plenum).

[0228] As will be appreciated, during non-shutdown operating conditions of the gas turbine engine 10, the reverse bleed system 500 need not be operated. Accordingly, in order to use certain aspects of the reverse bleed system 500 during non-shutdown operating conditions, the inlet 114 of the cold side bleed assembly 108 of the CCA system 100 is in fluid communication with the reverse bleed duct 502 of the reverse bleed system 500. In particular, the CCA system 100 further includes a flow control valve 415. The inlet 114 of the cold side bleed assembly 108 is in fluid communication with the reverse bleed duct 502 through the flow control valve 415. Still more particularly, the flow control valve 415 is positioned within the reverse bleed duct 502 of the reverse bleed system 500, and the inlet 114 of the cold side bleed assembly 108 is in fluid communication with the reverse bleed duct 502 through the flow control valve 415.

[0229] The flow control valve 415 can be moved between a first position, in which the flow of gas through the reverse bleed duct 502 is directed into the working gas flow path 37, and a second position, in which the flow of gas through the reverse bleed duct 502 is directed into the cold side bleed assembly 108 of the CCA system 100. In this way, the flow control valve 415 can be moved to the first position during shutdown operating conditions of the gas turbine engine 10, and can be moved to the second position during non-shutdown operating conditions of the gas turbine engine 10.

[0230] Additionally, the flow control valve 415 can be moved to one or more intermediate positions to split the flow of gas through the reverse bleed duct 502 to both the working gas flow path 37 and the cold side bleed assembly 108. This can be useful when, for example, the CCA heat exchanger 112 remains relatively hot after, for example, a flight operating condition of the gas turbine engine 10, as the split flow can both cool the CCA heat exchanger 112 and provide a flow of gas to the working gas flow path 37 to combat rotor bowing conditions.

[0231] Furthermore, in other embodiments, the flow control valve 415 can be moved to a closed position. In the closed position, the flow of gas through the reverse bleed duct 502 can be prevented from reaching either the working gas flow path 37 or the cold side bleed assembly 108.

[0232] Still referring to FIG. 1, the CCA system 100 further includes a cold side bleed assembly 108. The cold side bleed assembly 108 is in fluid communication with the working gas flow path 37 through the inlet 114 of the cold side bleed assembly 108. In particular, the cold side bleed assembly 108 is in fluid communication with the working gas flow path 37 through the inlet 114 of the cold side bleed assembly 108 and the flow control valve 415. Figure 32In this way, the reverse bleed air blower 504 of the reverse bleed air system 500 can be operable to increase the pressure of the airflow through the reverse bleed air duct 502 and provided to the cold side bleed air assembly 108 of the CCA system 100, thereby increasing the operability of the CCA system 100 during certain operations. This arrangement can allow the CCA system 100 to utilize relatively low pressure airflow from, for example, the bypass passage 56, thereby resulting in increased efficiency of the gas turbine engine. Alternatively, this arrangement can allow the CCA system 100 to be designed with a more efficient CCA heat exchanger 112, the more efficient CCA heat exchanger 112 having a higher pressure drop and more efficiently exchanging heat from the cooled airflow to the cooling airflow.

[0233] In this way, the reverse bleed air blower 504 of the reverse bleed air system 500 can be operable to increase the pressure of the airflow through the reverse bleed air duct 502 and provided to the cold side bleed air assembly 108 of the CCA system 100, thereby increasing the operability of the CCA system 100 during certain operations. This arrangement can allow the CCA system 100 to utilize relatively low pressure airflow from, for example, the bypass passage 56, thereby resulting in increased efficiency of the gas turbine engine. Alternatively, this arrangement can allow the CCA system 100 to be designed with a more efficient CCA heat exchanger 112, the more efficient CCA heat exchanger 112 having a higher pressure drop and more efficiently exchanging heat from the cooled airflow to the cooling airflow.

[0234] It will be appreciated, however, that other suitable configurations can be provided in other example embodiments. For example, in other embodiments, the cold location of the gas turbine engine 10 can be any of the cold locations of the gas turbine engine 10 described above. For example, now briefly referring to Figure 33 A schematic view of a gas turbine engine 10 having a reverse bleed air system 500 and a CCA system 100 according to another example embodiment of the present disclosure is provided, the cold location can be the bypass passage 56 of the gas turbine engine 10.

[0235] Further, now referring to Figure 33 A schematic view of a gas turbine engine 10 having a reverse bleed air system 500 and a CCA system 100 according to another example embodiment of the present disclosure is provided. Figure 31 The example gas turbine engine 10 of Figure 27 or Figure 8 may be configured in a similar manner to the example gas turbine engine 10 of

[0236] For example, Figure 33 The example reverse bleed air system 500 depicted in FIG. 1 includes a reverse bleed air duct 502 and a reverse bleed air blower 504 in fluid communication with the reverse bleed air duct 502. Similarly, Figure 32The depicted CCA system 100 includes a cold side bleed assembly 108, a hot side bleed assembly 110, and a CCA heat exchanger 112. For the depicted embodiment, however, the reverse bleed conduit 502 includes a first portion 514 and a second portion 516, and similarly, the cold side bleed assembly 108 includes a first portion 518 and a second portion 520. For the depicted embodiment, the first portion 514 of the cold side bleed assembly 108 is integrated with the first portion 518 of the reverse bleed conduit 502. In this manner, the inlet 508 of the reverse bleed conduit 502 is co-located with the inlet 114 of the cold side bleed assembly 108. As will be appreciated, for the depicted embodiment, the CCA heat exchanger 112 is positioned within the first portion 514 of the cold side bleed assembly 108, and the reverse bleed blower 504 is positioned within the first portion 518 of the reverse bleed conduit 502 at a location downstream of the CCA heat exchanger 112. The CCA system 100 also includes a flow control valve 415, wherein the second portion 516 of the cold side bleed assembly 108 is coupled to the first portion 514 of the cold side bleed assembly 108 by the flow control valve 415. Similarly, the second portion 520 of the reverse bleed conduit 502 is in fluid communication with the first portion 518 of the reverse bleed conduit 502 by the flow control valve 415.

[0237] For this configuration, operation of the blower 420 during, for example, a shutdown operating condition necessarily generates a flow of gas through the CCA heat exchanger 112 during such operating conditions, which can cool the CCA heat exchanger 112.

[0238] The flow control valve 415 and the reverse bleed blower 504 can operate similarly to the example flow control valve 415 and reverse bleed blower 504 discussed above with reference to Figure 2 The example flow control valve 415 and reverse bleed blower 504 operate. Notably, in certain example embodiments, the flow control valve 415, the reverse bleed blower 504, or both can be operably coupled to a controller, such as the example controller 130 described above with reference to Figure 34

[0239] Referring now to Figures 31 to 33 , a method 600 of operating a gas turbine engine in accordance with example aspects of the present disclosure is provided. The method 600 can be used with one or more of the example gas turbine engines described above with reference to Figure 35

[0240] ​​The method 600 includes receiving, at (602), data indicative of an operating condition of the gas turbine engine, and actuating, at (604), a flow control valve of a cooled cooling air (CCA) system to distribute airflow from a reverse bleed assembly of the CCA system between a cold side bleed assembly of the CCA system and a working gas flow path of a turbine of the gas turbine engine in response to receiving the data indicative of the operating condition of the gas turbine engine.

[0241] In certain example embodiments, the operating condition is a shutdown operating condition. For this example aspect, actuating, at (604), the flow control valve of the cooled cooling air system includes actuating, at (606), the flow control valve of the cooled cooling air system to distribute the airflow from the reverse bleed conduit to the working gas flow path of the turbine of the gas turbine engine. This arrangement can reduce the risk of rotor bowing or bending.

[0242] Briefly, it will be appreciated that when the operating condition is a shutdown operating condition, actuating, at (604), the flow control valve of the cooled cooling air system can also include actuating, at (605), the flow control valve of the cooled cooling air system to distribute the airflow from the reverse bleed conduit to the working gas flow path of the turbine and the cold side bleed assembly of the cooling air system for a first amount of time, and actuating, at (607), the flow control valve of the cooled cooling air system to distribute the airflow from the reverse bleed conduit to only the working gas flow path of the turbine for a second amount of time after the first amount of time. In this manner, the method 600 can allow for cooling of a cooled cooling air heat exchanger of the cooled cooling air system after a shutdown of the gas turbine engine (e.g., a shutdown operating condition of the gas turbine engine), and then transition to a more complete rotor bowing mitigation.

[0243] Additionally, in other example aspects, the operating condition can be a flight operating condition (non-shutdown operating condition). For this example aspect, actuating, at (604), the flow control valve of the cooled cooling air system includes actuating, at (608), the flow control valve of the cooled cooling air system to distribute the airflow from the reverse bleed conduit to the cold side bleed assembly of the cooled cooling air system. This arrangement can allow for operation of the cooled cooling air system.

[0244] Notably, in certain example aspects, the power source for the blower can vary based on the operating condition of the gas turbine engine. For example, in certain example embodiments, the operating condition can be a first operating condition. For this example aspect, the example method 600 can also include powering, at (610), the blower of the reverse bleed system with a first power source in response to receiving, at (602), the data indicative of the gas turbine engine being in the first operating condition.

[0245] The method can also include receiving, at (612), data indicative of a second operating condition of the gas turbine, the second operating condition being different than the first operating condition, and powering, at (614), the blower of the reverse bleed air system with a second power source separate from the first power source in response to receiving the data indicative of the second operating condition.

[0246] For example, the first power source can be an engine power source, such as a generator of a gas turbine engine, an energy storage unit of a gas turbine engine, etc. The second power source can be an aircraft power source, a ground power source, or other power source separate from the gas turbine engine.

