Gas turbine engine defining rotor cavity

Abstract

A gas turbine engine includes a compressor section including a compressor, a combustion section and a turbine section arranged in a series flow sequence and defining a working gas flow path, the compressor including a last compressor stage; a spool drivingly coupled to the compressor; a stator vane stage downstream of the last compressor stage; and a stator housing, the spool and the stator housing together defining a rotor cavity in fluid communication with the working gas flow path, the stator vane stage including a first stator vane defining a fluid passage, and a stator housing defining a gas chamber in fluid communication with the fluid passage in the first stator vane, and a supplemental gas flow passage in fluid communication with the gas chamber in the first stator vane. A supplemental airflow passage is in fluid communication with the plenum and the rotor cavity for providing airflow to the rotor cavity.

Description

[0001] Priority information

[0002] This application claims priority to Polish patent application No. P.445204 filed on June 14, 2023. Technical Field

[0003] The present disclosure relates to a gas turbine engine having a rotor cavity, such as a gas turbine engine having a rotor cavity and a compressor discharge pressure seal for sealing the rotor cavity. Background Art

[0004] A gas turbine engine typically includes a turbine and a rotor assembly. A gas turbine engine, such as a turbofan engine, can be used for aircraft propulsion. In the case of a turbofan engine, the rotor assembly can be configured as a fan assembly. The turbine includes a compressor section and a combustion section. The compressor section includes a high-pressure compressor driven by a high-pressure spool. The turbine also includes an exhaust nozzle stage at a downstream position of the high-pressure compressor to guide airflow into the combustion section of the turbine. The support structure of the exhaust nozzle can define a rotor cavity together with the high-pressure spool. The turbine can include a compressor exhaust pressure seal, which can be between the support structure of the exhaust nozzle and the high-pressure spool, at a position to seal the rotor cavity. BRIEF DESCRIPTION OF THE DRAWINGS

[0005] A full 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 with reference to the accompanying drawings, in which:

[0006] Figure 1 is a cross-sectional view of a gas turbine engine according to an exemplary aspect of the present disclosure.

[0007] Figure 2 is a schematic cross-sectional view of a portion of a compressor section and a combustion section of an exemplary gas turbine engine.

[0008] Figure 3 yes Figure 2 A close-up schematic diagram of the discharge nozzle stage and final compressor stage of a compressor of an exemplary embodiment of the present invention.

[0009] Figure 4 yes Figure 3 A cross-section of the stator housing is viewed from front to back, as shown in FIG. Figure 3 As shown in line 4-4.

[0010] Figure 5 Provides includes and falls into Figure 6 Table of corresponding values ​​for several gas turbine engines within the ranges shown.

[0011] Figure 6Provides the SATR value on the x-axis and the swirl angle θ on the y-axis Swirl .

[0012] Figure 7 is a close-up schematic diagram of a stator vane stage and a final compressor stage of a HP compressor of a gas turbine engine configured in accordance with another exemplary embodiment of the present disclosure.

[0013] Figure 8 is a close-up schematic diagram of a stator vane stage and a final compressor stage of a HP compressor of a gas turbine engine configured in accordance with yet another exemplary embodiment of the present disclosure.

[0014] Fig. 9 is a schematic cross-sectional view of a portion of a compressor section and a combustion section of a gas turbine engine according to an exemplary aspect of the present disclosure.

[0015] Fig.10 is a close-up schematic illustration of a seal tooth assembly according to an exemplary aspect of the present disclosure.

[0016] Fig.11 yes Fig.10 Along Fig.10 Schematic cross-sectional view of the seal tooth assembly along line 11-11.

[0017] Fig.12 is a close-up schematic illustration of a seal tooth assembly according to another exemplary aspect of the present disclosure.

[0018] Fig.13 is a perspective view of a piston ring according to an exemplary aspect of the present disclosure.

[0019] Fig.14 is a schematic cross-sectional view of a portion of a compressor section and a combustion section of a gas turbine engine according to an exemplary aspect of the present disclosure. DETAILED DESCRIPTION

[0020] Reference will now be made in detail to the present embodiments of the present disclosure, one or more examples of which are shown in the accompanying drawings. The detailed description uses numerical and letter references to refer to features in the drawings. Similar or similar reference numbers in the drawings and description have been used to refer to similar or similar parts of the present disclosure.

[0021] The word "exemplary" as used herein means "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. In addition, unless expressly stated otherwise, all embodiments described herein should be considered exemplary.

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

[0023] In a context such as “at least one of A, B, and C,” the term “at least one of” means only A, only B, only C, or any combination of A, B, and C.

[0024] The term “turbomachine” refers to a machine that includes one or more compressors, a heat generating section (eg, a combustion section), and one or more turbines that together produce a torque output.

[0025] The term "gas turbine engine" refers to an engine having a turbine as all or part of its power source. Example gas turbine engines include turbofan engines, turboprop engines, turbojets, turboshaft engines, etc., as well as hybrid-electric versions of one or more of these engines.

[0026] The term "combustion section" refers to any heat addition system for a turbomachine. For example, the term combustion section may refer to a section that includes one or more of a deflagration combustion assembly, a rotating detonation combustion assembly, a pulse detonation combustion assembly, or other suitable heat addition assemblies. In certain example embodiments, the combustion section may include an annular combustor, a can combustor, a can-annular combustor, a trapped vortex combustor (TVC), or other suitable combustion systems, or combinations thereof.

[0027] The terms "low" and "high" or their respective comparatives (e.g., more—, if applicable), when used with a compressor, turbine, shaft or spool component, etc., each refer to a relative speed within the engine unless otherwise specified. For example, a "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.

[0028] The terms "upstream" and "downstream" refer to relative directions with respect to the flow of a fluid in a fluid path. For example, "upstream" refers to the direction from which the fluid is flowing, and "downstream" refers to the direction to which the fluid is flowing.

[0029] In general, a gas turbine engine includes a turbine, which in turn includes a compressor section, a combustion section, a turbine section, and an exhaust section, and defines a working gas flow path passing therethrough. More specifically, the compressor section includes a high-pressure compressor, and the turbine includes an exhaust nozzle stage located at a downstream position of the high-pressure compressor, fluidly connecting the high-pressure compressor to the combustion section. The high-pressure spool is rotatably coupled to the high-pressure compressor, and the stator housing is coupled to the exhaust nozzle stage inside the exhaust nozzle stage in a radial direction of the gas turbine engine. The stator housing and the high-pressure spool together define a rotor cavity that is fluidly coupled to the working gas flow path. The stator housing and the high-pressure spool also together form a compressor exhaust pressure seal at one end of the rotor cavity. The high-pressure spool includes a cantilever member as a part of the compressor exhaust pressure seal.

[0030] The inventors of the present disclosure have discovered that during certain operating conditions of a gas turbine engine, the cantilever member may be susceptible to undesirable oscillations due to the fact that the mechanical resonant frequency of the cantilever member and the acoustic resonant frequency of the cantilever member overlap or define a relatively small margin. The cantilever member may oscillate at its mechanical resonant frequency due to its connection to the rotating parts of the engine. The cantilever member may oscillate at its acoustic frequency due to its exposure to the working gas flow path passing through the rotor cavity. Specifically, the airflow through the working gas flow path immediately downstream of the rotor blade stage of the high-pressure compressor generally flows in a pulsed manner due to passing through the high-pressure compressor rotor blades during the operating conditions of the gas turbine engine. The pulsation of the airflow may cause the cantilever member to oscillate at its acoustic resonant frequency. When these mechanical and acoustic resonant frequencies overlap or define a relatively small margin, the cantilever member may wear or fail prematurely due to severe oscillations.

[0031] Therefore, the inventors of the present disclosure sought a means to increase the margin between the acoustic resonant frequency and the mechanical resonant frequency of a cantilever component during various operating conditions of a gas turbine engine. In the process of designing such an engine, the inventors unexpectedly discovered that by modifying the manner in which a supplemental airflow (e.g., an airflow in addition to the airflow directly from the working gas flow path) is provided to the rotor cavity, the airflow can affect the acoustic resonant frequency of the cantilever component, thereby increasing the margin between the acoustic resonant frequency and the mechanical resonant frequency of the cantilever component. More specifically, the inventors discovered in the process of designing several gas turbine engines that provide airflow to the rotor cavity in various different ways that there is a relationship between the temperature of the airflow provided to the rotor cavity (relative to the compressor outlet temperature) and the circumferential angle or swirl angle of the airflow provided to the rotor cavity, so that a system for providing airflow to the rotor cavity including one aspect of the present disclosure can bring a net benefit to the entire gas turbine engine design, and in particular can increase the margin between the acoustic resonant frequency and the mechanical resonant frequency of the cantilever component.

