Seal assembly for a turbine engine

Abstract

A turbine engine includes a rotor, a stator having an aft wall, and a seal assembly having a plurality of seal segments disposed between the rotor and the stator. The rotor, the stator, and the seal assembly are arranged together to define a high pressure region and a low pressure region. The turbine engine also includes at least one biasing member engaged with one or more of the plurality of sealing segments. The plurality of seal segments includes a primary seal segment and a secondary seal segment connected together via a flexible joint. Thus, the flexible joint allows angular misalignment between the primary and secondary seal segments, allowing the primary seal segment to move with the rotor while the secondary seal segment remains in contact with the rear wall of the stator.

Description

TECHNICAL FIELD

[0001] The present disclosure generally relates to turbine engines and, more particularly, to a seal assembly for a turbine engine. BACKGROUND OF THE DISCLOSURE

[0002] A gas turbine engine, such as a turbofan engine, can be used for aircraft propulsion. A turbofan engine typically includes a bypass fan section and a turbine, such as a gas turbine engine, to drive the bypass fan. The turbine typically includes a compressor section, a combustion section, and a turbine section in a series flow arrangement. Both the compressor section and the turbine section are driven by one or more rotor shafts and typically include multiple rows or stages of rotor blades coupled to the rotor shafts. Each individual row of rotor blades is axially spaced from a consecutive row of rotor blades by a corresponding row of stator or stationary vanes. A radial gap is formed between the inner surface of the stator vanes and the outer surface of the rotor shaft. A gas turbine engine can also include various seals to reduce and / or prevent leakage of flow (e.g., working fluid flow) between various components of the gas engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0003] A complete and enabling disclosure of the present invention, including the best mode thereof, to one of ordinary skill in the art, is set forth in the specification, which makes reference to the accompanying drawings, in which:

[0004] Figure 1 is a cross-sectional view of a gas turbine engine in accordance with an aspect of the present disclosure.

[0005] Figure 2 is Figure 1 a cross-sectional schematic view of a portion of the turbine of

[0006] Figure 3 is a detailed schematic cross-sectional view of a portion of the turbine of Figure 2 taken along line 3-3.

[0007] Figure 4 is Figure 3 a detailed schematic cross-sectional view of a rotor, stator, carrier, and seal assembly of

[0008] Figure 5 is a detailed perspective view of a portion of a seal assembly in accordance with an aspect of the present disclosure, particularly showing a rear side of a seal segment of the seal assembly.

[0009] Figure 6 is another detailed perspective view of a portion of a seal assembly in accordance with an aspect of the present disclosure, particularly showing a rotor side of a seal segment of the seal assembly. DETAILED DESCRIPTION

[0010] Reference will now be made in detail to the present embodiments of the disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and alphabetical identifiers to refer to features in the drawings. Similar or like identifiers in the drawings and description have been used to refer to similar or like parts of the disclosure.

[0011] Unless the context clearly dictates otherwise, the singular forms "a", "an", and "the" include plural referents.

[0012] For example, the term "at least one" in the context of "at least one of A, B, and C" refers to only A, only B, only C, or any combination of A, B, and C.

[0013] The term "turbine" refers to a machine that includes one or more compressors, a heat addition section (e.g., a combustion section), and one or more turbines that together produce a torque output.

[0014] The term "gas turbine engine" or "turbine engine" refers to an engine that has a turbine as all or part of its power source. Example gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, etc., as well as hybrid electric versions of one or more of these engines.

[0015] The term "combustion section" refers to any heat addition system of a turbine. 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 components. In certain example embodiments, the combustion section may include an annular burner, a can burner, a tubo burner, a trapped vortex combustor (TVC), or other suitable combustion systems, or combinations thereof.