[0247] As discussed above, aspects of the present disclosure relate to a gas turbine engine having an active clearance control (ACC) system within a turbine section of the gas turbine engine to maintain a desired clearance between turbine rotor blades and a casing of a turbine of the turbine section. The ACC system can operate by providing a flow of gas onto the casing of the turbine to control thermal growth of the casing relative to the turbine rotor blades. Given the relatively low temperature of the flow of gas, certain ACC systems use a low pressure flow of gas, such as a bypass flow of gas. However, as discussed above, depending on, for example, the complexity of the ACC system, a higher pressure flow of gas can be required to drive the flow of gas through manifolds of the ACC system. Accordingly, a blower (e.g., pump) can be provided to push the flow of gas through the ACC system.

[0248] As will be further appreciated, a reverse bleed air system can be provided to reduce or minimize, for example, coke formation within a fuel nozzle of a gas turbine engine, as well as rotor bowing conditions within the engine. The reverse bleed air system generally includes a reverse bleed air duct extending from a cold location to a hot location, such as a compressor discharge cavity, and a blower operable during a shutdown operating condition of the gas turbine engine to circulate a flow of gas through a working gas flow path of the gas turbine engine and reduce, for example, rotor bowing conditions.

[0249] The inventors of the present disclosure have surprisingly discovered that the blower of the reverse bleed air system can provide sufficient pressurization of a low pressure flow of gas to use the low pressure flow of gas for the ACC system during non-shutdown operating conditions of the gas turbine engine. Such an arrangement can allow for improved overall engine efficiency by reducing or eliminating extraction of a flow of gas from the working gas flow path of the engine downstream of one or more compression stages, as well as by utilizing components that are not used during non-shutdown operating conditions of the gas turbine engine.

[0250] In particular, reference is now made to Figure 35 , providing a schematic cross-sectional view of a gas turbine engine 10 according to another example embodiment of the present disclosure. Figures 1 to 34The exemplary gas turbine engine 10 may be configured as described above with reference to Figure 35 One or more of the exemplary gas turbine engine 10 and CCA system 100 are configured in a similar manner.

[0251] For example, Figure 35 The exemplary gas turbine engine 10 generally includes a fan section 14 and a turbine 16, wherein the turbine 16 includes a compressor section, a combustion section 26, and a turbine section that collectively at least partially define a working gas flow path 37. The turbine section includes a first turbine (or more specifically, an HP turbine 28 in the illustrated embodiment) and a second turbine (or more specifically, an LP turbine 30 in the illustrated embodiment).

[0252] However, for Figure 35 In an embodiment of the present invention, the turbine 16 further includes a reverse bleed air system 500 and an ACC system 400. The reverse bleed air system 500 can be constructed in a manner similar to one or more of the exemplary reverse bleed air systems 500 described above. For example, the reverse bleed air system 500 includes a reverse bleed air duct 502 and a reverse bleed air blower 504 in fluid communication with the reverse bleed air duct 502. The reverse bleed air duct 502 is also in fluid communication with the working gas flow path 37 for providing air flow to the working gas flow path 37 during operation of the reverse bleed air system 500. In addition, an inlet 508 of the reverse bleed air duct 502 is in fluid communication with a cold position of the gas turbine engine 10, for Figure 2 In the exemplary embodiment, the cold location is the bypass passage 56 of the gas turbine engine 10 .

[0253] However, in other embodiments, the cold location may be one or more of the other example cold locations described above.

[0254] Similarly, the ACC system 400 can be configured in a manner similar to one or more of the exemplary ACC systems 400 described above. For example, the ACC system 400 generally includes an inlet 402, a heat transfer assembly disposed around a turbine of the turbine section, and an ACC duct assembly 404 extending from the inlet 402 to the heat transfer assembly. More specifically, for the illustrated embodiment, the turbine is the HP turbine 28, and the heat transfer assembly is a first heat transfer assembly 406. The exemplary ACC system 400 also includes a second heat transfer assembly 408 disposed around the LP turbine 30. The ACC duct assembly 404 extends to the first heat transfer assembly 406 and the second heat transfer assembly 408.

[0255] In particular, for the illustrated embodiment, the ACC duct assembly 404 includes a first portion 410 that extends to the first heat transfer assembly 406 and a second portion 412 that extends to the second heat transfer assembly 408. The first portion 410 and the second portion 412 each include a separate inlet 402 (fluidly coupled to the reverse bleed duct 502, as discussed in greater detail below). In brief, it will be appreciated that, for the depicted embodiment, the ACC system 400 includes one or more flow control valves that are operable with the first portion 410 of the ACC duct assembly 404, the second portion 412 of the ACC duct assembly 404, or both. More specifically, the ACC system 400 includes a first flow control valve 416 that is operable with the first portion 410 of the ACC duct assembly 404, and a second flow control valve 418 that is operable with the second portion 412 of the ACC duct assembly 404. The first flow control valve 416, the second flow control valve 418, and the flow control valve 415 can each be operably connected to a controller (such as the exemplary controller 130) for controlling operation of the respective valve and gas flow through the respective duct. Figure 35

[0256] In the depicted embodiment, to more fully utilize aspects of the reverse bleed system 500 under various operating conditions of the gas turbine engine 10, the inlets 402 of the ACC system 400 are in fluid communication with the reverse bleed duct 502 of the first bleed system. As noted above, the reverse bleed system 500 can be operated during shutdown operating conditions of the gas turbine engine 10. In contrast, the ACC system 400 can be operated during non-shutdown operating conditions of the gas turbine engine 10. Thus, for the exemplary configuration of Figure 35

[0257] Still referring to Figure 36 More specifically, the first portion 410 is in fluid communication with the reverse bleed duct 502 at a first location along the reverse bleed duct 502, and the second portion 412 is in fluid communication with the reverse bleed duct 502 at a second location along the reverse bleed duct 502. In the illustrated embodiment, the reverse bleed blower 504 is in fluid communication with the first bleed duct at a third location between the first location and the second location. In this manner, the reverse bleed duct 502 can be configured to increase the pressure of the gas flow provided to the second heat transfer assembly 408 around the LP turbine 30 during non-shutdown operating conditions of the gas turbine engine 10. As discussed above, this can allow the ACC system 400 to be operated in a desired manner, given the complexity of the second heat transfer assembly 408 around the LP turbine 30.

[0258] ​​Furthermore, it will be appreciated that, for the exemplary ACC system 400 depicted, the second flow control valve 418 is positioned in the second position and is configured as a variable three-way valve. In this manner, the second flow control valve 418 can be configured to regulate airflow between the reverse bleed air conduit 502 (and more specifically, the section of the reverse bleed air conduit 502 extending from the second position to the working gas flow path 37) and the second portion 412 of the ACC conduit assembly 404. Consequently, the second portion 412 of the ACC conduit assembly 404 is in fluid communication with the reverse bleed air conduit 502 via the second flow control valve 418.

[0259] When the gas turbine engine 10 is in a shutdown operating condition, the first flow control valve 416 can be moved to a closed position, thereby preventing air flow through the first portion 410 of the ACC duct assembly 404. Also during the shutdown operating condition, the second flow control valve 418 can be moved to a position in which all or substantially all of the air flow from the reverse bleed air blower 504 through the reverse bleed air duct 502 is provided through the second position and maintained to flow through the reverse bleed air duct 502 to the working gas flow path 37.

[0260] In contrast, when the gas turbine engine 10 is in a non-shutdown operating condition, the first flow control valve 416 can be moved to an open position, thereby allowing airflow through the first portion 410 of the ACC duct assembly 404 to flow therethrough. Also during non-shutdown operating conditions, the second flow control valve 418 can be moved to a different position, wherein all or substantially all of the airflow from the reverse bleed air blower 504 through the reverse bleed air duct 502 is provided to the second portion 412 of the ACC duct assembly 404 and allowed to flow from the reverse bleed air duct 502 through the second portion 412 of the ACC duct assembly 404 (and to the second heat transfer assembly 408). Notably, the reverse bleed air system 500 includes a check valve 512 within the reverse bleed air duct 502 at a location downstream of the inlet 402 of the second portion 412 of the ACC duct 404. The check valve 512 is positioned to prevent airflow from the outlet of the reverse bleed air duct 502 from flowing upstream toward the reverse bleed air blower 504.

[0261] However, it will be understood that in other exemplary embodiments, other suitable configurations may be provided. Figure 37 , provides a schematic diagram of a gas turbine engine 10 having a reverse bleed air system 500 and an ACC system 400 according to another exemplary embodiment of the present disclosure, which may provide different valve arrangements. For the depicted embodiment, the second portion 412 of the ACC piping assembly 404 includes a second flow control valve 418 downstream of the second location, and the reverse bleed air system 500 includes a separate flow control valve 522 downstream of the second location.

[0262] Further, now referring to Figure 35 , another exemplary embodiment of a gas turbine engine 10 is provided in accordance with embodiments of the present disclosure, which can again be constructed in a similar manner as the exemplary gas turbine engine 10 described above with reference to Figure 38 . Like or similar numbers can refer to like or similar parts.

[0263] For example, the exemplary gas turbine engine 10 generally includes a reverse bleed system 500 and an ACC system 400. An inlet 402 of the ACC system 400 is in fluid communication with a reverse bleed duct 502 of the reverse bleed system 500. However, for the depicted embodiment, an ACC duct assembly 404 of the ACC system 400 includes a first portion 410 extending to a first heat transfer assembly 406 of the ACC system 400, a second portion 412 extending to a second heat transfer assembly 408 of the ACC system 400, and an upstream portion 414 extending from the inlet 402 of the ACC system 400 to the first and second portions 410, 412. For this exemplary embodiment, the reverse bleed system 500 further includes a flow control valve 522, wherein the inlet 402 of the ACC duct assembly 404 is in fluid communication with the reverse bleed duct 502 through the flow control valve 522. The flow control valve 522 can be moved between a first position, in which a flow of gas from a reverse bleed blower 504 of the reverse bleed system 500 is provided to the ACC duct assembly 404 through the reverse bleed duct 502 and the inlet 402 (e.g., during non-shutdown operating conditions), and a second position, in which the flow of gas from the reverse bleed blower 504 is provided to a working gas flow path 37 of the gas turbine engine 10 through the reverse bleed duct 502 (e.g., during shutdown operating conditions).