[0032] In order to achieve an improved gas turbine engine capable of increasing the margin between the acoustic resonant frequency and the mechanical resonant frequency that can act on a cantilevered component, the inventors proceeded during the design of several different types of systems for introducing airflow into the rotor cavity in the following manner: designing a gas turbine engine having a system for providing airflow to the rotor cavity at different temperatures and different swirl angles; checking the operability and resonant frequency characteristics of the designed gas turbine engine; redesigning the gas turbine engine to change the noted parameters based on the impact on other aspects of the gas turbine engine; rechecking the operability and resonant frequency characteristics of the redesigned gas turbine engine; and so on, including the system described herein, which is described in more detail below.

[0033] Referring now to the drawings, in which like reference numerals refer to like elements throughout, Figure 1 is a schematic cross-sectional view of a gas turbine engine 10 according to an exemplary embodiment of the present disclosure. More specifically, Figure 1 In an embodiment, the gas turbine engine is a high bypass turbofan jet engine. Figure 1 As shown, the gas turbine engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 for reference) and a radial direction R. Generally speaking, the gas turbine engine 10 includes a fan section 14 and a turbine 16 disposed downstream of the fan section 14 .

[0034] The illustrated exemplary turbine 16 generally includes a substantially tubular outer casing 18 defining an annular inlet 20. The outer casing 18 encloses in series flow relationship: a compressor section including a supercharger 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 or spool 34 drivingly connects the HP turbine 28 to the HP compressor 24. A low pressure (LP) shaft or spool 36 drivingly connects the LP turbine 30 to the LP compressor 22. In addition, the compressor section, the combustion section 26, and the turbine section together at least partially define a working gas flow path 37 extending therethrough.

[0035] For the depicted embodiment, the fan section 14 includes a variable pitch fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced apart manner. As shown, the fan blades 40 extend outwardly from the disk 42 generally in a radial direction R. Each fan blade 40 can rotate relative to the disk 42 about a pitch axis P by virtue of the fan blades 40 being operably coupled to a suitable actuation member 44, which is configured to collectively and uniformly change the pitch of the fan blades 40. The fan blades 40, the disk 42, and the actuation member 44 are capable of rotating together about the longitudinal centerline 12 through the LP spool 36 across a power gearbox 46. The power gearbox 46 includes a plurality of gears for reducing the rotational speed of the LP spool 36 to a more efficient fan rotational speed.

[0036] Still reference Figure 1In the exemplary embodiment of the present invention, the disk 42 is covered by a rotatable forward nacelle 48 that is aerodynamically formed to facilitate airflow through the plurality of fan blades 40. In addition, the exemplary fan section 14 includes an annular fan case or outer nacelle 50 that circumferentially surrounds the fan 38 and / or at least a portion of the turbine 16. For the depicted embodiment, the nacelle 50 is supported relative to the turbine 16 by a plurality of circumferentially spaced outlet guide vanes 52. In addition, a downstream section 54 of the nacelle 50 extends over an outer portion of the turbine 16 to define a bypass airflow passage 56 therebetween.

[0037] During operation of the gas turbine engine 10, a volume of air 58 enters the gas turbine engine 10 through the nacelle 50 and / or the associated inlet 60 of the fan section 14. As the volume of air 58 passes through the fan blades 40, a first portion of the air 58, as indicated by arrow 62, is directed or delivered into the bypass airflow passage 56, and a second portion of the air 58, as indicated by arrow 64, is directed or delivered into the LP compressor 22. The ratio between the first portion of air 62 and the second portion of air 64 is generally referred to as the bypass ratio. The second portion of air 64 then increases in pressure as it is delivered through the high pressure (HP) compressor 24 and into the combustion section 26, where it is mixed with fuel and combusted to provide combustion gases 66.

[0038] The combustion gases 66 are routed through the HP turbine 28, where a portion of the heat and / or kinetic energy from the combustion gases 66 is extracted via successive stages of HP turbine stator blades 68 coupled to the outer casing 18 and turbine rotor blades 70 coupled to the HP shaft or spool 34, thereby causing the HP shaft or spool 34 to rotate, thereby supporting operation of the HP compressor 24. The combustion gases 66 are then routed through the LP turbine 30, where a second portion of the heat and kinetic energy from the combustion gases 66 is extracted via successive stages of LP turbine stator blades 72 coupled to the outer casing 18 and LP turbine rotor blades 74 coupled to the LP shaft or spool 36, thereby causing the LP shaft or spool 36 to rotate, thereby supporting operation of the LP compressor 22 and / or rotation of the fan 38.

[0039] The combustion gases 66 are then routed through the jet exhaust nozzle section 32 of the turbine 16 to provide propulsive thrust. At the same time, the pressure of the first portion of air 62 is significantly increased when it is routed through the bypass airflow passage 56 before being discharged from the fan nozzle exhaust section 76 of the gas turbine engine 10, which also provides propulsive thrust. The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section 32 at least partially define a hot gas path 78 for routing the combustion gases 66 through the turbine 16.

[0040] However, it should be understood that Figure 1The exemplary gas turbine engine 10 depicted in FIG. 1 is provided as an example only, and in other exemplary embodiments, the gas turbine engine 10 may have any other suitable configuration. For example, in other exemplary embodiments, the gas turbine engine 10 may be a direct drive turbofan engine (i.e., not including a power gearbox 46), may include a fixed pitch fan 38, etc. Additionally or alternatively, aspects of the present disclosure may be incorporated into any other suitable gas turbine engine, such as a turboshaft engine, a turboprop engine, a turbojet engine, a land-based gas turbine engine for power generation, an aeroderivative gas turbine engine, etc.

[0041] See now Figure 2 , provides Figure 1 FIG. 1 is a schematic cross-sectional view of a portion of a compressor section and a combustion section 26 of an exemplary gas turbine engine 10. More specifically, in addition to the combustion section 26, Figure 2 Also shown is the rear end of the HP compressor 24 of the compressor section.

[0042] As described above, during operation of the gas turbine engine 10, the airflow through the working gas flow path 37 of the gas turbine engine 10 is compressed sequentially as it flows through the compressor section (or more specifically, as it flows through the LP compressor 22 and the HP compressor 24). The compressed air from the compressor section is then provided to the combustion section 26, where at least a portion of the compressed air is mixed with fuel and combusted to produce combustion gases 66. The combustion gases 66 flow from the combustion section 26 to the turbine section, and more specifically, for the illustrated embodiment, sequentially through the HP turbine 28 and the LP turbine 30, driving the HP turbine 28 and the LP turbine 30 (see Figure 1 HP spool 34 is drivingly coupled to both HP turbine 28 and HP compressor 24 (see Figure 1 ).

[0043] See especially Figure 2 , the HP compressor 24 includes a plurality of compressor stages 80, each compressor stage 80 including, for example, a plurality of HP compressor rotor blades 82 and a rotor 84. Each of the respective compressor stages 80 is drivingly coupled to the HP spool 34 so that the HP turbine 28 can drive the HP compressor 24 through the HP spool 34. Among the plurality of compressor stages 80 of the HP compressor 24, the last compressor stage 80A is located at the rear end of the HP compressor 24. The compressor rotor blades 82 of the last compressor stage 80A are attached to the corresponding rotor 84 of the last compressor stage 80A.

[0044] The last compressor stage 80A provides compressed air to the combustion section 26. More specifically, in the illustrated embodiment, the gas turbine engine 10 includes a stator vane stage located downstream of the last compressor stage 80A, fluidly connecting the last compressor stage 80A of the HP compressor 24 to the combustion section 26. More specifically, for the depicted embodiment, the stator vane stage is configured to fluidly connect the last compressor stage 80A of the HP compressor 24 to a discharge nozzle 92 stage of the combustion section 26.

[0045] More specifically, the combustion section 26 defines a diffuser cavity 98, and the exhaust nozzle 92 stage is located downstream of the last compressor stage 80A of the HP compressor 24 and upstream of the diffuser cavity 98, so that compressed air from the last compressor stage 80A is provided to the diffuser cavity 98 through the exhaust nozzle 92 stage. The compressed air in the diffuser cavity 98 is in turn provided to the combustor assembly 96, where the compressed air is mixed with fuel and combusted to produce the combustion gases 66. As shown, the combustor assembly 96 generally includes a fuel nozzle 100, an inner liner 102, and an outer liner 104, which together form a combustion chamber 106. It should be understood that the combustor assembly 96 can be configured as any suitable combustor assembly 96 for the gas turbine engine 10. For example, in certain embodiments, the combustor assembly 96 can be configured as an annular combustor assembly, a can combustor assembly, or an annular can combustor assembly.

[0046] Still see Figure 2 As previously described, the HP spool 34 is drivingly connected to the HP compressor 24. For the illustrated embodiment, the HP spool 34 generally includes a center spool section and a forward spool section 109, the center spool section including a center spool member 108, the forward spool section 109 including a forward spool member 110. For the illustrated embodiment, the center spool member 108 extends generally along the axial direction A at a location radially inward of the combustor assembly 96 of the combustion section 26. In certain exemplary embodiments, the center spool member 108 may define a substantially cylindrical shape. The forward spool member 110 generally extends from the center spool member 108 to the last compressor stage 80A of the HP compressor 24 of the compressor section. More specifically, for the depicted embodiment, the HP spool 34 partially forms a compressor discharge pressure seal 112.