[0016] The terms "low" and "high" or their respective comparatives (e.g., "lower" and "higher" where applicable), when used in conjunction with components such as compressors, turbines, shafts, or spools, refer to relative speeds within the engine unless otherwise specified. For example, a "low turbine" or "low-speed turbine" defines a component configured to operate at a rotational speed (e.g., a maximum allowable speed) lower than that of a "high turbine" or "high-speed turbine" of the engine.

[0017] The terms "front" and "rear" refer to relative positions within a gas turbine engine or a vehicle. For example, for a gas turbine engine, front refers to a position closer to the engine inlet, and rear refers to a position closer to the engine nozzle or exhaust.

[0018] The terms "upstream" and "downstream" refer to relative directions with respect to fluid flow in a fluid path. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction towards which the fluid flows.

[0019] The present disclosure generally relates to an aerodynamic seal assembly for a turbine of a gas turbine engine. A turbine typically includes a compressor section including a low-pressure compressor and a high-pressure compressor, a combustion section, and a turbine section including a high-pressure turbine and a low-pressure turbine arranged in a series flow order. Each of the low-pressure compressor, high-pressure compressor, high-pressure turbine, and low-pressure turbine includes a continuous row of stationary or stator vanes axially spaced apart by a continuous row of rotor blades. The rotor blades are typically coupled to a rotor shaft, and the stator vanes are circumferentially mounted in an annular configuration around the outer surface of the rotor shaft. A radial gap is formed between the outer surface of the rotor shaft and the inner portion of each ring or row of stator vanes. A radial gap may also be formed between the outer surface of the rotor shaft and the inner portion of a non-rotating stationary part of the engine.

[0020] During operation, it is desirable to control (reduce or prevent) leakage of compressed air flow or combustion gas flow through these radial gaps. Accordingly, the seal assembly includes a seal segment for sealing these radial gaps. Additionally, during operation of a gas turbine engine, the stator is fixed in place while the rotor rotates relative to the stator. Due to centrifugal loads, thermal loads (e.g., non-uniform temperatures of the rotor and stator), and pressure-induced forces caused by the rotation, there is relative movement between the rotor and the stator. The relative movement between the rotor and the stator can be a relative radial displacement or a relative angular displacement about a tangential axis (e.g., coning about a tangential axis). During such relative angular displacement (e.g., about a tangential axis) between the rotor and the stator, the seal assembly (and thus the seal segment) would be desired to cone or pitch with the rotor, but this cannot be done because the seal segment is loaded against the back wall of the stator.

[0021] Accordingly, an aerodynamic seal assembly is disclosed herein having a seal segment disposed between a rotor and a stator. As used herein, an aerodynamic seal generally refers to a mechanical seal that uses a dynamic rotor and one or more grooves on the rotor or stator that form an air film and the opposing seal surfaces ride on the air film. Specifically, the rotor, stator, and seal assembly are arranged together to define a high-pressure region and a low-pressure region. A biasing member (e.g., a spring) engages the seal segment. Additionally, the seal assembly is segmented into a plurality of seal segments. Specifically, in an embodiment, the plurality of seal segments includes a primary seal segment and a secondary seal segment connected together via a flexure joint. Thus, the flexure joint allows for angular misalignment between the rotor and the stator. More specifically, the primary seal segment moves with the rotor, and the secondary seal segment is loaded by the back wall of the stator. In an embodiment, the seal segment can be a two-piece (or two-body) seal including a primary seal segment and a secondary seal segment.

[0022] In addition, in an embodiment, to minimize leakage of high-pressure fluid from the high-pressure region to the low-pressure region through the split dual seal, the seal assembly may further include a static seal disposed between the primary seal section and the secondary seal section. In such an embodiment, one or both of the primary seal section and the secondary seal section may include one or more grooves for receiving the static seal.