[0264] Such a configuration can result in an overall more efficient engine by more fully utilizing components throughout the operating window of the gas turbine engine.

[0265] As discussed above, aspects of the present disclosure relate to a reverse bleed system that can be provided to reduce or minimize coke formation within one or more portions of a gas turbine engine (e.g., a fuel nozzle of the gas turbine engine), and / or rotor bowing conditions within the engine. The reverse bleed system generally includes a reverse bleed duct extending from a cold location to a hot location (such as a compressor discharge plenum) and a blower that is operable during shutdown operating conditions of the gas turbine engine to circulate a flow of gas through a working gas flow path of the gas turbine engine and reduce, for example, rotor bowing conditions.

[0266] Additionally, various accessory systems of the gas turbine engine, and particularly various accessory systems of the turbine of the gas turbine engine, require cooling during non-shut down operating conditions. Many of these accessory systems are located in the undercowl region (the region below the outer casing of the turbine and outside the working gas flow path of the gas turbine engine). In particular, cooling of these accessory systems in the undercowl region can become more difficult as the engine is pushed towards more efficient (and hotter) operation.

[0267] However, the inventors of the present disclosure have discovered that a blower utilizing a reverse bleed system can provide sufficient pressurization of the low pressure cold air stream to allow for additional cooling of one or more of these accessory systems. In particular, by pressurizing the low pressure cold air, more airflow and more impingement cooling can be possible for these accessory systems. This arrangement can allow for increased overall engine efficiency by more fully utilizing components that are not used during non-shut down operating conditions of the gas turbine engine to provide additional cooling for certain engine systems.

[0268] In particular, referring now to Figure 38 , a schematic cross-sectional view of a gas turbine engine 10 according to another example embodiment of the present disclosure is provided. Figures 1 to 37 The example gas turbine engine 10 of Figure 38 may be constructed in a similar manner as one or more of the example gas turbine engines 10 described above with reference to

[0269] For example, Figure 38 The example gas turbine engine 10 generally includes a fan section 14 and a turbine 16, where the turbine 16 includes a compressor section, a combustion section 26, and a turbine section that collectively at least partially define a working gas flow path 37. The turbine 16 additionally includes an operability bleed assembly 180 having an operability bleed conduit 182 that extends between an inlet 184 in fluid communication with the working gas flow path 37 and an outlet 186 in fluid communication with a bypass passage 56 of the gas turbine engine 10. In particular, the inlet 184 of the operability bleed conduit 182 is in fluid communication with the working gas flow path 37 at a location between the LP compressor 22 and the HP compressor 24.

[0270] Further, the depicted exemplary turbine 16 includes a reverse bleed system 500. The reverse bleed system 500 can be constructed in a similar manner as one or more of the exemplary reverse bleed systems 500 described above. For example, the reverse bleed system 500 generally includes a reverse bleed duct 502 and a reverse bleed blower 504 in fluid communication with the reverse bleed duct 502. The reverse bleed duct 502 is also in fluid communication with the working gas flow path 37 for providing a flow of air to the working gas flow path 37 during operation of the reverse bleed system 500.

[0271] As with one or more of the exemplary embodiments discussed above, Figure 38 The exemplary embodiments include an apparatus for utilizing one or more aspects of the reverse bleed system 500 during non-shutdown operating conditions of the gas turbine engine 10. In particular, the exemplary gas turbine engine 10 also includes an accessory system 700 and the turbine 16 also includes an accessory cooling system 702. The accessory cooling system 702 includes a cooling duct 704 that defines an inlet 706 in fluid communication with the reverse bleed duct 502. The cooling duct 704 is configured to provide a flow of air onto the accessory system 700 to cool the accessory system 700.

[0272] In the depicted embodiment, the accessory cooling system 702 also includes a flow control valve 708. The inlet 402 of the accessory cooling system 702 is in fluid communication with the reverse bleed duct 502 through the flow control valve 708. In particular, for the depicted embodiment, the flow control valve 708 is a variable three-way valve positioned within the reverse bleed duct 502 of the reverse bleed system 500 with the inlet 706 of the accessory cooling system 702 at the variable three-way valve. In this manner, the variable three-way valve can be moved between a first position and a second position. In the first position, the variable three-way valve can allow the flow of air through the reverse bleed duct 502 to flow past the variable three-way valve and remain within the reverse bleed duct 502. In contrast, in the second position, the variable three-way valve diverts the flow of air through the reverse bleed duct 502 into the cooling duct 704 of the accessory cooling system 702. The variable three-way valve can be in the first position when the gas turbine engine 10 is in a shutdown operating condition and in the second position when the gas turbine engine 10 is in a non-shutdown operating condition.

[0273] In the depicted embodiment, the reverse bleed air blower 504 is in fluid communication with the reverse bleed air duct 502 at a location upstream of the flow control valve 708. The reverse bleed air blower 504 can be configured to increase the pressure of the airflow received from the inlet 114 of the reverse bleed air duct 502 through the reverse bleed air duct 502. The inlet 114 of the reverse bleed air duct 502 is in fluid communication with a cold location of the gas turbine engine 10. For the depicted embodiment, the cold location is the bypass passage 56 of the gas turbine engine 10.

[0274] However, as depicted in dashed lines, in other example embodiments, the cold location can be one or more other suitable cold locations, such as one or more of the other cold locations discussed above, such as the operability bleed air duct 182 of the turbine 16.

[0275] Further, for embodiments in which the accessory system 700 is the first accessory system 700A, and the gas turbine engine 10 further includes a second accessory system 700B, the cooling duct 704 includes a first portion 710 extending toward the first accessory system 700A and in fluid communication with the reverse bleed air duct 502 at a first location along the reverse bleed air duct 502. In particular, the first portion 710 is in fluid communication with the reverse bleed air duct 502 through the flow control valve 708. Figure 39

[0276] The cooling duct 704 further includes a second portion 712 extending toward the second accessory system 700B and in fluid communication with the reverse bleed air duct 502 at a second location along the reverse bleed air duct 502. In the depicted embodiment, the reverse bleed air blower 504 is in fluid communication with the reverse bleed air duct 502 at a third location between the first location and the second location. In particular, the reverse bleed air blower 504 is in fluid communication with the reverse bleed air duct 502 at a location downstream of the second location and upstream of the first location. In this manner, the accessory cooling system 702 can be configured to provide passive airflow cooling to the second accessory system 700B, and can be configured to provide active or pressurized airflow cooling to the first accessory system 700A using the reverse bleed air blower 504, for example, during non-shutdown operating conditions.

[0277] Reference is now made to Figure 39 , providing a schematic view of a cooling duct 704 of an accessory cooling system 702 according to example aspects of the present disclosure. Figure 38 The example cooling duct 704 of the accessory cooling system 702 depicted in Figure 39 may be configured in a similar manner to the example cooling duct 704 of the accessory cooling system 702 described above with reference to

[0278] For example embodiments in which the accessory system 700 is the first accessory system 700A, and the gas turbine engine 10 further includes a second accessory system 700B, the cooling duct 704 includes a first portion 710 extending toward the first accessory system 700A and in fluid communication with the reverse bleed air duct 502 at a first location along the reverse bleed air duct 502. In particular, the first portion 710 is in fluid communication with the reverse bleed air duct 502 through the flow control valve 708. Figure 39 ​In embodiments of the present disclosure, the cooling duct 704 of the accessory cooling system 702 includes a cooling tip 714, where the cooling tip 714 is oriented toward the accessory system 700 to provide a cooling airflow 716 onto the accessory system 700. The cooling tip 714 can include a nozzle or diffuser configuration to jet the flow at the accessory system to increase local cooling or to diffuse the flow to cover a wide area of the accessory system.

[0279] More specifically, in Figure 2 In embodiments of the present disclosure, the accessory system 700 is a first accessory system 700A of a plurality of accessory systems 700 of the gas turbine engine 10, and the cooling tip 714 is a first cooling tip 714A of a plurality of cooling tips 714. More specifically, the cooling duct 704 includes a plurality of branches 718, and each of the plurality of branches 718 includes (or defines) a respective one of the plurality of cooling tips 714. Each of the plurality of cooling tips 714 is oriented toward a respective one of the plurality of accessory systems 700 to provide a cooling airflow 716 onto the respective accessory system 700.

[0280] In this manner, the accessory cooling system 702 can use components of the reverse bleed system 500 to provide impingement cooling to various different accessory systems 700 of the gas turbine engine 10 that would otherwise not be operated during non-shutdown operating conditions of the gas turbine engine 10.