[0047] In addition, as shown in the figure, for Figure 2In an embodiment of the present invention, the engine includes a stator housing 94, which is coupled to the discharge nozzle 92 stage at a position located inside the discharge nozzle 92 stage along the radial direction R. The stator housing 94 also partially forms a compressor discharge pressure seal 112, so that the HP spool 34 and the stator housing 94 form the compressor discharge pressure seal 112. As shown, the stator housing 94 forms the stator portion of the compressor discharge pressure seal 112, and the HP spool 34 forms the rotor portion 116 of the compressor discharge pressure seal 112 (the rotor portion 116 can rotate relative to the stator portion). The stator portion typically includes a sealing gasket 118, and the rotor portion 116 typically includes a plurality of sealing teeth 120, and the sealing teeth 120 are configured to form a seal with the sealing gasket 118. As will be understood, the rotor portion 116 is generally cantilevered from the front spool member 110.

[0048] However, it should be appreciated that in other exemplary embodiments, the compressor discharge pressure seal 112 may have any other suitable configuration.

[0049] Still see Figure 2 It should be understood that the HP spool 34 and stator housing 94 together define a rotor cavity 122 in fluid communication with the working gas flow path 37. Specifically, the rotor cavity 122 extends between a rotor cavity opening 123 that fluidly connects the rotor cavity 122 with the working gas flow path 37 and the compressor discharge pressure seal 112.

[0050] In addition, from Figure 2 The gas turbine engine 10 also includes a system for providing airflow to the rotor cavity 122 to, for example, affect the acoustic resonant frequency of one or more components exposed to the rotor cavity 122. Specifically, the discharge nozzle 92 stage includes a first discharge nozzle 92 defining a fluid passage 124 therethrough. Figure 2 In the embodiment of the present invention, the fluid passage 124 of the first discharge nozzle 92 is in fluid communication with the compressor of the compressor section and is configured to receive the bleed air flow from the compressor section. More specifically, the fluid passage 124 of the first discharge nozzle 92 is configured to receive the bleed air flow from the HP compressor 24, and more specifically, to receive the bleed air flow from the last compressor stage 80A of the HP compressor 24. Specifically, the gas turbine engine 10 includes an air flow delivery system 126 having an air flow duct 128, which is configured to receive the bleed air flow and provide the bleed air flow to the first discharge nozzle 92.

[0051] As will be appreciated, during operating conditions of the gas turbine engine 10 (e.g., takeoff operating conditions, climb operating conditions, or cruise operating conditions), the gas flow exiting the HP compressor 24 through the working gas flow path 37 defines a compressor outlet temperature T3 in degrees Rankine. When the gas flow is provided to the rotor cavity 122, the gas flow provided by the gas flow delivery system 126 during the operating conditions may be at a passage temperature TP in degrees Rankine, as will be explained in more detail below. In at least some exemplary embodiments, the passage temperature T P It may be useful to affect the acoustic resonant frequency of the rotor section 116 of the compressor discharge pressure seal 112 in a desired manner less than the compressor outlet temperature T3. Specifically, for the illustrated embodiment, the gas turbine engine 10 includes a heat exchanger 130 in fluid communication with the fluid passage 124 of the first discharge nozzle 92 at a position downstream of the compressor section and upstream of the first discharge nozzle 92. More specifically, the heat exchanger 130 is thermally connected to the airflow duct 128 for reducing the temperature of the bleed air flow passing through the airflow duct. For the illustrated embodiment, the airflow duct 128 is configured to receive the bleed air flow from the diffuser cavity 98 and provide the bleed air flow to the fluid passage 124.

[0052] Although not shown, it should be understood that the heat exchanger 130 may be further thermally coupled to a heat sink, such as the bypass airflow passage 56 of the gas turbine engine 10 , a fuel flow to the combustion section 26 of the gas turbine engine 10 , etc. Additionally, a controller (not depicted) of the gas turbine engine 10 may be operably coupled to one or more components to control the heat rejected by the heat exchanger 130 .

[0053] Now also refer to Figure 3 , provided above reference Figure 2 A close up schematic diagram of the discharge nozzle 92 stage and the final compressor stage 80A of the HP compressor 24 is depicted.

[0054] As will be appreciated, for the illustrated embodiment, the stator housing 94 further defines a plenum 132 and a supplemental air flow passage 134. The plenum 132 is in fluid communication with the fluid passage 124 in the first discharge nozzle 92 (which, as described above, receives bleed air flow from the compressor section in the illustrated embodiment), and the supplemental air flow passage 134 is in fluid communication with the plenum 132 and the rotor cavity 122 for providing supplemental air flow from the plenum 132 and the fluid passage 124 to the rotor cavity 122. The supplemental air flow provided to the rotor cavity 122 may be used to purge the rotor cavity 122 and affect the acoustic resonant frequency of the rotor section 116 of the compressor discharge pressure seal 112.

[0055] In short, it should be understood that the stator housing 94 and the HP spool 34 further define the supplemental gas flow passage 134. The rotor cavity opening 123 is defined at a location having a minimum cross-sectional area between the rotor cavity 122 and the working gas flow path. The minimum cross-sectional area may be referred to as the cross-sectional area A of the rotor cavity opening 123. FO .

[0056] In addition, now refer to Figure 4 , provides a front-to-rear view of a cross section of the stator housing 94, as shown Figure 3 As shown in line 4-4.

[0057] As described above, the stator housing 94 defines the air chamber 132 and the supplemental air flow channel 134. More specifically, for the depicted embodiment, the air chamber 132 extends in the circumferential direction C of the gas turbine engine 10. In addition, the supplemental air flow channel 134 is a first supplemental air flow channel 134A in an array of supplemental air flow channels 134 defined by the stator housing 94. Each supplemental air flow channel 134 in the array of supplemental air flow channels 134 is fluidly coupled to the air chamber 132 and the rotor cavity 122 for providing supplemental air flow to the rotor cavity 122. The array of supplemental air flow channels 134 is arranged along the circumferential direction C of the gas turbine engine 10, so that the array of supplemental air flow channels 134 can provide air flow to the rotor cavity 122 in a distributed manner along the circumferential direction C.

[0058] As briefly mentioned above, the swirl angle θ that provides supplemental airflow to the rotor cavity 122 Swirl The acoustic resonant frequency of the rotor section 116 of the compressor discharge pressure seal 112 may be further influenced. Figure 4 In the embodiment, the array of supplementary air flow channels 134 is arranged to have a swirl angle θ greater than zero degrees. Swirl Supplementary airflow is provided to the rotor cavity 122. For example, the gas turbine engine 10 is defined in the axial direction A (see Figure 3 ;Enter Figure 4 The reference plane 136 extending in the radial direction R and the first supplementary airflow channel 134A passes through the reference plane 136. The first supplementary airflow channel 134A includes an airflow outlet 138 defining an airflow outlet direction 140. The airflow outlet direction 140 defines a swirl angle θ greater than 0 degrees with the reference plane 136. Swirl Specifically, for the embodiment shown, the swirl angle θ with respect to the reference plane 136 is Swirl is greater than 0 degrees and less than or equal to 85 degrees. For example, the swirl angle θ Swirl It may be greater than or equal to 15 degrees and less than or equal to 75 degrees.

[0059] It is worth noting that it should be understood that each of the plurality of supplemental airflow channels 134 is configured in a similar manner as the first supplemental airflow channel 134A. Specifically, each of the plurality of supplemental airflow channels 134 similarly defines an airflow outlet direction 140 that defines a swirl angle θ with a local reference plane (see, for example, local reference plane 136'). Swirl Each local reference plane passes through a corresponding supplementary airflow channel 134 in the plurality of supplementary airflow channels 134 and extends in the axial direction A (see Figure 3 ) and extends in the radial direction R. Each swirl angle θ Swirl The swirl angle θ defined by the airflow outlet direction 140 of the airflow outlet 138 of the first supplemental airflow channel 134A may be Swirl Same as above.

[0060] In short, it should be understood that the airflow outlets 138 of all of the supplemental airflow channels 134 may collectively define a total cross-sectional area A SAP_Total The cross-sectional area A of the rotor chamber opening 123 is FO The cross-sectional area A of the multiple supplementary air flow channels SAP_Total The ratio (A FO Ratio A SAP_Total ) may affect the ability of the supplemental airflow channel to achieve the desired changes described herein.

[0061] It should be understood that although a single discharge nozzle 92 is shown including a fluid passage 124 (see, e.g., Figure 2 and Figure 3 ), but in other exemplary embodiments, the gas turbine engine 10 may include a plurality of discharge nozzles 92 arranged in the circumferential direction C, each discharge nozzle 92 including one or more corresponding fluid passages 124, each fluid passage 124 being fluidly coupled to the air chamber 132. Additionally, as an alternative, the stator housing 94 may include a plurality of air chambers 132 arranged in the circumferential direction C, each air chamber 132 being fluidly coupled to one or more fluid passages 124 extending through the corresponding one or more discharge nozzles 92.