[0023] Referring now to the drawings, in which like numerals represent like elements throughout the figures, Figure 1 is a schematic cross-sectional view of a gas turbine engine 10 according to an embodiment of the present disclosure. More specifically, for Figure 1 the embodiment, the gas turbine engine 10 is a high-bypass turbofan jet engine, sometimes also referred to as a "turbofan engine". As Figure 1 shown, the gas turbine engine 10 defines an axial direction A (extending parallel to a reference longitudinal centerline 12), a radial direction R, and a circumferential direction C extending around the longitudinal centerline 12. Generally, the gas turbine engine 10 includes a fan section 14 and a turbine 16 located downstream of the fan section 14.

[0024] The turbine 16 shown generally includes a substantially tubular outer casing 18 that defines an annular inlet 20. The outer casing 18 surrounds a compressor section in a series flow relationship, the 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 exhaust nozzle section 32. A high-pressure (HP) shaft 34 (which may alternatively or additionally be a spool) drivingly connects the HP turbine 28 to the HP compressor 24. A low-pressure (LP) shaft 36 (which may alternatively or additionally be a spool) drivingly connects the LP turbine 30 to the LP compressor 22. The compressor section, the combustion section 26, the turbine section, and the exhaust nozzle section 32 together define a working gas flow path 37.

[0025] For the illustrated embodiment, the fan section 14 includes a fan 38 having a plurality of fan blades 40 that are coupled to a disk 42 in a spaced-apart manner. As shown, the fan blades 40 extend generally radially outwardly from the disk 42 along a radial direction R. Each fan blade 40 is rotatable relative to the disk 42 about a pitch axis P in that the fan blades 40 are operatively coupled to a suitable pitch change mechanism 44 configured to commonly change (e.g., uniformly change) the pitch of the fan blades 40. The gas turbine engine 10 further includes a power gearbox 46, and the fan blades 40, disk 42, and pitch change mechanism 44 can rotate together about a longitudinal centerline 12 across the power gearbox 46 via an LP shaft 36. The power gearbox 46 includes a plurality of gears for regulating the rotational speed of the fan 38 relative to the LP shaft 36 such that the fan 38 can rotate at a more efficient fan speed.

[0026] Still referring to Figure 1 the embodiment, the disk 42 is covered by a rotatable front hub 48 (sometimes also referred to as a “ spinner ”) of the fan section 14. The front hub 48 has an aerodynamic profile to facilitate air flow through the plurality of fan blades 40.

[0027] In addition, the fan section 14 includes an annular fan casing or nacelle 50 that circumferentially surrounds at least a portion of the fan 38 and / or the turbine 16. It should be understood that in the illustrated embodiment, the nacelle 50 is supported relative to the turbine 16 by a plurality of circumferentially spaced outlet guide vanes 52. Further, a downstream section 54 of the nacelle 50 extends over an outer portion of the turbine 16, thereby defining a bypass air flow passage 56 therebetween.

[0028] During operation of the gas turbine engine 10, a quantity of air 58 enters the gas turbine engine 10 through the nacelle 50 and an associated inlet 60 of the fan section 14. As the quantity of air 58 passes through the fan blades 40, a first portion 62 of the air is directed or guided into the bypass air flow passage 56, while a second portion 64 of the air, as indicated by arrow 64, is directed or guided into the working gas flow path 37, or more specifically, into the LP compressor 22. The ratio between the first portion 62 of the air and the second portion 64 of the air is commonly referred to as the bypass ratio. Then, the pressure of the second portion 64 of the air increases as it is directed through the HP compressor 24 and into the combustion section 26, where it is mixed with fuel and burned to produce combustion gases 66.

[0029] The combustion gases 66 are directed through the HP turbine 28, where a portion of the thermal and / or kinetic energy from the combustion gases 66 is extracted via successive 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, causing the HP shaft 34 to rotate and thereby supporting the operation of the HP compressor 24. The combustion gases 66 are then directed through the LP turbine 30, where a second portion of the thermal and kinetic energy is extracted from the combustion gases 66 via successive 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, causing the LP shaft 36 to rotate and thereby supporting the operation of the LP compressor 22 and / or the rotation of the fan 38.