[0281] In certain example embodiments, the accessory systems 700 can be spaced apart along the axial and circumferential directions within a shroud down region of the turbine 16. The accessory systems 700 can include one or more of: a HP turbine ACC valve (e.g., the first flow control valve 416), a LP turbine ACC valve (e.g., the second flow control valve 418), a combustor igniter system, a core compartment cooling valve, a booster anti-ice valve, a nacelle anti-ice valve, a start bleed valve, a transient bleed valve, a regulated turbine cooling valve, an engine controller (such as the controller 130 in Figure 40

[0282] Reference is now made to Figure 40 , providing a schematic view of a gas turbine engine 10 according to another example embodiment of the present disclosure. Figure 38 The example gas turbine engine 10 of Figure 40 may be constructed in substantially the same manner as the example gas turbine engine 10 of

[0283] For example, Figure 40 ​The exemplary gas turbine engine 10 includes an accessory system 700 and a turbine machine 16 having a reverse bleed air system 500 and an accessory cooling system 702 for cooling the accessory system 700. However, for embodiments of the present disclosure, the accessory system 700 is a first accessory system 700A and the gas turbine engine 10 further includes a second accessory system 700B. Moreover, the accessory cooling system 702 includes a cooling conduit 704 having a first portion 710 extending toward the first accessory system 700A and in fluid communication with the reverse bleed air conduit 502 at a first location along the reverse bleed air conduit 502, and a second portion 712 extending toward the second accessory system 700B and in fluid communication with a cool location of the gas turbine engine 10 separate from the reverse bleed air system 500. In particular, for the depicted embodiment, the reverse bleed air system 500 includes a reverse bleed air conduit 502 having an inlet 508 in fluid communication with the bypass passage 56 of the gas turbine engine 10, and the cool location with which the second portion 712 of the cooling conduit 704 is in fluid communication is the operability bleed air conduit 182 of the turbine machine 16. Figure 41

[0284] However, it is noted that in other exemplary embodiments, the accessory cooling system 702 and the reverse bleed air system 500 can have other configurations. For example, it will be appreciated that in other embodiments, the accessory cooling system 702 can include a flow control valve 708 downstream of a second location along the reverse bleed air conduit 502 (e.g., a location at which the accessory cooling system 702 is in fluid communication with the reverse bleed air conduit 502). For such exemplary aspects, the second portion 712 of the accessory cooling system 702 can be in fluid communication with the reverse bleed air conduit 502 through the flow control valve 708, and the reverse bleed air system 500 can further include a separate RBS flow control valve (not shown herein) downstream of the second location.

[0285] It will be appreciated that the embodiments provided above are provided by way of example only. Features of each of these embodiments can be combined with features of any other embodiment to create additional embodiments within the scope of the present disclosure, unless it is clear from the context that it would not be combinable.

[0286] It will also be appreciated that while the exemplary embodiments described above are described in the context of a turbofan engine (e.g., having a fan, a turbine machine, and an outer nacelle surrounding at least a portion of the fan and turbine machine), in other exemplary embodiments, the embodiments described herein can also be incorporated into an open rotor engine (e.g., having a non-ducted fan driven by a turbine machine). ​ ​An example of such an engine is provided, providing a schematic view of an open rotor engine 800 having an un-tubulated fan 802 and a turbine 804. The turbine 804 generally includes a compressor section 806, a combustion section 808, and a turbine section 810 arranged in serial flow order. The turbine 804 also includes an outer casing 812 surrounding the compressor section 806, the combustion section 808, and the turbine section 810. The outer casing 812 partially defines an undercowl region 814, which can house one or more of an ACC system, a CCA system, a reverse bleed system, an accessory cooling system, etc.

[0287] Further aspects are provided by the subject matter of the following clauses:

[0288] A gas turbine engine comprising: a turbine including a compressor section having a low pressure compressor and a high pressure compressor, a combustion section defining a compressor discharge plenum, and a turbine section, the turbine defining a working gas flow path and further including a cooled cooling air (CCA) system, the CCA system including: a cold side bleed assembly defining an inlet in fluid communication with the working gas flow path through the compressor section at a location between or both of the low pressure compressor, the low pressure compressor and the high pressure compressor; a CCA heat exchanger in thermal communication with the cold side bleed assembly downstream of the inlet of the cold side bleed assembly; and a hot side bleed assembly defining an inlet in fluid communication with the working gas flow path through the compressor section, at the compressor discharge plenum, or both, the hot side bleed assembly in thermal communication with the CCA heat exchanger to cool a flow through the hot side bleed assembly, the hot side bleed assembly further in thermal communication with a hot component of the turbine to cool the hot component of the turbine.

[0289] The gas turbine engine according to any preceding clause, wherein the turbine section includes a high pressure turbine, and wherein the hot component is the high pressure turbine.

[0290] The gas turbine engine according to any preceding clause, wherein the hot component is a rotor bore of a turbine of the turbine section, an airfoil of the turbine, a rotor bore of the high pressure compressor, an airfoil of the high pressure compressor, an oil sump within the turbine section, a turbine mid-frame, a turbine aft-frame, or a combination thereof.

[0291] The gas turbine engine according to any preceding clause, wherein the turbine defines an undercowl plenum, and wherein the CCA heat exchanger is located in the undercowl plenum of the turbine.

[0292] The gas turbine engine of any preceding paragraph, wherein the gas turbine engine includes a bifurcation defining a bifurcation cavity, wherein the turbine defines a undercowl cavity, and wherein the CCA heat exchanger is located in the undercowl cavity of the turbine, the bifurcation cavity, or both.

[0293] The gas turbine engine of any preceding paragraph, wherein the gas turbine engine is a turbofan engine including a fan driven by the turbine and an outer nacelle surrounding the fan and defining a bypass passage with the turbine.

[0294] The gas turbine engine of any preceding paragraph, wherein the turbine includes an operability bleed assembly defining an operability bleed conduit extending between an inlet and an outlet, the inlet in fluid communication with the working gas flow path at a location between the low pressure compressor and the high pressure compressor, wherein the inlet of the cold side bleed assembly is in fluid communication with the operability bleed conduit.

[0295] The gas turbine engine of any preceding paragraph, wherein the operability bleed assembly further includes a variable bleed valve in communication with the operability bleed conduit at a location upstream of the inlet of the cold side bleed assembly.

[0296] The gas turbine engine of any preceding paragraph, wherein the inlet is in direct fluid communication with the working gas flow path.

[0297] The gas turbine engine of any preceding paragraph, wherein the CCA heat exchanger is a first CCA heat exchanger, wherein the CCA system further includes a second CCA heat exchanger, wherein the cold side bleed assembly includes a first portion and a second portion in a parallel flow arrangement, wherein the first portion is in thermal communication with the first CCA heat exchanger and the second portion is in thermal communication with the second CCA heat exchanger.

[0298] The gas turbine engine of any preceding paragraph, wherein the hot side bleed assembly is a first hot side bleed assembly and is in thermal communication with the first CCA heat exchanger, wherein the CCA system further includes a second hot side bleed assembly in thermal communication with the second CCA heat exchanger.

[0299] The gas turbine engine of any preceding paragraph, wherein the hot component is a first hot component, wherein the CCA heat exchanger is a first CCA heat exchanger, wherein the CCA system further comprises a third CCA heat exchanger, wherein the cold side bleed assembly comprises first and third portions in a parallel flow arrangement, wherein the first portion is in thermal communication with the first CCA heat exchanger and the third portion is in thermal communication with the third CCA heat exchanger, and wherein the hot side bleed assembly is a first hot side bleed assembly and is in thermal communication with the first CCA heat exchanger, wherein the CCA system further comprises a third hot side bleed assembly in thermal communication with the third CCA heat exchanger and a hot component of the gas turbine engine separate from the first hot component.

[0300] The gas turbine engine of any preceding paragraph, wherein the CCA system further comprises a flow control valve in operable communication with the cold side bleed assembly.

[0301] The gas turbine engine of any preceding paragraph, wherein the CCA system further comprises a flow control valve in operable communication with the hot side bleed assembly.

[0302] The gas turbine engine of any preceding paragraph, wherein the inlet of the hot side bleed assembly is in fluid communication with the working gas flow path at the compressor discharge plenum.

[0303] The gas turbine engine of any preceding paragraph, wherein the hot side bleed assembly comprises a first portion in thermal communication with the CCA heat exchanger and a bypass portion that bypasses the CCA heat exchanger.

[0304] The gas turbine engine of any preceding paragraph, wherein, when operating at rated speed during standard day operating conditions, the gas turbine engine defines a total pressure ratio greater than or equal to 50: 1 and less than or equal to 70: 1.

[0305] The gas turbine engine of any preceding paragraph, further comprising: a fan section having a fan driven by the turbine, wherein, when the engine is operating under cruise conditions, the fan defines a fan pressure ratio less than or equal to 1.6.

[0306] The gas turbine engine of any preceding paragraph, wherein the gas turbine engine is an open rotor gas turbine engine.

[0307] The gas turbine engine of any preceding paragraph, wherein the gas turbine engine is a turbofan engine.

[0308] A gas turbine engine comprising: a bifurcation; and a turbine machine comprising a compressor section, a combustion section defining a compressor discharge cavity, and a turbine section, the turbine machine defining a working gas flow path and further comprising a cooled cooling air (CCA) system, the CCA system comprising: a cold side bleed assembly defining an inlet positioned in fluid communication with an airflow above the bifurcation; a CCA heat exchanger in thermal communication with the cold side bleed assembly downstream of the inlet of the cold side bleed assembly; and a hot side bleed assembly defining an inlet in fluid communication with the working gas flow path through the compressor section, at the compressor discharge cavity, or both, the hot side bleed assembly in thermal communication with the CCA heat exchanger to cool an airflow through the hot side bleed assembly, the hot side bleed assembly further in thermal communication with a hot component of the turbine machine to cool the hot component of the turbine machine.

[0309] The gas turbine engine of any preceding paragraph, wherein the inlet is on a flow path surface of the bifurcation.

[0310] The gas turbine engine of any preceding paragraph, wherein the cold side bleed assembly comprises a scoop at the inlet.

[0311] The gas turbine engine of any preceding paragraph, wherein the cold side bleed assembly comprises a variable geometry component at the inlet.

[0312] The gas turbine engine of any preceding paragraph, wherein the bifurcation is an upper bifurcation of the gas turbine engine.

[0313] The gas turbine engine of any preceding paragraph, wherein the gas turbine engine is a turbofan engine comprising a fan driven by the turbine machine and an outer nacelle surrounding the fan and defining a bypass passage with the turbine machine, wherein the upper bifurcation extends between the turbine machine and the outer nacelle.

[0314] The gas turbine engine of any preceding paragraph, further comprising: an accessory system cooling system comprising a cooling system inlet on the bifurcation and a duct in fluid communication with the cooling system inlet, wherein the inlet is in fluid communication with the duct of the accessory system cooling system at a location downstream of the cooling system inlet.

[0315] The gas turbine engine of any preceding paragraph, wherein the accessory system cooling system is an environmental control system.