[0062] As mentioned previously, the inventors, in the process of designing a gas turbine engine that provides airflow to a rotor cavity—for example, designing a gas turbine engine that provides airflow to a rotor cavity at various temperatures and various swirl angles, and evaluating the overall performance and resonant frequency characteristics of the engine—unexpectedly discovered that there is a significant relationship between the temperature of the airflow provided to the rotor cavity (relative to the compressor outlet temperature) and the circumferential angle or swirl angle at which the airflow is provided to the rotor cavity. This relationship is an indicator of the ability of the gas turbine engine to effectively increase the margin between the acoustic resonant frequency and the mechanical resonant frequency of the cantilevered component exposed to the rotor cavity.

[0063] The relationship is applicable to a gas turbine engine configured to provide supplemental gas flow to a rotor cavity extending between a working gas flow path of the gas turbine engine and a seal having a cantilevered member. The relationship relates a swirl angle provided by the supplemental gas flow to the rotor cavity to a temperature ratio of the gas flow provided to the rotor cavity to identify a system capable of achieving a desired increase in margin between acoustic and mechanical resonant frequencies of the cantilevered member, as described herein.

[0064] Specifically, the inventors of the present disclosure have discovered that by incorporating a circumferential vortex into a rotor cavity having a cantilevered member exposed thereto, it is possible to act to counteract the airflow pressure pulses through the rotor cavity imposed in part by the passage of the rotor blades of the final stage of the HP compressor. In this manner, the inventors have discovered that the effects of the airflow pressure pulses from the rotor blades are reduced, making the cantilevered member less susceptible to oscillations. In this manner, more vortex generally results in reduced effects and fewer oscillations.

[0065] However, the inventors have further discovered that utilizing too high a swirl angle may be counterproductive as it may result in a design that is difficult to manufacture and may lead to an inefficient air flow delivery path.

[0066] In addition, the inventors of the present disclosure have found that by reducing the temperature of the supplementary airflow provided to the rotor cavity, the acoustic dwell frequency of the cantilever member can be affected to increase the margin between the acoustic resonant frequency and the mechanical resonant frequency. Specifically, the inventors have found that the acoustic resonant frequency of the cantilever member is a function of temperature, so that reducing the temperature changes the acoustic resonant frequency.

[0067] However, the inventors have found that the benefits may be reduced by incorporating a temperature reduction above a certain threshold in the supplemental air flow provided to the rotor cavity due to the amount of energy required to achieve such a temperature reduction.

[0068] In short, the inventors have further discovered that in order to ensure that the supplementary airflow introduced into the rotor cavity for vortex and / or temperature change has the desired effect, the inventors have discovered that the size of the opening providing the supplementary airflow must be within a certain relative size range of the cross-sectional area of ​​the rotor cavity opening of the rotor cavity. By maintaining the ratio of the cross-sectional area of ​​the rotor cavity opening to the total cross-sectional area of ​​the plurality of openings within a certain range, the supplementary airflow provided to the rotor cavity has the desired effect described herein.

[0069] Thus, the relationships discovered below make it possible to identify a gas turbine engine that provides airflow to the rotor cavity, is capable of achieving the desired increase in the margin between the acoustic and mechanical resonant frequencies of the cantilevered component, while avoiding the sharp drop in efficiency of the gas turbine engine and the other disadvantages mentioned above, and is suitable for specific mission requirements, taking into account efficiency, weight, complexity, reliability and other factors that affect the optimal choice of a gas turbine engine that provides airflow to the rotor cavity.

[0070] In addition to producing an improved gas turbine engine that provides supplemental airflow to the rotor cavity, as explained in detail above, by utilizing this relationship, the inventors have discovered that the number of suitable or feasible gas turbine engine designs that can meet the above-mentioned design requirements can be greatly reduced, thereby facilitating a more rapid down-selection of designs to be considered when developing a gas turbine engine. Such a benefit can provide a deeper understanding of the requirements of a given gas turbine engine before the specific technical, integration and system requirements are fully developed. Such a benefit can also avoid late-stage redesigns.

[0071] A desirable relationship discovered by the inventors to provide an improved gas turbine engine is the relationship between the swirl angle of the supplemental air flow passage defined by the stator housing and the supplemental air flow temperature ratio (SATR), expressed as:

[0072] SATR=T P / T3,

[0073] Where T P is the passage temperature in degrees Rankine during the operating condition of the gas turbine engine, and T3 is the compressor outlet temperature in degrees Rankine during the operating condition of the gas turbine engine. More specifically, the inventors of the present disclosure have found that when the SATR is less than 1.0, such as less than or equal to 0.9, the swirl angle of the plurality of supplementary air flow passages should be greater than or equal to 0 degrees and less than or equal to 85 degrees, and when the SATR is greater than or equal to 1.0, the swirl angle of the plurality of supplementary air flow passages should be greater than 0 degrees and less than or equal to 85 degrees. This configuration may allow the gas turbine engine to achieve a desired increase in the margin between the acoustic resonant frequency and the mechanical resonant frequency of the cantilever component.

[0074] Now refer to Figure 5 and Figure 6 Various gas turbine engines according to one or more aspects of the present disclosure are provided, which can provide supplemental airflow to a rotor cavity through an array of supplemental airflow channels. Figure 5 Provides includes and falls into Figure 6 Table of corresponding values ​​for several gas turbine engines within the ranges shown. Figure 6 The SATR value on the x-axis 202 and the swirl angle θ on the y-axis 204 are provided. Swirl Figure 200.

[0075] like Figure 6 As shown, the graph 200 depicts a first range 206 in which the swirl angles θ of the plurality of supplemental airflow channels are greater than or equal to 0.8 and less than 1. Swirl is greater than or equal to 0 degrees and less than or equal to 85 degrees, and when the SATR value is greater than or equal to 1 and less than or equal to 1.15, the swirl angle θ of the plurality of supplementary air flow channelsSwirl Greater than 0 degrees and less than or equal to 85 degrees. In addition, Figure 6 A second range 208 is depicted, which is a subset of the first range 206, wherein the relationship between the swirl angle and the SATR is further provided by the following relationship:

[0076] θ ≥ 199 × SATR 2 -186×SATR+7.

[0077] The inventors of the present disclosure have discovered that a gas turbine engine having a plurality of supplemental airflow passages configured to provide airflow to the rotor cavity in this manner can effectively increase the margin between the acoustic resonant frequency and the mechanical resonant frequency of the cantilevered member.

[0078] The working gas flow path 37 of the rotor cavity has a cross-sectional area A defined by the opening FO , and the plurality of supplementary airflow channels define a total cross-sectional area A SAP_Total The inventors of the present invention have also found that the cross-sectional area A of the opening of the working gas flow path 37 is FO The cross-sectional area A of the multiple supplementary air flow channels SAP_Total The ratio (A FO Ratio A SAP_Total ) can further indicate whether the structure of the gas turbine engine can effectively increase the margin between the acoustic resonant frequency and the mechanical resonant frequency of the cantilever component.

[0079] Specifically, the inventors of the present invention found that when A FO With A SAP_Total When the ratio is greater than or equal to 0.05 and less than 0.25, the structure of the gas turbine engine configured according to one or more of the above exemplary aspects can effectively increase the margin between the acoustic resonance frequency and the mechanical resonance frequency that can act on the cantilever member. For example, the supplementary airflow can effectively affect the vortex of the airflow in the rotor cavity, can effectively affect the temperature of the airflow in the rotor cavity, or both.

[0080] Although the above discussion generally relates to a gas turbine engine having a thermal management system having a heat exchanger integrated into a stator vane stage (e.g., an exhaust nozzle), wherein the stator vanes each define an internal hot fluid passage that is in fluid communication with the compressor via an air flow duct, in other exemplary embodiments, other configurations may also be provided.

[0081] For example, see now Figure 7 , provides a close-up schematic diagram of a stator vane stage, and more specifically, a close-up schematic diagram of a discharge nozzle 92 stage, and a final compressor stage 80A of an HP compressor 24 of a gas turbine engine 10 configured in accordance with another exemplary embodiment of the present disclosure. Figure 7 The exemplary embodiments can be described with reference to Figure 3 The exemplary embodiments described are configured in substantially the same manner. The same or similar numbers may refer to the same or similar parts.