[0030] Subsequently, the combustion gases 66 are directed through the jet exhaust nozzle section 32 of the turbine 16 to provide propulsive thrust. At the same time, when a first portion 62 of the air is directed through the bypass air flow passage 56 before being discharged from the fan nozzle exhaust section 76 of the gas turbine engine 10, the pressure of the first portion 62 of the air is significantly increased, also providing 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 directing the combustion gases 66 through the turbine 16.

[0031] However, it should be recognized that Figure 1 the gas turbine engine 10 shown is by way of example only, and in other embodiments, the gas turbine engine 10 may have any other suitable configuration. For example, although the gas turbine engine 10 shown is configured as a ducted gas turbine engine (e.g., including an outer nacelle 50), in other embodiments, the gas turbine engine 10 may be a non-ducted gas turbine engine (such that the fan 38 is a non-ducted fan and the exit guide vanes 52 project cantilevered from, for example, the outer casing 18).

[0032] Additionally, or alternatively, although the gas turbine engine 10 shown is configured as a geared gas turbine engine (e.g., including a power gearbox 46) and a variable pitch gas turbine engine (e.g., including a fan 38 configured as a variable pitch fan), in other embodiments, the gas turbine engine 10 may 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 cannot rotate about the pitch axis P), or both. It should also be understood that in other embodiments, aspects of the present disclosure may be incorporated into any other suitable gas turbine engine. For example, in other embodiments, aspects of the present disclosure may (as appropriate) be incorporated into, for example, a turboprop gas turbine engine, a turboshaft gas turbine engine, or a turbojet gas turbine engine.

[0033] Now referring toFigure 2 and Figure 3 , provide various views of a portion of turbine 16 of Figure 1 . Specifically, Figure 2 shows Figure 1 a schematic cross-sectional view of a portion of turbine 16 of Figure 3 shows Figure 1 a detailed schematic cross-sectional view of a portion of turbine 16 of Figure 2 and Figure 3 generally shown in, turbine 16 typically includes a rotor 100, a stator 102, and a seal assembly 106 disposed between the rotor 100 and the stator 102. The rotor 100 can be any rotor of the turbine 16, such as the LP shaft 36, the HP shaft 34, etc. As an example, briefly referring back to Figure 1 , a circle SA has been added to Figure 1 to provide an example location where the seal assembly 106 of the present disclosure can be incorporated into the turbine of the present disclosure.

[0034] As Figure 3 specifically shown in, the stator 102 further includes stator vanes 115, and the seal assembly 106 is located at the inner end of the stator vanes 115 in the radial direction R of the turbine 16 in the illustrated embodiment. The turbine 16 further includes a first stage 117 and a second stage 121 of rotor blades 119 spaced apart along the axial direction A of the gas turbine engine 10. The seal assembly 106 is located between the first stage 117 and the second stage 121 of the rotor blades 119 along the axial direction A.

[0035] In the illustrated embodiment, the seal assembly 106 is located within the turbine section of the gas turbine engine 10, such as within the HP turbine 28 or the LP turbine 30. In this way, it can be understood that the rotor 100 can be a rotor coupled to the HP turbine 28, such as the HP shaft 34, or a rotor coupled to the LP turbine 30, such as the LP shaft 36. More specifically, in the illustrated embodiment, the rotor 100 is a connector extending between a disk 123 of the first stage 117 of the rotor blades 119 and a disk 125 of the second stage 121 of the rotor blades 119. However, it can be understood that in other embodiments, the seal assembly 106 can be integrated into, for example, the compressor section of the gas turbine engine 10.

[0036] Still referring to Figure 2 , as will be explained in more detail below, the seal assembly 106 includes a plurality of seal segments 108 extending between the rotor 100 and the stator 102 in the circumferential direction C. In addition, in an embodiment, as Figure 2 and Figure 4 shown, the seal segments 108 together define a sealing surface 116. In addition, as Figure 4As specifically shown, the rotor 100, stator 102, and seal section 108 are arranged together to define a high-pressure region 110 and a low-pressure region 112. In a particular embodiment, as Figure 4 shown, the high-pressure region 110 is located in front of the low-pressure region 112.