[0316] The gas turbine engine according to any preceding paragraph, wherein the CCA system further comprises a flow control valve, wherein the flow control valve is located at the inlet to control a flow split between the duct of the accessory system cooling system and the cold side bleed assembly.

[0317] The gas turbine engine according to any preceding paragraph, wherein the turbine section comprises a high pressure turbine, and wherein the hot component is the high pressure turbine.

[0318] The gas turbine engine according to any preceding paragraph, wherein the hot component is a rotor bore of a turbine of the turbine section, an airfoil of the turbine, a rotor bore of the high pressure compressor, an airfoil of the high pressure compressor, an oil sump within the turbine section, a turbine mid-frame, a turbine aft-frame, or a combination thereof.

[0319] The gas turbine engine according to any preceding paragraph, wherein the turbomachine defines an undercowl cavity, and wherein the CCA heat exchanger is located in the undercowl cavity of the turbomachine.

[0320] The gas turbine engine according to any preceding paragraph, wherein the bifurcation defines a bifurcation cavity, wherein the turbomachine defines an undercowl cavity, and wherein the CCA heat exchanger is located in the undercowl cavity of the turbomachine, the bifurcation cavity of the bifurcation, or both.

[0321] The gas turbine engine according to any preceding paragraph, wherein the CCA heat exchanger is a first CCA heat exchanger, wherein the CCA system further comprises a second CCA heat exchanger, wherein the cold side bleed assembly comprises a first portion and a second portion in a parallel flow arrangement, wherein the first portion is in thermal communication with the first CCA heat exchanger and the second portion is in thermal communication with the second CCA heat exchanger.

[0322] The gas turbine engine according to any preceding paragraph, wherein the hot side bleed assembly comprises a first portion and a second portion in a parallel flow arrangement, wherein the first portion is in thermal communication with the first CCA heat exchanger and the second portion is in thermal communication with the second CCA heat exchanger.

[0323] The gas turbine engine according to any preceding paragraph, wherein the CCA system further comprises a flow control valve in operable communication with the cold side bleed assembly.

[0324] The gas turbine engine according to any preceding paragraph, wherein the CCA system further comprises a flow control valve in operable communication with the hot side bleed assembly.

[0325] The gas turbine engine according to any preceding paragraph, wherein, when operating at rated speed during standard day operating conditions, the gas turbine engine defines a total pressure ratio greater than or equal to 50: 1 and less than or equal to 70: 1.

[0326] The gas turbine engine according to any preceding paragraph, further comprising: a fan section having a fan driven by the turbine, wherein, when the engine is operating under cruise conditions, the fan defines a fan pressure ratio less than or equal to 1.6.

[0327] The gas turbine engine according to any preceding paragraph, wherein the inlet of the hot side bleed assembly is in fluid communication with the working gas flow path at the compressor discharge plenum.

[0328] A gas turbine engine comprising: an extension; and a turbine defining a compressor section, a combustion section defining a compressor discharge plenum, and a turbine section, the turbine defining a working gas flow path and further comprising a cooled cooling air (CCA) system, the extension coupled to and extending outwardly from the turbine, the CCA system comprising: a cold side bleed assembly defining an inlet positioned in fluid communication with an airflow above the extension; a CCA heat exchanger in thermal communication with the cold side bleed assembly downstream of the inlet of the cold side bleed assembly; and a hot side bleed assembly defining an inlet in fluid communication with the working gas flow path through the compressor section, the compressor discharge plenum, or both, the hot side bleed assembly in thermal communication with the CCA heat exchanger to cool an airflow through the hot side bleed assembly, the hot side bleed assembly further in thermal communication with hot components of the turbine to cool the hot components of the turbine.

[0329] A gas turbine engine comprising: a turbomachine comprising a compressor section, a combustion section defining a compressor discharge plenum, and a turbine section, the turbomachine defining a working gas flow path and further comprising a cooled cooling air (CCA) system, the CCA system comprising: a hot side bleed assembly defining an inlet in fluid communication with the working gas flow path through the compressor section, at the compressor discharge plenum, or both, the hot side bleed assembly further in thermal communication with hot components of the turbomachine to cool the hot components of the turbomachine; a CCA heat exchanger in thermal communication with the hot side bleed assembly downstream of the inlet of the hot side bleed assembly; and a cold side bleed assembly defining an inlet and an outlet, the inlet in fluid communication with a cold location of the gas turbine engine, the cold side bleed assembly in thermal communication with the CCA heat exchanger for cooling a gas stream through the hot side bleed assembly, the outlet of the cold side bleed assembly in fluid communication with the working gas flow path at a location downstream of the turbine section, in fluid communication with a core cowl vent of the turbomachine, in fluid communication with a tunnel above the turbomachine, or a combination thereof.

[0330] The gas turbine engine of any preceding paragraph, wherein the outlet of the cold side bleed assembly is in fluid communication with the core cowl vent of the turbomachine.

[0331] The gas turbine engine of any preceding paragraph, wherein the cold side bleed assembly comprises one or more ducts extending from the CCA heat exchanger toward the core cowl vent to a location at an aft end of a low pressure turbine of the turbine section, or to a location aft of the low pressure turbine of the turbine section.

[0332] The gas turbine engine of any preceding paragraph, wherein the turbomachine defines an undercowl flow path for gas flow through the outlet of the cold side bleed assembly to the core cowl vent, and wherein the turbomachine comprises a thermal shield between the undercowl flow path and a low pressure turbine of the turbine section.

[0333] The gas turbine engine of any preceding paragraph, wherein the cold side bleed assembly comprises a diffuser at the outlet.

[0334] The gas turbine engine according to any preceding clause, wherein the turbine includes a transient bleed assembly defining a transient bleed duct extending between an inlet and an outlet, wherein the inlet is in fluid communication with the working gas flow path at a high pressure compressor of the compressor section or the compressor discharge plenum, wherein the outlet of the cold side bleed assembly is in fluid communication with the transient bleed duct.

[0335] The gas turbine engine according to any preceding clause, further comprising: a bifurcation connected to the turbine, wherein the outlet of the cold side bleed assembly is in fluid communication with an opening in a flow path surface of the bifurcation at a trailing edge of the bifurcation.

[0336] The gas turbine engine according to any preceding clause, wherein the cold location of the gas turbine engine is through a low pressure compressor, the working gas flow path between or both of the low pressure compressor and a high pressure compressor.

[0337] The gas turbine engine according to any preceding clause, further comprising: a bifurcation connected to the turbine, wherein the cold location of the gas turbine engine is a flow path surface of the bifurcation.

[0338] The gas turbine engine according to any preceding clause, wherein the cold side bleed assembly includes a scoop at the inlet.

[0339] The gas turbine engine according to any preceding clause, wherein the cold side bleed assembly includes a variable geometry component at the inlet.

[0340] The gas turbine engine according to any preceding clause, wherein the bifurcation is an upper bifurcation of the gas turbine engine.

[0341] The gas turbine engine according to any preceding clause, wherein the gas turbine engine defines a bypass passage above the turbine, and wherein the cold location is the bypass passage.

[0342] The gas turbine engine according to any preceding clause, wherein the turbine section includes a high pressure turbine, and wherein the hot component is the high pressure turbine.

[0343] The gas turbine engine according to any preceding clause, wherein the hot component is a rotor bore of a turbine of the turbine section, an airfoil of the turbine, a rotor bore of the high pressure compressor, an airfoil of the high pressure compressor, an oil sump within the turbine section, a turbine mid-frame, a turbine aft-frame, or a combination thereof.

[0344] The gas turbine engine according to any preceding clause, wherein the turbine defines a shroud down cavity, and wherein the CCA heat exchanger is located in the shroud down cavity of the turbine.

[0345] The gas turbine engine according to any preceding clause, wherein the bifurcation defines a bifurcation cavity, wherein the turbine defines a shroud down cavity, and wherein the CCA heat exchanger is located in the shroud down cavity of the turbine, the bifurcation cavity of the bifurcation, or both.

[0346] The gas turbine engine according to any preceding clause, wherein the CCA system further comprises a flow control valve in operable communication with the cold side bleed assembly.

[0347] The gas turbine engine according to any preceding clause, wherein the CCA system further comprises a flow control valve in operable communication with the hot side bleed assembly.

[0348] The gas turbine engine according to any preceding clause, wherein the inlet of the hot side bleed assembly is in fluid communication with the working gas flow path at the compressor discharge cavity.

[0349] The gas turbine engine according to any preceding clause, wherein the turbine includes a transient bleed assembly defining a transient bleed duct extending between an inlet and an outlet, wherein the inlet is in fluid communication with the hot side bleed assembly, wherein the outlet of the cold side bleed assembly is in fluid communication with the transient bleed duct.

[0350] The gas turbine engine according to any preceding clause, wherein the outlet of the transient bleed duct is in fluid communication with the hot side bleed assembly at a location upstream of the CCA heat exchanger.

[0351] A gas turbine engine comprising: a turbomachine comprising a compressor section, a combustion section defining a compressor discharge plenum, and a turbine section, the turbomachine further comprising a cooled cooling air (CCA) system, the CCA system comprising: a CCA heat exchanger; a cold-side bleed assembly defining an inlet and an outlet, the inlet being in fluid communication with a cold location of the gas turbine engine; a hot-side bleed assembly defining an inlet, the inlet being in fluid communication with the working gas flow path through the compressor section, at the compressor discharge plenum, or both, the CCA heat exchanger being in thermal communication with both the cold-side bleed assembly and the hot-side bleed assembly to cool a flow of air through the hot-side bleed assembly, the hot-side bleed assembly further being in thermal communication with a hot component of the turbomachine to cool the hot component of the turbomachine; and a valve assembly operable with the cold-side bleed assembly and the hot-side bleed assembly, the valve assembly being structured to cooperatively regulate a flow of air through the cold-side bleed assembly and the hot-side bleed assembly.