[0082] For example, Figure 7 The exemplary embodiment of generally includes a discharge nozzle 92 stage located downstream of the last compressor stage 80A. The discharge nozzle 92 stage generally includes a first discharge nozzle 92 defining a fluid passage 124. The fluid passage 124 of the first discharge nozzle 92 is in fluid communication with the HP compressor 24, in direct fluid communication with the working gas flow path 37 through the discharge nozzle 92 stage, or both. More specifically, for the illustrated embodiment, the fluid passage 124 of the first discharge nozzle 92 is in direct fluid communication with the working gas flow path 37 through the discharge nozzle 92 stage. Specifically, the first discharge nozzle 92 defines an opening 144 in an outer surface 146 at a position within the working gas flow path 37 for directly receiving a gas flow from the working gas flow path 37. More specifically, in the depicted embodiment, the opening 144 is located at the leading edge. The fluid passage 124 is in fluid communication with the opening 144 to receive a gas flow.

[0083] Although opening 144 is depicted as being located at the leading edge of first discharge nozzle 92, in other embodiments, opening 144 may be located on the pressure side of discharge nozzle 92, on the suction side of first discharge nozzle 92, or in any other suitable location. Figure 7 A single opening 144 is shown in FIG. 1 , but in other embodiments, the first discharge nozzle 92 may define a plurality of openings 144 in fluid communication with the fluid passage 124 , with the plurality of openings 144 positioned at any suitable locations.

[0084] As will be appreciated, utilizing Figure 7 configuration, the SATR value (e.g., the channel temperature T in degrees Rankine during the operating conditions of the gas turbine engine) P The ratio of the compressor outlet temperature T3 in degrees Rankine during the operating conditions of the gas turbine engine may be equal to 1, or even greater than 1. Therefore, in order to achieve the benefits described above, the plurality of supplemental air flow passages 134 defined by the stator housing 94 may define a swirl angle θ greater than 0 degrees. Swirl .

[0085] In addition, now refer to Figure 8 , provides a close-up schematic illustration of the exhaust nozzle 92 stage of the gas turbine engine 10 and the final compressor stage 80A of the HP compressor 24 configured in accordance with yet another exemplary embodiment of the present disclosure. Figure 8 The exemplary embodiments can be described with reference to Figure 3 The exemplary embodiments described are configured in substantially the same manner. The same or similar numbers may refer to the same or similar parts.

[0086] For example, Figure 8 The exemplary embodiment of the invention generally includes a stator housing 94 coupled to a discharge nozzle stage 92. In the illustrated embodiment, the stator housing 94 includes an anti-eddy current material strip 148 exposed to the rotor cavity 122 to help control eddy currents within the rotor cavity 122. For the depicted embodiment, the anti-eddy current material strip 148 may be a corrugated material strip positioned on the front side of the stator housing 94.

[0087] Additionally, for the depicted embodiment, the stator housing 94 also includes a second strip of anti-eddy current material 148 positioned at a location downstream of the compressor discharge pressure seal 112 and a third strip of anti-eddy current material 148 located between adjacent teeth 120 of the rotor segment 116 of the compressor discharge pressure seal 112. The second and third strips of anti-eddy current material 148 may also help control eddy currents between the stator housing 94 and the HP spool 34.

[0088] In short, it should also be understood that Figure 8 In the exemplary embodiment, stator housing 94 does not define a dedicated plenum, but rather is configured to provide supplemental air flow from fluid passage 124 through discharge nozzle 92 directly to one or more supplemental air flow passages 134 defined in stator housing 94 .

[0089] Now refer to Fig. 9 , provides a schematic cross-sectional view of a portion of a compressor section and a combustion section 26 of a gas turbine engine 10 according to another exemplary embodiment of the present disclosure. In at least some exemplary embodiments, the gas turbine engine 10 may be configured to operate in a manner similar to that of FIG. Figure 1 The exemplary gas turbine engine 10 is configured in a similar manner. The same or similar numbers may refer to the same or similar parts.

[0090] like Fig. 9 As shown, the HP compressor 24 of the compressor section includes a plurality of compressor stages 80, each compressor stage 80 including, for example, a plurality of HP compressor rotor blades 82 and a rotor 84. Each of the respective compressor stages 80 is drivingly coupled to the HP spool 34 so that the HP turbine 28 (see Figure 1 ) may drive the HP compressor 24 through the HP spool 34. Among the plurality of compressor stages 80 of the HP compressor 24, the last compressor stage 80A is located at the rear end of the HP compressor 24. The compressor rotor blades 82 of the last compressor stage 80A are attached to the corresponding rotor 84 of the last compressor stage 80A.

[0091] The last compressor stage 80A provides compressed air to the combustion section 26. More specifically, in the illustrated embodiment, the gas turbine engine 10 includes a stator vane stage located downstream of the last compressor stage 80A, fluidly connecting the last compressor stage 80A of the HP compressor 24 to the combustion section 26. More specifically, for the depicted embodiment, the stator vane stage is configured to fluidly connect the last compressor stage 80A of the HP compressor 24 to a discharge nozzle 92 stage of the combustion section 26.

[0092] More specifically, the combustion section 26 defines a diffuser cavity 98, and the exhaust nozzle 92 stage is located downstream of the last compressor stage 80A of the HP compressor 24 and upstream of the diffuser cavity 98, so that compressed air from the last compressor stage 80A is provided to the diffuser cavity 98 through the exhaust nozzle 92 stage. The compressed air in the diffuser cavity 98 is in turn provided to the combustor assembly 96, where the compressed air is mixed with fuel and combusted to produce the combustion gases 66. As shown, the combustor assembly 96 generally includes a fuel nozzle 100, an inner liner 102, and an outer liner 104, which together form a combustion chamber 106. It should be understood that the combustor assembly 96 can be configured as any suitable combustor assembly 96 for the gas turbine engine 10. For example, in certain embodiments, the combustor assembly 96 can be configured as an annular combustor assembly, a can combustor assembly, or an annular can combustor assembly.

[0093] Still see Fig. 9 As previously described, the HP spool 34 is drivingly connected to the HP compressor 24. For the illustrated embodiment, the HP spool 34 generally includes a center spool section and a forward spool section 109, the center spool section including a center spool member 108, the forward spool section 109 including a forward spool member 110. For the depicted embodiment, the center spool member 108 extends generally in the axial direction A at a location along the radial direction R inside the combustor assembly 96 of the combustion section 26. In certain exemplary embodiments, the center spool member 108 defines a substantially cylindrical shape. The forward spool member 110 extends generally from the center spool member 108 to the last compressor stage 80A of the HP compressor 24 of the compressor section. More specifically, for the depicted embodiment, the HP spool 34 partially forms a compressor discharge pressure seal 112.

[0094] In addition, as shown in the figure, for Fig. 9In an embodiment of the present invention, the engine includes a stator housing 94, which is coupled to the discharge nozzle 92 stage at a position located in the inner stage of the discharge nozzle 92 along the radial direction R. The stator housing 94 also partially forms a compressor discharge pressure seal 112, so that the HP spool 34 and the stator housing 94 form the compressor discharge pressure seal 112. As shown, the stator housing 94 forms a stator portion of the compressor discharge pressure seal 112, and the HP spool 34 forms a rotor portion 116 of the compressor discharge pressure seal 112 (the rotor portion 116 can rotate relative to the stator portion). The stator portion generally includes a sealing gasket 118, and the rotor portion 116 generally includes a plurality of sealing teeth 120, and the sealing teeth 120 are configured to form a seal with the sealing gasket 118.

[0095] However, it should be appreciated that in other exemplary embodiments, the compressor discharge pressure seal 112 may have any other suitable configuration.

[0096] Still see Fig. 9 It should be appreciated that the HP spool 34 , and more specifically, the rotor portion 116 of the compressor discharge pressure seal 112 , includes a seal tooth assembly 150 .

[0097] See now Fig.10 , provides a close-up schematic illustration of a seal tooth assembly 150 according to an exemplary aspect of the present disclosure. Fig.10 The exemplary seal tooth assembly 150 may be similar to that described above with reference to Fig. 9 The exemplary seal tooth assembly 150 described is configured in substantially the same manner.

[0098] For the illustrated embodiment, the seal tooth assembly 150 includes the seal teeth 120, and more specifically, first and second seal teeth 120A, 120B, extending toward the seal pad 118 to prevent or minimize airflow therethrough. In addition, the seal tooth assembly 150 generally includes a seal support extension 152, and a damper 154, with the first and second seal teeth 120A, 120B extending from the seal support extension 152 toward the seal pad 118 of the stator housing 94, and the damper 154 is operable with the seal support extension 152 to inhibit movement in the radial direction R.

[0099] Additionally, for the illustrated embodiment, the seal tooth assembly 150 also includes a base extension 156 extending substantially parallel to the seal support extension 152 (e.g., within 15 degrees of each other), wherein the dampener 154 is positioned between the base extension 156 and the seal support extension 152. Specifically, for the depicted embodiment, the base extension 156 extends parallel to the seal support extension 152. Notably, in the illustrated embodiment, the seal support extension 152 and the base extension 156 both extend along the axial direction A.