[0037] It should be understood that the seal section 108 can be in fluid communication with a high-pressure air source to supply high-pressure fluid (e.g., P HIGH ) flow to the seal section 108. In at least some aspects, the high-pressure air source can be the working gas flow path 37 provided by the gas turbine engine 10 and the seal assembly 106, e.g., at the high-pressure region 110 of the seal assembly 106.

[0038] In addition, as Figure 2 , Figure 3 and Figure 4 shown, the seal assembly 106 includes at least one biasing member 114, such as an annular spring, which engages the seal section 108 to urge contact between the seal section 108 and the rotor 100. Specifically, as Figure 2 , Figure 3 and Figure 4 shown, the biasing member 114 is shown surrounding multiple seal sections 108 simultaneously but not connected to the stator 102. In such an embodiment, the biasing member 114 is configured as an annular spring that urges the seal section 108 toward the rotor 100. Alternative embodiments are possible, where a compression spring or a helical spring connects the seal section 108 to the stator 102. In such an embodiment, the compression spring pushes the seal section 108 away from the rotor 100. Such an embodiment is generally referred to herein as a retraction spring. Further, in such an embodiment, in the absence of differential seal pressure, the retraction spring pulls the seal section 108 away from the rotor 100 or "retracts" it away from the rotor 100. Additionally, upon pressurization, the pressure overcomes the biasing force of the biasing member 114 and pushes the seal section 108 toward the rotor 100. In further embodiments, other types of biasing members (e.g., tension springs, beam flexures, shape memory components, etc.) can be used to connect the seal section 108 to the stator 102 and / or the rotor 100.

[0039] In addition, as Figure 3 and Figure 4 shown, the seal assembly 106 can include a groove 118 in the sealing surface 116 of the seal section 108 and a primary seal dam 104. Further, as Figure 6 shown, in an embodiment, the groove 118 and the primary seal dam 104 are rotated approximately 90 degrees about the edge of a seal section 108, indicating that the groove 118 and the primary seal dam 104 are not only circumferential in direction.

[0040] Specific reference is made to Figure 4 and Figure 5, showing the Figure 2 Various cross-sectional views of different embodiments of portions of the turbine 16 are shown. In particular, as shown in FIG. Figure 4 and Figure 5 As shown, the seal assembly 106 is divided into a plurality of seal segments 108. More specifically, as shown, the plurality of seal segments 108 include a plurality of seal segments 108 connected to the seal assembly 106 by a flexure joint 126 ( Figure 4 ) is connected together with a primary seal segment 122 and a secondary seal segment 124. For example, in an embodiment, the flexure joint 126 can be a flexure pivot joint, a hinge joint, a flexure beam, a pin joint or the like. In addition, as shown, the secondary seal segment 124 is located behind the primary seal segment 122. Therefore, the flexure joint 126 allows relative movement between the primary seal segment 122 and the secondary seal segment 124, so that the primary seal segment 122 can track the rotor 100 and the secondary seal segment 124 can remain aligned with the stator 102. In general, this configuration allows the flexure joint 126 to allow the seal segment 108 to absorb the angular misalignment between the rotor 100 and the stator 102. As used herein, a flexure joint can ensure bending movement (e.g., bending), thereby allowing rotational freedom between adjacent elements connected by the bending.

[0041] More specifically, due to centrifugal loads caused by rotation, thermal loads (e.g., uneven temperatures of the rotor 100 and the stator 102), and pressure-induced forces, there is a relative motion between the rotor 100 and the stator 102. The relative motion between the rotor 100 and the stator 102 can be a relative radial displacement around a tangential axis or a relative angular displacement (e.g., a taper around one of the tangential axes T1, T2, T3, T4 of the corresponding seal segment 108, such as Figure 3 During such relative angular displacements between the rotor 100 and the stator 102 (about respective tangential axes T1, T2, T3, T4), the seal segment 108 allows the primary seal segment 122 to taper / pitch with the rotor 100, wherein the secondary seal segment 124 remains loaded against the rear wall 142 of the stator 102.