[0352] The gas turbine engine of any preceding paragraph, wherein the valve assembly comprises a cold-side valve section operable with the cold-side bleed assembly, a hot-side valve section operable with the hot-side bleed assembly, and an actuator coupled to both the cold-side valve section and the hot-side valve section.

[0353] The gas turbine engine of any preceding paragraph, wherein the valve assembly further comprises a shaft movable by the actuator, the actuator being coupled to both the cold-side valve section and the hot-side valve section through the shaft.

[0354] The gas turbine engine of any preceding paragraph, wherein the actuator is a rotary actuator operable to rotate the shaft about an axis of the shaft.

[0355] The gas turbine engine of any preceding paragraph, wherein the actuator is a linear actuator operable to move the shaft along an axis of the shaft.

[0356] The gas turbine engine of any preceding paragraph, wherein the valve assembly further comprises a shaft, a first pull rod, and a second pull rod, wherein the actuator is coupled to the shaft, wherein the first pull rod and the second pull rod are rotatably coupled to an end of the shaft opposite the actuator at a first connection point, wherein the first pull rod is further coupled to the cold-side valve section, and wherein the second pull rod is further coupled to the hot-side valve section.

[0357] The gas turbine engine according to any preceding paragraph, wherein the valve assembly is operable with the cold side bleed assembly at a location upstream of the CCA heat exchanger and operable with the hot side bleed assembly at a location upstream of the CCA heat exchanger.

[0358] The gas turbine engine according to any preceding paragraph, wherein the valve assembly is operable with the cold side bleed assembly at a location upstream of the CCA heat exchanger and operable with the hot side bleed assembly at a location downstream of the CCA heat exchanger.

[0359] The gas turbine engine according to any preceding paragraph, wherein the turbine section includes a high pressure turbine, and wherein the hot component is the high pressure turbine.

[0360] The gas turbine engine according to any preceding paragraph, wherein the hot component is a rotor bore of a turbine of the turbine section, an airfoil of the turbine, a rotor bore of the high pressure compressor, an airfoil of the high pressure compressor, an oil sump within the turbine section, a turbine mid-frame, a turbine aft-frame, or a combination thereof.

[0361] The gas turbine engine according to any preceding paragraph, wherein the turbomachine defines an undercowl cavity, and wherein the CCA heat exchanger and the valve assembly are each located in the undercowl cavity of the turbomachine.

[0362] The gas turbine engine according to any preceding paragraph, further comprising a bifurcation coupled to the turbomachine, wherein the bifurcation defines an inner cavity, wherein the turbomachine defines an undercowl cavity, and wherein the CCA heat exchanger and the valve assembly are each located in the undercowl cavity of the turbomachine, the inner cavity of the bifurcation, or both.

[0363] The gas turbine engine according to any preceding paragraph, wherein the cold side bleed assembly further defines an outlet, wherein the outlet of the cold side bleed assembly is in fluid communication with the working gas flow path at a location downstream of the turbine section, in fluid communication with a core cowl vent of the turbomachine, in fluid communication with a passage above the turbomachine, or a combination thereof.

[0364] The gas turbine engine according to any preceding paragraph, wherein the cold location of the gas turbine engine is through a low pressure compressor, the working gas flow path between the low pressure compressor and a high pressure compressor, or both.

[0365] The gas turbine engine according to any preceding paragraph, wherein the gas turbine engine defines a bypass passage above the turbomachine, and wherein the cold location is the bypass passage.

[0366] The gas turbine engine of any preceding paragraph, wherein the turbine section comprises a high pressure turbine, and wherein the hot component is the high pressure turbine.

[0367] A method of operating a valve assembly of a cooled cooling air (CCA) system of a gas turbine engine, the method comprising: receiving data indicative of an operating condition of the gas turbine engine; and responsive to receiving the data indicative of the operating condition of the gas turbine engine, actuating a valve assembly of the CCA system to cooperatively regulate a first airflow through a cold side bleed assembly of the CCA system and a second airflow through a hot side bleed assembly of the CCA system.

[0368] The method of any preceding paragraph, wherein the data indicative of the operating condition is data indicative of a power level of the gas turbine engine.

[0369] The method of any preceding paragraph, wherein the valve assembly comprises a cold side valve section operable with the cold side bleed assembly, a hot side valve section operable with the hot side bleed assembly, and an actuator coupled to both the cold side valve section and the hot side valve section, and wherein actuating the valve assembly comprises moving a shaft of the actuator assembly with the actuator, the shaft coupled to both the cold side valve section and the hot side valve section.

[0370] The method of any preceding paragraph, further comprising: receiving data indicative of a degradation parameter of one or more aspects of the gas turbine engine; and responsive to receiving the data indicative of a degradation parameter, actuating the valve assembly.

[0371] A gas turbine engine comprising: a turbomachine comprising a compressor section, a combustion section defining a compressor discharge cavity, and a turbine section comprising a turbine, the turbomachine further comprising: an active clearance control (ACC) system comprising an inlet, a heat transfer assembly disposed about the turbine of the turbine section, and an ACC duct assembly extending from the inlet to the heat transfer assembly; and a cooled cooling air (CCA) system comprising: a CCA heat exchanger; a hot side bleed assembly defining an inlet in fluid communication with the working gas flow path through the compressor section, at the compressor discharge cavity, or both, the hot side bleed assembly further in thermal communication with a hot component of the turbomachine to cool the hot component of the turbomachine; and a cold side bleed assembly in thermal communication with an airflow through the hot side bleed assembly across the CCA heat exchanger, the cold side bleed assembly defining an inlet in fluid communication with the ACC duct assembly.

[0372] The gas turbine engine of any preceding paragraph, wherein the turbine is a first turbine, and wherein the heat transfer assembly is a first heat transfer assembly, wherein the turbine section further comprises a second turbine, and wherein the ACC system further comprises a second heat transfer assembly disposed about the second turbine, wherein the ACC duct assembly extends to both the first heat transfer assembly and the second heat transfer assembly.

[0373] The gas turbine engine of any preceding paragraph, wherein the ACC duct assembly comprises a first portion extending to the first heat transfer assembly, a second portion extending to the second heat transfer assembly, and an upstream portion extending from the inlet of the ACC system to the first and second portions.

[0374] The gas turbine engine of any preceding paragraph, wherein the inlet of the cold side bleed assembly is in fluid communication with the upstream portion of the ACC duct assembly through a variable control valve of the CCA system.

[0375] The gas turbine engine of any preceding paragraph, wherein the ACC system comprises one or more flow control valves operable with the first portion of the ACC duct assembly, the second portion of the ACC duct assembly, or both.

[0376] The gas turbine engine of any preceding paragraph, wherein the ACC system comprises a blower in fluid communication with the upstream portion of the ACC duct assembly for pushing a flow of air through the ACC duct assembly.

[0377] The gas turbine engine of any preceding paragraph, wherein the inlet of the cold side bleed assembly is in fluid communication with the upstream portion of the ACC duct assembly at a location downstream of the blower.

[0378] The gas turbine engine of any preceding paragraph, wherein the CCA system comprises a blower in fluid communication with the cold side bleed assembly for pushing a flow of air through the cold side bleed assembly.

[0379] The gas turbine engine of any preceding paragraph, wherein the gas turbine engine defines a bypass passage above the turbomachinery, and wherein the inlet of the ACC duct assembly is in fluid communication with the bypass passage.

[0380] The gas turbine engine according to any preceding clause, wherein the ACC system includes a blower in fluid communication with the second portion of the ACC duct assembly for pushing a flow of air through the second portion of the ACC duct assembly.

[0381] The gas turbine engine according to any preceding clause, wherein the inlet of the cold side bleed assembly is in fluid communication with the second portion of the ACC duct assembly at a location downstream of the blower.

[0382] The gas turbine engine according to any preceding clause, wherein the cold side bleed assembly further defines an outlet, wherein the outlet is selectively in fluid communication with the first portion of the ACC duct assembly, the second portion of the ACC duct assembly, or both.

[0383] The gas turbine engine according to any preceding clause, wherein the cold side bleed assembly further defines an outlet, wherein the outlet is selectively in fluid communication with the ACC duct assembly at a location upstream of the heat transfer assembly.

[0384] The gas turbine engine according to any preceding clause, wherein the inlet of the ACC system is in fluid communication with a cold location of the gas turbine engine, wherein the cold location of the gas turbine engine is through a low pressure compressor, the working gas flow path between the low pressure compressor and a high pressure compressor, or both.

[0385] The gas turbine engine according to any preceding clause, wherein the turbine includes an operability bleed assembly defining an operability bleed duct extending between an inlet and an outlet, the inlet being in fluid communication with the working gas flow path at a location between the low pressure compressor and the high pressure compressor, wherein the inlet of the ACC system is in fluid communication with the operability bleed duct.

[0386] The gas turbine engine according to any preceding clause, wherein the inlet of the ACC system is in fluid communication with a cold location of the gas turbine engine, wherein the gas turbine engine further includes a bifurcation connected to the turbine, wherein the cold location of the gas turbine engine is a flow path surface of the bifurcation.

[0387] The gas turbine engine according to any preceding clause, wherein the turbine section includes a high pressure turbine, and wherein the hot component is the high pressure turbine.

[0388] The gas turbine engine of any preceding paragraph, wherein the hot component is a rotor bore of a turbine of the turbine section, an airfoil of the turbine, a rotor bore of a high pressure compressor, an airfoil of the high pressure compressor, an oil sump within the turbine section, a turbine mid-frame, a turbine aft-frame, or a combination thereof.

[0389] The gas turbine engine of any preceding paragraph, wherein the turbine defines a shroud down cavity, and wherein the CCA heat exchanger is located in the shroud down cavity of the turbine.

[0390] The gas turbine engine of any preceding paragraph, wherein the inlet of the hot side bleed assembly is in fluid communication with the working gas flow path at the compressor discharge cavity.