[0100] The seal support extension 152 is cantilevered from the base extension 156. Specifically, the seal support extension 152 includes a front end 158, and the base extension 156 similarly includes a front end 160. The front end 158 of the seal support extension 152 is coupled to the front end 160 of the base extension 156. In this manner, it should be appreciated that the seal support extension 152 may bend along the radial direction R in response to various operating conditions of the gas turbine engine 10. For example, it should be appreciated that the rotor cavity 122 is defined between the stator housing 94 and the HP spool 34 including the seal tooth assembly 150. The rotor cavity 122 may define a pressure that generally increases with the pressure at the outlet of the HP compressor 24 (see Fig. 9 ). When the pressure within the rotor cavity 122 increases, the seal support extension 152 may generally bend inwardly along the radial direction R. Conversely, when the rotational speed of the HP spool 34 increases, the centrifugal force may cause the seal support extension 152 to bend outwardly along the radial direction R.

[0101] Furthermore, it should be appreciated that, depending on certain operating conditions of gas turbine engine 10, seal support extension 152 may be susceptible to oscillations, ie, acoustic resonances, caused by pressure fluctuations within rotor cavity 122. Damper 154 may be configured to dampen such oscillations, as described below.

[0102] See now Fig.11 , provides Fig.10 The seal tooth assembly 150 is Fig.10 11-11 in the schematic cross-sectional view of the seal support extension 152 and the base extension 156, it should be understood that the seal support extension 152 and the base extension 156 define a gap 162 therebetween, and the gap 162 has a gap height 164 along the radial direction R. When the gas turbine engine 10 is in a shut-down operating condition (i.e., when the HP spool 34 is not rotating), the gap height 164 can be defined. The damper 154 similarly defines a damper height 166 along the radial direction R that is less than the gap height 164.

[0103] More specifically, it should be understood that for the depicted embodiment, the damper 154 is a first damper 154A coupled to and extending from the seal support extension 152, and the damper height 166 is a first damper height 166A. The seal tooth assembly 150 also includes a second damper 154B coupled to and extending from the base extension 156 and defining a second damper height 166B along the radial direction R, the second damper height 166B also being less than the gap height 164. In this manner, the first damper 154A and the second damper 154B can prevent the seal support extension 152 from moving along the radial direction R toward the base extension 156 beyond a predetermined amount (i.e., an amount of allowable movement equal to the difference between the gap height 164 and the larger of the first damper height 166A and the second damper height 166B).

[0104] In addition, if Fig.11 As can be understood from the view of FIG. 1 , the first damper 154A defines a first lip 168A extending from a distal end of the first damper 154A toward the second damper 154B, and the second damper 154B defines a second lip 168B extending from a distal end of the second damper 154B toward the first damper 154A. The first lip 168A and the second lip 168B overlap along the radial direction R (i.e., the first lip 168A and the second lip 168B each extend through a common circumferential position at an axial position). In this manner, the first damper 154A and the second damper 154B can further prevent the seal support extension 152 from moving away from the base extension 156 along the radial direction R by more than a predetermined amount (e.g., more than an amount equal to the gap between the first lip 168A and the second lip 168B along the radial direction R).

[0105] In particular, it should be understood that the first lip 168A and the second lip 168B together define a radial gap R therebetween. Gap In addition, the second lip 168B defines a circumferential or lateral gap T between the second lip 168B and the first damper 154A. Gap In some embodiments, the lateral gap T GapThe second lip 168B may be very small (e.g., less than 0.1 mm) or zero. With this configuration, when the seal support extension 152 moves relative to the base extension 156, the second lip 168B may rub against the first damper 154A, creating friction and inhibiting such relative movement. In addition, when the seal support extension 152 moves toward the base extension 156 so that the first damper 154A contacts the base extension 156 and the second damper 154B contacts the seal support extension 152, the friction between these components inhibits relative circumferential movement between the seal support extension 152 and the base extension 156. This may occur when the seal teeth 120 temporarily rub into the wear-resistant material of the seal pad 118.

[0106] Note that in the illustrated embodiment, the first and second dampeners 154A, 154B are integrally formed with the seal support extension 152 and the base extension, respectively.

[0107] However, it should be understood that in other exemplary embodiments, the first damper 154A and the second damper 154B can be removably coupled to the seal support extension 152 and the base extension, respectively, using, for example, rotational fasteners. Additionally or alternatively, they can also be coupled using press fit, welding, brazing, or a combination thereof. In this case, the material forming the first damper 154A and the second damper 154B can be the same as the material forming the seal support extension 152 and the base extension, respectively, or can be different.

[0108] Fig.10 and Fig.11 The configuration may thus prevent or reduce the amplitude of oscillation of the seal support extension 152 relative to the base extension 156 by more than a predetermined amount. In addition, when the first damper 154A, the second damper 154B, or both contact the seal support extension 152, one or more of the base extension 156, or the other of the first and second dampers 154A, 154B, this configuration may change the resonant frequency of the seal tooth assembly 150, effectively reducing the sensitivity of the seal tooth assembly 150 to a frequency that matches or is very close to the cavity acoustic frequency, thereby causing undesirable oscillations.

[0109] In addition, when the first damper 154A, the second damper 154B, or both contact the seal support extension 152, the base extension 156 and / or the other of the first damper 154A or the second damper 154B, such contact creates friction between the components, which further suppresses oscillation of the seal support extension 152 relative to the base extension 156.

[0110] Still see Fig.10 and Fig.11, it should be appreciated that the seal support extension 152 and the base extension 156 are integrally formed with each other as a unitary component, such that there is little or no risk of air leakage between the components. Additionally, for the depicted embodiment, the seal tooth assembly 150 is integrally formed with the rotor arm of the HP spool 34 (and more specifically, the front spool member 110 of the HP spool 34) as a unitary component, further reducing the risk of air leakage between the components. For example, in at least some exemplary embodiments, the integrally formed components may be integrally formed via an additive manufacturing process.

[0111] However, in other exemplary embodiments, the seal tooth assembly 150 may be configured according to any other suitable configuration. Fig.12 , provides a close-up schematic illustration of a seal tooth assembly 150 according to another exemplary aspect of the present disclosure. Fig.12 The exemplary seal tooth assembly 150 may be similar to that described above with reference to Fig.10 The depicted example seal tooth assembly 150 is configured in a similar manner, and the same or similar numbers may refer to the same or similar parts.

[0112] For example, the exemplary seal tooth assembly 150 shown generally includes a seal support extension 152, a seal tooth 120 extending from the seal support extension 152 toward the seal gasket 118 (see, e.g., Fig. 9 ), and a damper 154 capable of operating with the seal support extension 152 to inhibit movement along the radial direction R and the axial direction A.

[0113] However, as will be appreciated, Fig.12 In an embodiment, the seal tooth assembly 150 is formed separately from and coupled to the rotor arm of the HP spool 34 (or more specifically, the front spool member 110 of the HP spool 34 in the illustrated embodiment). More specifically, the seal support extension 152 includes a first flange 170 at a front end 158 and a second flange 172 at a rear end 174. The rotor arm includes a first bracket 175 defining a first gap 176, and the first flange 170 is positioned within the first gap 176. The first bracket 175 can be formed integrally with the rotor arm. The rotor arm also includes a second bracket 178 defining a second gap 180, and the second flange 172 is positioned within the second gap 180. The second bracket 178 is configured as a two-piece bracket having a first member 182 formed integrally with the rotor arm and a second member 184 coupled to the first member 182 by a mechanical fastener 186. The second gap 180 is defined by the second member 184. It is worth noting that due to Fig.12 The exemplary seal tooth assembly 150 is modular so that the flanges 170, 172 are able to move within the first gap 176 and the second gap 180, respectively. This movement creates frictional resistance at the interface.

[0114] In addition, for Fig.12 In the embodiment of the present invention, the damper 154 is configured as a sealing ring positioned on the inner side 188 of the seal support extension 152 in the radial direction R. More specifically, the sealing ring is configured as a piston ring contacting the inner side 188 of the seal support extension 152 and is configured to bias the seal support extension 152 outward in the radial direction R.

[0115] The term "seal ring" may refer to any suitable annular / 360 degree structure capable of forming a fluid seal with the inner side 188 of the seal support extension 152. A "piston ring" is a 360 degree ring that elastically deforms when installed so that it presses the inner side 188 of the seal support extension 152 outwardly in the radial direction R. For example, briefly referring to Fig.13 , provides a perspective view of a piston ring 300 according to the present disclosure, which can be used as Fig.12 The piston ring 300 is shown in an undeformed state, wherein the piston ring 300 defines a diameter 302 that is larger than the diameter of the location in which it will be installed (i.e., the inner side 188 of the seal support extension 152 is along the inner side 188 of the seal support extension 152). Fig.12 The piston ring 300 also defines an overlap joint 304 which will overlap when elastically deformed into place.