[0042] In a specific embodiment, if Figure 4 and Figure 5 As shown, the primary seal segment 122 defines an L-shaped cross-sectional shape in the axial direction A. Further, as shown, the L-shaped cross-sectional shape generally defines a flange 128. Further, as shown, the secondary seal segment 124 generally defines a substantially rectangular cross-sectional shape in the axial direction A. Thus, as shown in the illustrated embodiment, the secondary seal segment 124 is located on the flange 128 of the L-shaped cross-sectional shape of the primary seal segment 122.

[0043] In a further embodiment, Figures 4 to 6As shown, the sealing section 108, the rotor 100, or both may include at least one passage 130 to connect the high-pressure region 110 to the recess 118. Alternatively, the primary seal member 122 may include one or more holes formed therein in the radial direction R, for example, to convey high-pressure air from the high-pressure region 110 to the recess 118. Additionally, the passage 130 is configured to separate at least two aerodynamic features 132 (such as Figure 6 shown as Rayleigh steps). The aerodynamic features 132 represent a portion of the primary seal section 122 that is radially farther from the rotor 100 than the radial face 136 of the primary seal section 122. Nominally, the film thickness of the aerodynamic features 132 of the primary seal section 122 is slightly greater than that of the radial face 136 of the primary seal section 122. In the presence of a rotational speed, this variation in film thickness results in an aerodynamic lift and film between the primary seal section 122 and the rotating rotor 100. Additionally, in an embodiment, the aerodynamic features 132 may be spiral grooves, herringbone grooves, or the like.

[0044] Specifically referring to Figure 4 and Figure 5 , the secondary seal section 124 is in physical contact with the rear wall 142. In such an embodiment, the contact is made through the secondary seal dam 134, which surrounds to form an inner pocket 138 between the secondary seal section 124 and the rear wall 142. Thus, the air in the high-pressure region 110 can fill the inner pocket 138 and assist in the axial force balance of the seal assembly 106, while the secondary seal dam 134 ensures minimal air leakage from the inner pocket 138 to the downstream cavity.

[0045] Still referring to Figure 4 and Figure 5 , the seal assembly 106 may further include a static seal 140 disposed between the primary seal section 122 and the secondary seal section 124. For example, in an embodiment, the static seal 140 may be a spline seal, a C-type seal, an E-type seal, a rope seal, or a W-type seal. In such embodiments, the static seal 140 reduces the leakage of air from the high-pressure region 110 to the downstream cavity through the gap formed between the primary seal section 122 and the secondary seal section 124. In additional embodiments, the seal assembly 106 may further include one or more coatings to minimize radial friction or wear. In an embodiment, for example, the coating may be applied to the inner pocket 138 and / or the rear wall 142.

[0046] It should now be understood that the present disclosure is generally directed to an aerodynamic seal assembly for a turbine of a gas turbine engine. More specifically, the present disclosure is directed to a seal assembly including a seal segment for sealing a radial gap between a rotor and a stator. Additionally, the seal assembly includes a biasing member engaged with the seal segment. The seal segment includes a primary seal segment and a secondary seal segment connected together via a flexure joint. Thus, the flexure joint allows for angular misalignment between the rotor and the stator. More specifically, the primary seal segment moves with the rotor, and the secondary seal segment is loaded by a rear wall of the stator. In an embodiment, the seal segment may be a two-piece seal including the primary seal segment and the secondary seal segment. Additionally, to minimize leakage of high-pressure fluid from a high-pressure region to a low-pressure region through the split two-piece seal, the seal assembly may further include a static seal disposed between the primary seal segment and the secondary seal segment. In such an embodiment, one or more of the primary seal segment or the secondary seal segment may include one or more grooves for receiving the static seal.