[0391] A gas turbine engine comprising: a turbine including a compressor section, a combustion section defining a compressor discharge cavity, and a turbine section, the compressor section, the combustion section, and the turbine section collectively partially defining a working gas flow path, the turbine further comprising: a reverse bleed system including a reverse bleed duct and an RBS blower in fluid communication with the reverse bleed duct, the reverse bleed duct being in fluid communication with the working gas flow path; and a cooled cooling air (CCA) system, the CCA system including: a CCA heat exchanger; a hot side bleed assembly defining an inlet in fluid communication with the working gas flow path through the compressor section, the compressor discharge cavity, or both, the hot side bleed assembly further being in thermal communication with a hot component of the turbine to cool the hot component of the turbine; and a cold side bleed assembly in thermal communication with a flow of gas through the hot side bleed assembly across the CCA heat exchanger, the cold side bleed assembly defining an inlet in fluid communication with the reverse bleed duct.

[0392] The gas turbine engine of any preceding paragraph, wherein the CCA system further includes a flow control valve, wherein the inlet of the cold side bleed assembly is in fluid communication with the reverse bleed duct through the flow control valve.

[0393] The gas turbine engine of any preceding paragraph, wherein the flow control valve is movable between a first position in which a flow of gas through the reverse bleed duct is directed into the working gas flow path and a second position in which the flow of gas through the reverse bleed duct is directed into the cold side bleed assembly.

[0394] A gas turbine engine as claimed in any preceding clause, wherein the flow control valve is movable to one or more intermediate positions to divert the flow of gas through the reverse bleed air duct to the working gas flow path and the cold side bleed air assembly.

[0395] A gas turbine engine as claimed in any preceding clause, wherein the flow control valve is movable to a closed position in which gas flow through the reverse bleed duct is prevented from reaching the working gas flow path or the cold side bleed assembly.

[0396] The gas turbine engine of any preceding clause, wherein the blower of the reverse bleed air assembly is positioned upstream of the flow control valve within the reverse bleed air duct.

[0397] The gas turbine engine of any preceding clause, wherein the compressor section, the combustion section, or both define a compressor bleed air plenum, wherein the reverse bleed air conduit is in fluid communication with the working gas flow path through the compressor bleed air plenum.

[0398] The gas turbine engine of any preceding clause, wherein the reverse bleed air duct is further in fluid communication with a cold location of the gas turbine engine, the cold location acting as a source of airflow for the reverse bleed air system.

[0399] The gas turbine engine of any preceding clause, further comprising a fork connected to the turbine, wherein the cold location of the gas turbine engine is a flow path surface of the fork.

[0400] A gas turbine engine according to any preceding clause, wherein the gas turbine engine defines a bypass passage above the turbine, and wherein the cold location is the bypass passage, the working gas flow path through the low-pressure compressor, the working gas flow path between the low-pressure compressor and the high-pressure compressor, or a combination thereof.

[0401] A gas turbine engine according to any preceding clause, wherein the turbine includes an operable bleed air assembly having an operable bleed air duct, the operable bleed air duct extending between an inlet and an outlet, the inlet being in fluid communication with the working gas flow path at a position between the low-pressure compressor and the high-pressure compressor, wherein the inlet of the reverse bleed air duct is in fluid communication with the operable bleed air duct.

[0402] The gas turbine engine of any preceding paragraph, wherein the outlet of the operable bleed conduit is in fluid communication with the bypass passage, wherein the cold location is the working gas flow path between the low pressure compressor and the high pressure compressor when the gas turbine engine is in an operating condition, and wherein the cold location is the bypass passage when the gas turbine engine is in a shutdown operating condition.

[0403] The gas turbine engine of any preceding paragraph, wherein the reverse bleed conduit comprises a first portion and a second portion, wherein the cold side bleed assembly comprises a first portion and a second portion, wherein the first portion of the cold side bleed assembly is integrated with the first portion of the reverse bleed conduit.

[0404] The gas turbine engine of any preceding paragraph, wherein the CCA heat exchanger is positioned within the first portion of the cold side bleed assembly.

[0405] The gas turbine engine of any preceding paragraph, wherein the RBS blower is positioned within the first portion of the reverse bleed conduit at a location downstream of the CCA heat exchanger.

[0406] A method of operating a gas turbine engine, the method comprising: receiving data indicative of an operating condition of the gas turbine engine; and responsive to receiving the data indicative of the operating condition of the gas turbine engine, actuating a flow control valve of a cooled cooling air (CCA) system to distribute a flow of air from a reverse bleed conduit of a reverse bleed assembly between a cold side bleed assembly of the CCA system and a working gas flow path of a turbine of the gas turbine engine.

[0407] The method of any preceding paragraph, wherein the operating condition is a shutdown operating condition, and wherein actuating the flow control valve of the CCA system comprises actuating the flow control valve of the CCA system to distribute the flow of air from the reverse bleed conduit to the working gas flow path of the turbine of the gas turbine engine.

[0408] The method of any preceding paragraph, wherein the operating condition is a flight operating condition, and wherein actuating the flow control valve of the CCA system comprises actuating the flow control valve of the CCA system to distribute the flow of air from the reverse bleed conduit to the cold side bleed assembly of the CCA system.

[0409] The method of any preceding clause, wherein the operating condition is a first operating condition, and wherein the method further comprises: powering a blower of the reverse bleed system with a first power source in response to receiving the data indicative of the gas turbine engine being in the first operating condition; receiving data indicative of a second operating condition of the gas turbine, the second operating condition being different than the first operating condition; and powering the blower of the reverse bleed system with a second power source separate from the first power source in response to receiving the data indicative of the second operating condition.

[0410] The method of any preceding clause, wherein the operating condition is a shutdown operating condition, and wherein actuating the flow control valve of the CCA system comprises: actuating the flow control valve of the CCA system for a first amount of time to distribute the gas flow from the reverse bleed duct to both the working gas flow path of the turbomachine and the cold side bleed assembly; and actuating the flow control valve of the CCA system for a second amount of time after the first amount of time to distribute the gas flow from the reverse bleed duct to only the working gas flow path of the turbomachine.

[0411] A gas turbine engine comprising: a turbomachine comprising a compressor section, a combustion section defining a compressor discharge plenum, and a turbine section, the compressor section, the combustion section, and the turbine section collectively partially defining a working gas flow path, the turbomachine further comprising: a reverse bleed system comprising a reverse bleed duct and an RBS blower in fluid communication with the reverse bleed duct, the reverse bleed duct being in fluid communication with the working gas flow path; and an active clearance control (ACC) system comprising an inlet, a heat transfer assembly arranged about the turbine of the turbine section, and an ACC duct assembly extending from the inlet to the heat transfer assembly, the inlet of the ACC system being in fluid communication with the reverse bleed duct.

[0412] The gas turbine engine of any preceding clause, wherein the reverse bleed system further comprises a flow control valve, wherein the inlet of the ACC system is in fluid communication with the reverse bleed duct through the flow control valve.

[0413] The gas turbine engine of any preceding clause, wherein the turbine is a first turbine, and wherein the heat transfer assembly is a first heat transfer assembly, wherein the turbine section further comprises a second turbine, and wherein the ACC system further comprises a second heat transfer assembly arranged about the second turbine, wherein the ACC duct assembly extends to both the first heat transfer assembly and the second heat transfer assembly.

[0414] The gas turbine engine of any preceding paragraph, wherein the ACC duct assembly includes a first portion extending to the first heat transfer assembly, a second portion extending to the second heat transfer assembly, and an upstream portion extending from the inlet of the ACC system to the first portion and the second portion.

[0415] The gas turbine engine of any preceding paragraph, wherein the ACC system includes one or more flow control valves operable with the first portion of the ACC duct assembly, the second portion of the ACC duct assembly, the upstream portion of the ACC duct assembly, or a combination thereof.

[0416] The gas turbine engine of any preceding paragraph, wherein the ACC duct assembly includes a first portion extending to the first heat transfer assembly and in fluid communication with the reverse bleed duct at a first location along the reverse bleed duct, wherein the ACC duct assembly further includes a second portion extending to the second heat transfer assembly and in fluid communication with the reverse bleed duct at a second location along the reverse bleed duct.

[0417] The gas turbine engine of any preceding paragraph, wherein the RBS blower is in fluid communication with the reverse bleed duct at a third location between the first location and the second location.

[0418] The gas turbine engine of any preceding paragraph, wherein the reverse bleed system includes a flow control valve, wherein the flow control valve is a variable three-way valve.

[0419] The gas turbine engine of any preceding paragraph, wherein the reverse bleed system further includes a flow control valve at the second location along the reverse bleed duct, wherein the second portion is in fluid communication with the reverse bleed duct through the flow control valve.

[0420] The gas turbine engine of any preceding paragraph, wherein the ACC system further includes a flow control valve downstream of the second location, and wherein the reverse bleed system further includes an RBS flow control valve downstream of the second location.

[0421] The gas turbine engine of any preceding paragraph, wherein the reverse bleed system includes a check valve at a location downstream of the inlet of the ACC system, wherein the check valve is positioned to prevent airflow from an outlet of the reverse bleed duct toward the blower.

[0422] The gas turbine engine according to any preceding clause, wherein the compressor section, the combustion section, or both define a compressor bleed plenum, wherein the reverse bleed duct is in fluid communication with the working gas flow path through the compressor bleed plenum.

[0423] The gas turbine engine according to any preceding clause, wherein the reverse bleed duct is further in fluid communication with a cold location of the gas turbine engine as a source of airflow for the reverse bleed system.

[0424] The gas turbine engine according to any preceding clause, further comprising: a bifurcation connected to the turbine, wherein the cold location of the gas turbine engine is a flow path surface of the bifurcation.

[0425] The gas turbine engine according to any preceding clause, wherein the gas turbine engine defines a bypass passage above the turbine, and wherein the cold location is the bypass passage, the working gas flow path through a low pressure compressor, the working gas flow path between the low pressure compressor and a high pressure compressor, or a combination thereof.