[0116] See again Fig.12 , for the depicted embodiment, the rotor arm includes a seal housing 190, and the damper 154 is at least partially positioned within the seal housing 190 and is configured to form a fluid seal between the seal housing 190 and the seal support extension 152. In this manner, the damper 154 and the seal housing 190 can reduce a majority of the airflow leakage through the joint between the first flange 170 and the first gap 176, as well as any residual airflow leakage through the joint between the second flange 172 and the second gap 180. As will be appreciated, the joint between the first flange 170 and the first gap 176 is positioned in the high pressure cavity, while the joint between the second flange 172 and the second gap 180 is positioned in the low pressure cavity because a pressure drop occurs across the seal teeth 120.

[0117] Furthermore, when the seal support extension 152 is bent in the radial direction R, such movement may cause the damper 154 to move inward or outward in the radial direction R, resulting in friction between the seal ring and the seal housing 190. The friction generated between the damper 154 and the seal housing 190 may cause damping of such movement of the seal support extension 152 in the radial direction R. Furthermore, when the seal support extension 152 is bent in the radial direction R, all joints between adjacent portions (e.g., the joint between the first flange 170 and the first gap 176, and the joint between the second flange 172 and the second gap 180) may similarly generate friction to provide damping of such movement of the seal support extension 152 in the radial direction R.

[0118] However, it should be appreciated that in other exemplary embodiments, any other suitable dampener 154 operable with the seal support extension 152 to inhibit movement along the radial direction R may be provided.

[0119] For example, see now Fig.14 , provides a close-up schematic illustration of a seal tooth assembly 150 according to yet another exemplary aspect of the present disclosure. Fig.14 The exemplary seal tooth assembly 150 may be similar to that described above with reference to Fig.10 The depicted example seal tooth assembly 150 is configured in a similar manner, and the same or similar numbers may refer to the same or similar parts.

[0120] For example, Fig.14 The seal tooth assembly 150 in the embodiment of the present invention generally includes a seal support extension 152 and a seal tooth 120 extending from the seal support extension 152 toward the seal pad 118, and a damper 154 capable of operating with the seal support extension 152 to inhibit movement in the radial direction R. Specifically, for the illustrated embodiment, the seal support extension 152 defines a channel 192 extending in the circumferential direction C on an inner surface 194 of the seal support extension 152 in the radial direction R. For the depicted embodiment, the damper 154 is a damping ring that is at least partially positioned within the channel 192 defined by the seal support extension 152 and extends 360 degrees (or at least 300 degrees; see Fig.13 As the rotational speed of the HP spool 34 including the seal tooth assembly 150 increases, centrifugal force may press the damper 154 (or more specifically the damping ring) into the channel 192, increasing the stiffness of the seal support extension 152 and changing the resonant frequency of the seal support extension 152. In this manner, the damping ring may be configured to inhibit movement along the radial direction R.

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

[0122] A gas turbine engine defining a radial direction, the gas turbine engine comprising: a compressor section, a combustion section and a turbine section arranged in a series flow order and defining a working gas flow path, the compressor section comprising a compressor, the compressor comprising a last compressor stage; a spool, the spool being drivingly coupled to the compressor; a stator blade stage, the stator blade stage being located downstream of the last compressor stage; and a stator casing, the stator casing being coupled to the stator blade stage inside the stator blade stage along the radial direction, the spool and the stator casing together defining a rotor cavity in fluid communication with the working gas flow path, the stator blade stage comprising a first stator blade defining a fluid passage, and the stator casing defining an air chamber and a supplementary air flow passage, the air chamber being in fluid communication with the fluid passage in the first stator blade, the supplementary air flow passage being in fluid communication with the air chamber and the rotor cavity for providing air flow to the rotor cavity.

[0123] A gas turbine engine as claimed in any preceding clause, wherein the gas turbine engine defines a circumferential direction, and wherein the plenum extends in the circumferential direction.

[0124] A gas turbine engine as claimed in any preceding clause, wherein the gas turbine engine defines a circumferential direction, wherein the supplemental air flow channel is a first supplemental air flow channel in an array of supplemental air flow channels defined by the stator casing, the array of supplemental air flow channels being arranged along the circumferential direction.

[0125] A gas turbine engine according to any of the preceding clauses, wherein the gas turbine engine defines an axial direction and a reference plane extending in the axial direction and in a radial direction, wherein the supplementary air flow channel passes through the reference plane and includes an air flow outlet defining an air flow outlet direction, wherein the air flow outlet direction defines a swirl angle greater than 0 degrees with the reference plane.

[0126] A gas turbine engine as claimed in any preceding clause, wherein the swirl angle is greater than 0 degrees and less than 85 degrees.

[0127] A gas turbine engine as claimed in any preceding clause, wherein the fluid passage of the first stator vane is in fluid communication with the compressor.

[0128] A gas turbine engine according to any of the preceding clauses, wherein the compressor is a high-pressure compressor, and wherein the stator vane stage is an exhaust nozzle stage that fluidly connects the high-pressure compressor to the combustion section, and wherein the fluid passage of the first stator vane is in fluid communication with the high-pressure compressor, in direct fluid communication with the working gas flow path through the exhaust nozzle stage, or both.

[0129] A gas turbine engine as claimed in any preceding clause, further comprising a heat exchanger in fluid communication with the fluid passage of the first stator vane at a location downstream from the compressor section.

[0130] A gas turbine engine according to any preceding clause, wherein the first stator vane defines an opening at a location within the working gas flow path for receiving a working gas flow path gas flow therefrom, wherein the fluid passage is in fluid communication with the opening.

[0131] A gas turbine engine according to any preceding clause, wherein the stator casing and rotor disk together form a compressor discharge pressure seal and together define a rotor cavity opening, and wherein the rotor cavity extends between the compressor discharge pressure seal and the rotor cavity opening.

[0132] A gas turbine engine defines an axial direction, a radial direction, and a reference plane extending in the axial direction and the radial direction: a compressor section, a combustion section, and a turbine section arranged in a series flow order and defining a working gas flow path, the compressor section including a compressor, the compressor including a last compressor stage; a spool, the spool drivingly coupled to the compressor; a stator blade stage, the stator blade stage being located downstream of the last compressor stage; and a stator casing, the stator casing being coupled to the stator blade stage inside the stator blade stage along the radial direction, the spool and the stator casing together defining a rotor cavity in fluid communication with the working gas flow path, the stator blade stage including a first stator blade defining a fluid passage, and the stator casing defining a supplementary airflow passage, the supplementary airflow passage being in fluid communication with the fluid passage and the rotor cavity, the supplementary airflow passage passing through the reference plane and defining an airflow outlet direction, the airflow outlet direction defining a swirl angle greater than 0 degrees with the reference plane.

[0133] A gas turbine engine defining an axial direction and a radial direction, the gas turbine engine comprising: a compressor section, a combustion section, and a turbine section arranged in a series flow order and defining a working gas flow path, the compressor section comprising a compressor, the compressor comprising a final compressor stage and defining a compressor outlet temperature T3 in degrees Rankine during operating conditions of the gas turbine engine; a spool drivingly coupled to the compressor; a stator vane stage, the stator vane stage being located downstream of the final compressor stage; and a stator casing coupled to the stator vane stage in the radial direction inside the stator vane stage, the spool and the stator casing together defining a rotor cavity in fluid communication with the working gas flow path, wherein the stator vane stage defines one or more fluid passages, and the stator casing defines a plurality of supplemental gas flow passages in fluid communication with the one or more fluid passages and the rotor cavity, wherein the plurality of supplemental gas flow passages are configured to provide a flow path at a passage temperature T3 in degrees Rankine during the operating conditions of the gas turbine engine. P The rotor cavity is provided with a supplementary airflow, wherein each of the plurality of supplementary airflow channels defines a swirl angle θ at a local reference plane. Swirl , each local reference plane passes through a corresponding supplementary airflow channel of the plurality of supplementary airflow channels and extends in the axial direction and the radial direction, wherein the gas turbine engine defines a value equal to T P The ratio of the supplementary air flow temperature ratio (SATR) to T3, and wherein the SATR is less than 1.0 and θ Swirl is greater than or equal to 0 degrees and less than or equal to 85 degrees, or where SATR is greater than or equal to 1.0 and θ Swirl Greater than 0 degrees and less than or equal to 85 degrees.

[0134] A gas turbine engine according to any preceding clause, wherein the spool and stator housing together define a rotor cavity opening, wherein the rotor cavity opening defines a cross-sectional area A FO , wherein the plurality of supplemental airflow channels define a total cross-sectional area A SAP_Total , and where A FO With A SAP_Total The ratio is greater than or equal to 0.05 and less than 0.25.

[0135] A gas turbine engine as in any preceding clause, wherein the SATR is greater than or equal to 0.8 and less than or equal to 1.15.

[0136] A gas turbine engine according to any preceding clause, wherein θ Swirl The relationship between θ and STAR is as follows: Swirl ≥199×SATR2 -186×SATR+7.

[0137] A gas turbine engine as claimed in any preceding clause, wherein the stator casing defines a plenum, and wherein the plurality of supplemental gas flow passages are in fluid communication with the one or more fluid passages through the plenum.