[0047] A further aspect is provided by the subject matter of the following clauses.

[0048] A turbine engine, comprising: a rotor; a stator including a rear wall; a seal assembly including a plurality of seal segments disposed between the rotor and the stator, wherein the rotor, the stator, and the seal assembly are arranged together to define a high-pressure region and a low-pressure region; and at least one biasing member engaged with one or more of the plurality of seal segments, wherein the plurality of seal segments includes a primary seal segment and a secondary seal segment connected together via a flexible joint, and wherein the flexible joint allows for angular misalignment between the primary seal segment and the secondary seal segment, thereby allowing the primary seal segment to move with the rotor while the secondary seal segment remains in contact with the rear wall of the stator.

[0049] The turbine engine according to any one of the preceding clauses, wherein the secondary seal segment is located behind the primary seal segment.

[0050] The turbine engine according to any one of the preceding clauses, wherein the primary seal segment defines an L-shaped cross-sectional shape in an axial direction of the turbine engine, and the L-shaped cross-sectional shape defines a flange.

[0051] The turbine engine according to any one of the preceding clauses, wherein the secondary seal segment defines a generally rectangular cross-sectional shape in the axial direction of the turbine engine, and the secondary seal segment is positioned on the flange of the L-shaped cross-sectional shape of the primary seal segment.

[0052] A turbine engine according to any one of the preceding clauses, wherein at least one of the primary seal section or the rotor further includes at least one aerodynamic feature to allow the formation of an air film between the plurality of seal sections and the rotor.

[0053] A turbine engine according to any one of the preceding clauses, wherein the at least one aerodynamic feature includes at least one of a Rayleigh pad, a spiral groove, or a chevron groove.

[0054] A turbine engine according to any one of the preceding clauses, further including an inner pocket located at the rear side of at least one of the plurality of seal sections, the inner pocket being surrounded by secondary seal dams on multiple sides.

[0055] A turbine engine according to any one of the preceding clauses, wherein the seal assembly further includes a static seal disposed between the primary seal section and the secondary seal section.

[0056] A turbine engine according to any one of the preceding clauses, wherein the static seal includes at least one of a spline seal, a C - seal, an E - seal, a rope seal, or a W - seal.

[0057] A turbine engine according to any one of the preceding clauses, wherein the flexible joint includes at least one of a pivot joint, a hinge joint, a flexure beam, or a pin joint.

[0058] A turbine engine according to any one of the preceding clauses, wherein the at least one biasing member connects the plurality of seal sections to the stator and retracts the seal assembly away from the rotor in the absence of pressure.

[0059] A turbine engine according to any one of the preceding clauses, wherein the at least one biasing member connects the plurality of seal sections to the stator and pushes the seal assembly towards the rotor.

[0060] A turbine engine according to any one of the preceding clauses, wherein the seal assembly further includes one or more coatings to minimize radial friction or wear.

[0061] A sealing assembly for a turbomachine having a rotor and a stator, the sealing assembly comprising: a plurality of sealing segments, wherein, when arranged together, the rotor, the stator, and the sealing assembly define a high-pressure region and a low-pressure region, the plurality of sealing segments including a primary sealing segment and a secondary sealing segment connected together via a flexible joint; and at least one biasing member that engages one or more of the plurality of sealing segments, wherein the flexible joint allows angular misalignment between the primary sealing segment and the secondary sealing segment, thereby allowing the primary sealing segment to move with the rotor while the secondary sealing segment remains in contact with the rear wall of the stator.

[0062] The sealing assembly according to any one of the preceding clauses, wherein the secondary sealing segment is located behind the primary sealing segment, and wherein the primary sealing segment defines an L-shaped cross-sectional shape in the axial direction of the turbomachine, the L-shaped cross-sectional shape defining a flange, and wherein the secondary sealing segment defines a generally rectangular cross-sectional shape in the axial direction of the turbomachine, and the secondary sealing segment is positioned on the flange of the L-shaped cross-sectional shape of the primary sealing segment.