[0426] The gas turbine engine according to any preceding clause, wherein the turbine includes an operability bleed assembly defining an operability bleed duct extending between an inlet and an outlet, the inlet in fluid communication with the working gas flow path at a location between the low pressure compressor and the high pressure compressor, wherein the inlet of the reverse bleed system is in fluid communication with the operability bleed duct.

[0427] The gas turbine engine according to any preceding clause, wherein the outlet of the operability bleed duct is in fluid communication with the bypass passage, wherein the cold location is the working gas flow path between the low pressure compressor and the high pressure compressor when the gas turbine engine is in an operating condition, and wherein the cold location is the bypass passage when the gas turbine engine is in a shutdown operating condition.

[0428] The gas turbine engine according to any preceding clause, wherein the blower is a variable speed blower.

[0429] The gas turbine engine according to any preceding clause, further comprising: a fan section having a fan driven by the turbine, wherein the fan defines a fan pressure ratio less than or equal to 1.6 when the engine is operated in a cruise condition.

[0430] The gas turbine engine of any preceding paragraph, wherein the gas turbine engine is an open rotor gas turbine engine.

[0431] A gas turbine engine comprising: an accessory system; and a turbomachine comprising a compressor section, a combustion section defining a compressor discharge plenum, and a turbine section, the compressor section, the combustion section, and the turbine section collectively partially defining a working gas flow path, the turbomachine further comprising: a reverse bleed air system comprising a reverse bleed air duct and an RBS blower in fluid communication with the reverse bleed air duct, the reverse bleed air duct being in fluid communication with the working gas flow path; and an accessory cooling system comprising a cooling duct defining an inlet in fluid communication with the reverse bleed air duct, the accessory cooling duct comprising a cooling tip oriented toward the accessory system to provide a flow of air onto the accessory system.

[0432] The gas turbine engine of any preceding paragraph, wherein the accessory cooling system further comprises a flow control valve, wherein the inlet of the accessory cooling system is in fluid communication with the reverse bleed air duct through the flow control valve.

[0433] The gas turbine engine of any preceding paragraph, wherein the flow control valve is a variable three-way valve.

[0434] The gas turbine engine of any preceding paragraph, wherein the gas turbine engine further comprises a plurality of accessory systems, wherein the accessory system is a first accessory system of the plurality of accessory systems, wherein the accessory cooling system further comprises a plurality of cooling tips, wherein the cooling tip is a first cooling tip of the plurality of cooling tips, wherein a cooling duct comprises a plurality of branches, wherein each of the plurality of branches defines a respective one of the plurality of cooling tips, and wherein each of the plurality of cooling tips is oriented toward a respective accessory system of the plurality of accessory systems to provide a flow of air onto the respective accessory system.

[0435] The gas turbine engine of any preceding paragraph, wherein the accessory system is a first accessory system, wherein the gas turbine engine further comprises a second accessory system, wherein the cooling duct comprises a first portion extending toward the first accessory system and in fluid communication with the reverse bleed air duct at a first location along the reverse bleed air duct, wherein the cooling duct further comprises a second portion extending toward the second accessory system and in fluid communication with the reverse bleed air duct at a second location along the reverse bleed air duct.

[0436] The gas turbine engine according to any preceding paragraph, wherein the accessory cooling system further comprises a flow control valve at the second location along the reverse bleed duct, wherein the second portion is in fluid communication with the reverse bleed duct through the flow control valve.

[0437] The gas turbine engine according to any preceding paragraph, wherein the accessory cooling system further comprises a flow control valve downstream of the second location along the reverse bleed duct, wherein the second portion is in fluid communication with the reverse bleed duct through the flow control valve, and wherein the reverse bleed system further comprises an RBS flow control valve downstream of the second location.

[0438] The gas turbine engine according to any preceding paragraph, wherein the RBS blower is in fluid communication with the reverse bleed duct at a location upstream of the inlet of the cooling duct.

[0439] The gas turbine engine according to any preceding paragraph, wherein the reverse bleed system comprises a check valve at a location downstream of the inlet of the accessory cooling system, wherein the check valve is positioned to prevent gas flow from an outlet of the reverse bleed duct toward the blower.

[0440] The gas turbine engine according to any preceding paragraph, wherein the compressor section, the combustion section, or both define a compressor bleed plenum, wherein the reverse bleed duct is in fluid communication with the working gas flow path through the compressor bleed plenum.

[0441] The gas turbine engine according to any preceding paragraph, wherein the reverse bleed duct is further in fluid communication with a cold location of the gas turbine engine as a source of gas flow for the reverse bleed system.

[0442] The gas turbine engine according to any preceding paragraph, further comprising a bifurcation connected to the turbine, wherein the cold location of the gas turbine engine is a flow path surface of the bifurcation.

[0443] The gas turbine engine according to any preceding paragraph, wherein the gas turbine engine defines a bypass passage above the turbine, and wherein the cold location is the bypass passage, the working gas flow path through a low pressure compressor, the working gas flow path between the low pressure compressor and a high pressure compressor, or a combination thereof.

[0444] The gas turbine engine of any preceding paragraph, wherein the turbine includes an operability bleed assembly defining an operability bleed conduit extending between an inlet and an outlet, the inlet in fluid communication with the working gas flow path at a location between the low pressure compressor and the high pressure compressor, wherein the inlet of the reverse bleed system is in fluid communication with the operability bleed conduit.

[0445] The gas turbine engine of any preceding paragraph, wherein the outlet of the operability bleed conduit is in fluid communication with the bypass passage, wherein the cold location is the working gas flow path between the low pressure compressor and the high pressure compressor when the gas turbine engine is in an operating condition, and wherein the cold location is the bypass passage when the gas turbine engine is in a shutdown operating condition.

[0446] The gas turbine engine of any preceding paragraph, wherein the gas turbine engine defines a total pressure ratio greater than or equal to 50: 1 and less than or equal to 70: 1 when operating at rated speed during standard day operating conditions.

[0447] The gas turbine engine of any preceding paragraph, further comprising: a fan section having a fan driven by the turbine, wherein the fan defines a fan pressure ratio less than or equal to 1.6 when the engine is operating under cruise conditions.

[0448] The gas turbine engine of any preceding paragraph, wherein the gas turbine engine is an open rotor gas turbine engine.

[0449] The gas turbine engine of any preceding paragraph, wherein the gas turbine engine is a turbofan engine.

[0450] The gas turbine engine of any preceding paragraph, wherein the turbine defines a core mass flow entering the compressor section during operating conditions of the gas turbine engine, wherein the CCA system is configured to receive a bleed gas flow greater than or equal to 0.3% of the core mass flow and less than or equal to 13.5% of the core mass flow during the operating conditions of the gas turbine engine.

[0451] This written description uses examples to disclose the disclosure, including the best mode, and also to enable any person skilled in the art to practice the disclosure, including making and using any devices or systems and performing any incorporated methods. The patent scope of the disclosure is defined by the claims, and can include other examples that occur to those skilled in the art. Such other examples are intended to fall 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 languages of the claims.

Claims

1. A gas turbine engine characterized by, including: a turbine including a compressor section, a combustion section defining a compressor discharge plenum, and a turbine section, the compressor section, the combustion section, and the turbine section collectively partially defining a working gas flow path, the turbine further including: a reverse bleed air system including a reverse bleed air duct and a reverse bleed air system (RBS) blower in fluid communication with the reverse bleed air duct, the reverse bleed air duct in fluid communication with the working gas flow path; and an active clearance control (ACC) system including an inlet, a heat transfer assembly arranged about the turbine of the turbine section, and an ACC duct assembly extending from the inlet to the heat transfer assembly, the inlet of the ACC system in fluid communication with the reverse bleed air duct.

2. The gas turbine engine of claim 1, wherein, wherein, the reverse bleed air system further includes a flow control valve, wherein the inlet of the ACC system is in fluid communication with the reverse bleed air duct through the flow control valve.

3. The gas turbine engine of claim 1, wherein, wherein, the turbine is a first turbine, and wherein the heat transfer assembly is a first heat transfer assembly, wherein the turbine section further includes a second turbine, and wherein the ACC system further includes a second heat transfer assembly arranged about the second turbine, wherein the ACC duct assembly extends to both the first heat transfer assembly and the second heat transfer assembly.

4. The gas turbine engine of claim 3, wherein, wherein, the ACC duct assembly includes a first portion extending to the first heat transfer assembly, a second portion extending to the second heat transfer assembly, and an upstream portion extending from the inlet of the ACC system to the first portion and the second portion.

5. The gas turbine engine of claim 4, wherein, wherein, the ACC system includes one or more flow control valves operable with the first portion of the ACC duct assembly, the second portion of the ACC duct assembly, the upstream portion of the ACC duct assembly, or a combination thereof.

6. The gas turbine engine of claim 3, wherein, wherein, the ACC duct assembly includes a first portion extending to the first heat transfer assembly and in fluid communication with the reverse bleed air duct at a first location along the reverse bleed air duct, wherein the ACC duct assembly further includes a second portion extending to the second heat transfer assembly and in fluid communication with the reverse bleed air duct at a second location along the reverse bleed air duct.

7. The gas turbine engine of claim 6, wherein, wherein, the RBS blower is in fluid communication with the reverse bleed air duct at a third location between the first location and the second location.

8. The gas turbine engine of claim 6, wherein, wherein, the reverse bleed air system includes a flow control valve, wherein the flow control valve is a variable three-way valve.

9. The gas turbine engine of claim 6, wherein, wherein, the reverse bleed air system further includes a flow control valve at the second location along the reverse bleed air duct, wherein the second portion is in fluid communication with the reverse bleed air duct through the flow control valve.

10. The gas turbine engine of claim 9, wherein, wherein, The ACC system further includes a second ACC system flow control valve downstream of the second location, and wherein the reverse bleed system further includes a RBS flow control valve downstream of the second location.

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