[0138] A gas turbine engine as claimed in any preceding clause, wherein the gas turbine engine defines a circumferential direction, and wherein the plenum extends in the circumferential direction.

[0139] A gas turbine engine as claimed in any preceding clause, wherein said one or more fluid passages are in fluid communication with said compressor.

[0140] A gas turbine engine as claimed in any preceding clause, further comprising: a heat exchanger in fluid communication with the one or more fluid passages at a location downstream of the compressor, and wherein the SATR is less than 1.

[0141] A gas turbine engine as claimed in any preceding clause, wherein the compressor is a high pressure compressor, and wherein the stator bucket stage is an exhaust nozzle stage fluidly connecting the high pressure compressor to the combustion section.

[0142] A gas turbine engine defining radial, circumferential and axial directions, the gas turbine engine comprising: a compressor section, a combustion section and a turbine section arranged in a series flow order and defining a working gas flow path, the compressor section comprising a compressor, the compressor comprising a last compressor stage; a stator blade stage located downstream of the last compressor stage; a stator casing, the stator casing being connected to the stator blade stage inside the stator blade stage along the radial direction, the stator casing comprising a sealing gasket; and a spool drivingly connected to the compressor, the spool and the stator casing together defining a rotor cavity fluidically connected to the working gas flow path, the spool comprising a sealing tooth assembly, the sealing tooth assembly comprising a sealing support extension, sealing teeth extending from the sealing support extension to the sealing gasket, and a damper capable of operating with the sealing support extension.

[0143] A gas turbine engine as claimed in any preceding clause, wherein the damper is operable to inhibit movement of the seal support extension in the radial direction, the circumferential direction, the axial direction or a combination thereof.

[0144] A gas turbine engine as claimed in any preceding clause, wherein the seal tooth assembly includes a base extension extending substantially parallel to the seal support extension, and wherein the dampener is positioned between the base extension and the seal support extension.

[0145] A gas turbine engine as claimed in any preceding clause, wherein the seal tooth assembly includes a base extension extending parallel to the seal support extension, and wherein the dampener is positioned between the base extension and the seal support extension.

[0146] A gas turbine engine as claimed in any preceding clause, wherein the seal support extension is cantilevered from the base extension.

[0147] A gas turbine engine according to any preceding clause, wherein the seal support extension and the base extension define a gap having a gap height along the radial direction, wherein the damper defines a damper height along the radial direction that is less than the gap height.

[0148] A gas turbine engine according to any preceding clause, wherein the damper is a first damper coupled to and extending from the seal support extension, and wherein the seal tooth assembly further includes a second damper coupled to and extending from the base extension.

[0149] A gas turbine engine as claimed in any preceding clause, wherein the first damper defines a first lip, wherein the second damper defines a second lip, and wherein the first lip and the second lip overlap in the radial direction.

[0150] A gas turbine engine as claimed in any preceding clause, wherein the seal support extension and the base extension are integrally formed.

[0151] A gas turbine engine as claimed in any preceding clause, wherein the spool further comprises a rotor arm extending at least partially along the radial direction, and wherein the seal tooth assembly is integrally formed with the rotor arm.

[0152] The gas turbine engine of any preceding clause, wherein the spool further comprises a rotor arm extending at least partially along the radial direction, and wherein the seal tooth assembly is formed separately from and coupled to the rotor arm.

[0153] A gas turbine engine as claimed in any preceding clause, wherein the damper is configured as a seal ring positioned on an inner side of the seal support extension in the radial direction.

[0154] A gas turbine engine as claimed in any preceding clause, wherein the seal ring is a piston ring configured to bias the seal support extension outwardly in the radial direction.

[0155] A gas turbine engine as described in any preceding clause, wherein the rotor arm comprises an outer casing, and wherein the seal ring is at least partially positioned within the outer casing to form a fluid seal between the outer casing and the seal support extension.

[0156] A gas turbine engine according to any preceding clause, wherein the rotor cavity is in fluid communication with a portion of the working gas flow path, wherein the seal tooth assembly and the seal pad form a fluid seal with the rotor cavity.

[0157] A seal tooth assembly for a spool of a gas turbine engine, the gas turbine engine including a stator housing having a sealing gasket, the seal tooth assembly comprising: a seal support extension; a seal tooth extending from the seal support extension and further extending toward the sealing gasket of the stator housing when the seal tooth assembly is installed in the gas turbine engine; and a damper operable with the seal support extension to inhibit movement.

[0158] A seal tooth assembly as described in any preceding clause, wherein the dampener is operable to inhibit movement of the seal support extension in the radial direction, the circumferential direction, the axial direction, or a combination thereof.

[0159] The seal tooth assembly of any preceding clause, further comprising: a base extension extending substantially parallel to the seal support extension, wherein the dampener is positioned between the base extension and the seal support extension.

[0160] The seal tooth assembly of any preceding clause, further comprising: a base extension extending parallel to the seal support extension, wherein the dampener is positioned between the base extension and the seal support extension.

[0161] A seal tooth assembly as claimed in any preceding clause, wherein the seal support extension is cantilevered from the base extension.

[0162] The seal tooth assembly of any preceding clause, wherein the seal support extension and the base extension define a gap having a gap height along the radial direction, wherein the damper defines a damper height along the radial direction that is less than the gap height.

[0163] A seal tooth assembly according to any preceding clause, wherein the dampener is a first dampener coupled to and extending from the seal support extension, and wherein the seal tooth assembly further includes a second dampener coupled to and extending from the base extension.

[0164] A seal tooth assembly according to any preceding clause, wherein the first damper defines a first lip, wherein the second damper defines a second lip, and wherein the first lip and the second lip overlap in the radial direction.

[0165] The seal tooth assembly according to any preceding clause, wherein the dampener is configured as a seal ring positioned on an inner side of the seal support extension along the radial direction.

[0166] This written description uses examples to disclose the present disclosure, including the best mode, and also to enable any person skilled in the art to practice the present disclosure, including making and using any device or system and performing any incorporated method. The patentable scope of the present disclosure is defined by the claims, and may 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 that do not differ substantially from the literal language of the claims.

Claims

1. A gas turbine engine with a defined radial direction, characterized in that: The gas turbine engine comprises: a compressor section, a combustion section, and a turbine section arranged in series flow order and defining a working gas flow path, the compressor section including a compressor including a final compressor stage; a spool drivingly coupled to the compressor; a stator vane stage located downstream of the last compressor stage; and a stator casing, the stator casing being coupled to the stator vane stage in the radial direction inside the stator vane stage, the bobbin and the stator casing together defining a rotor cavity in fluid communication with the working gas flow path, the stator vane stage comprising a first stator vane defining a fluid passage, and the stator casing defining an air chamber and a supplementary air flow passage, the air chamber being in fluid communication with the fluid passage in the first stator vane, the supplementary air flow passage being in fluid communication with the air chamber and the rotor cavity for providing air flow to the rotor cavity.

2. The gas turbine engine according to claim 1, characterized in that Wherein the gas turbine engine defines a circumferential direction, and wherein the plenum extends in the circumferential direction.

3. The gas turbine engine according to claim 1, characterized in that The gas turbine engine defines a circumferential direction, wherein the supplemental airflow channel is a first supplemental airflow channel in an array of supplemental airflow channels defined by the stator casing, the array of supplemental airflow channels being arranged along the circumferential direction.

4. The gas turbine engine according to claim 1, characterized in that The gas turbine engine further defines an axial direction and a reference plane extending in the axial direction and in a radial direction, wherein the supplemental airflow passage passes through the reference plane and includes an airflow outlet defining an airflow outlet direction, wherein the airflow outlet direction defines a swirl angle greater than 0 degrees with the reference plane.

5. The gas turbine engine according to claim 4, characterized in that The swirl angle is greater than 0 degree and less than 85 degrees.

6. The gas turbine engine according to claim 1, characterized in that The fluid passage of the first stator blade is in fluid communication with the compressor.

7. The gas turbine engine according to claim 6, characterized in that wherein the compressor is a high-pressure compressor, and wherein the stator vane stage is an exhaust nozzle stage fluidly connecting the high-pressure compressor to the combustion section, and wherein the fluid passage of the first stator vane is in fluid communication with the high-pressure compressor, in direct fluid communication with the working gas flow path through the exhaust nozzle stage, or both.

8. The gas turbine engine according to claim 1, characterized in that Further including: A heat exchanger is in fluid communication with the fluid passage of the first stator vane at a location downstream of the compressor section.

9. The gas turbine engine according to claim 1, characterized in that The first stator vane defines an opening at a position within the working gas flow path for receiving a working gas flow path gas flow from the working gas flow path, wherein the fluid channel is in fluid communication with the opening.

10. The gas turbine engine according to claim 1, characterized in that wherein the stator housing and the rotor disk together form a compressor discharge pressure seal and together define a rotor cavity opening, and wherein the rotor cavity extends between the compressor discharge pressure seal and the rotor cavity opening.

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