[0063] The sealing assembly according to any one of the preceding clauses, wherein at least one of the primary sealing segment or the rotor further comprises at least one aerodynamic feature to allow the formation of an air film between the plurality of sealing segments and the rotor.

[0064] The sealing assembly according to any one of the preceding clauses, wherein the at least one aerodynamic feature comprises at least one of a Rayleigh pad, a spiral groove, or a chevron groove.

[0065] The sealing assembly according to any one of the preceding clauses, further comprising an inner pocket located at the rear side of at least one of the plurality of sealing segments, the inner pocket being surrounded by secondary sealing dams on multiple sides.

[0066] The sealing assembly according to any one of the preceding clauses, wherein the sealing assembly further comprises a static seal arranged between the primary sealing segment and the secondary sealing segment, and wherein the static seal comprises at least one of a spline seal, a C-type seal, an E-type seal, a rope seal, or a W-type seal.

[0067] The sealing assembly according to any one of the preceding clauses, wherein the flexible joint comprises at least one of a pivot joint, a hinge joint, a flexure beam, or a pin joint.

[0068] This written description uses examples to disclose the present disclosure, including the best mode, and also enables those 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. If these other examples 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, then these other examples are intended to be within the scope of the claims.

Claims

1. A turbomachine, characterized in that, Comprising: A rotor; A stator, said stator including a rear wall; A sealing assembly, said sealing assembly including a plurality of sealing segments disposed between said rotor and said stator, wherein said rotor, said stator and said sealing assembly are arranged together to define a high-pressure region and a low-pressure region; And At least one biasing member, said at least one biasing member engaging one or more of said plurality of sealing segments, Wherein said plurality of sealing segments includes a primary sealing segment and a secondary sealing segment connected together via a flexible joint, and Wherein said flexible joint allows an angular misalignment between said primary sealing segment and said secondary sealing segment, thereby allowing said primary sealing segment to move with said rotor while said secondary sealing segment remains in contact with said rear wall of said stator.

2. The turbine engine according to claim 1, characterized in that, Wherein said secondary sealing segment is located behind said primary sealing segment.

3. The turbine engine according to claim 1, characterized in that, Wherein said primary sealing segment defines an L-shaped cross-sectional shape in the axial direction of the turbine engine, said L-shaped cross-sectional shape defining a flange.

4. The turbine engine according to claim 3, characterized in that, Wherein said secondary sealing segment defines a generally rectangular cross-sectional shape in said axial direction of the turbine engine, said secondary sealing segment being positioned on said flange of said L-shaped cross-sectional shape of said primary sealing segment.

5. The turbine engine according to claim 1, characterized in that, Wherein at least one of said primary sealing segment or said rotor further includes at least one aerodynamic feature to allow the formation of an air film between said plurality of sealing segments and said rotor.

6. The turbine engine according to claim 5, characterized in that, Wherein said at least one aerodynamic feature includes at least one of a Rayleigh pad, a spiral groove or a chevron groove.

7. The turbomachine according to claim 1, characterized in that, Further including an inner pocket located at the rear side of at least one of said plurality of sealing segments, said inner pocket being surrounded by secondary sealing dams on multiple sides.

8. The turbomachine according to claim 1, characterized in that, Wherein said sealing assembly further includes a static seal disposed between said primary sealing segment and said secondary sealing segment.

9. The turbine engine according to claim 8, characterized in that, Wherein said static seal includes at least one of a spline seal, a C-shaped seal, an E-shaped seal, a rope seal or a W-shaped seal.

10. The turbomachine according to claim 1, characterized in that, Wherein said flexible joint includes at least one of a pivot joint, a hinge joint, a flexure beam or a pin joint.

Images