CN112648637A

CN112648637A - Seal assembly for a CMC liner-penetrating component

Info

Publication number: CN112648637A
Application number: CN202011070415.8A
Authority: CN
Inventors: 黄飞; D·柯特利; G·E·维赫
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 2019-10-09
Filing date: 2020-10-09
Publication date: 2021-04-13

Abstract

The invention relates to a seal assembly for a component penetrating a CMC liner. Specifically, a combustion section and a sealing system for a fuel ignition assembly of a combustion section of a gas turbine engine are provided. For example, a sealing system includes: a ferrule positioned on an outer surface of a Ceramic Matrix Composite (CMC) combustor liner; an aperture defined in the CMC liner; a sleeve positioned within the adapter of the fuel ignition assembly such that an inner end portion of the sleeve is in contact with the ferrule, the sleeve having an end wall that forms an inner boundary of a cavity defined by the sleeve; and a biasing member positioned within the cavity. The biasing member extends between the sleeve and the end wall of the sleeve. The biasing member continuously urges the sleeve into contact with the ferrule to seal the orifice to prevent fluid leakage therethrough. The exemplary sealing system may be part of a fuel ignition assembly of a combustion section of a gas turbine engine.

Description

Seal assembly for a CMC liner-penetrating component

Cross Reference to Related Applications

This application is a partial continuation of U.S. application serial No. 15/448938 filed on 3/2017 and is claimed to have priority, the contents of which are incorporated herein by reference.

Technical Field

The present subject matter relates generally to combustion assemblies for gas turbine engines. More specifically, the present subject matter relates to a seal assembly for sealing around a component that penetrates a combustor liner of a gas turbine engine combustion assembly, and most specifically penetrates a ceramic matrix composite combustor liner.

Background

Gas turbine engines generally include a fan and a core arranged in flow communication with each other. Further, the core of a gas turbine engine generally includes, in serial-flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section, where one or more axial compressors progressively compress the air until it reaches the combustion section, which includes a combustor defining a combustion chamber. Fuel is mixed with the compressed air and combusted within the combustion chamber to provide combustion gases. The combustion gases are channeled from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to the atmosphere.

The combustion section generally includes an annular inner liner, an annular outer liner radially spaced from the inner liner, and a combustor dome coupled to an upstream or forward end of the inner and outer liners. A fuel injector or nozzle extends through the dome and is configured to provide a fuel/air mixture to a combustion chamber defined between an inner liner and an outer liner. An outer casing or combustor casing circumferentially surrounds the outer liner and at least partially defines an outer plenum or passage between the combustor casing and the outer liner.

The combustion section also includes an ignition system having one or more igniter assemblies mounted or coupled to the housing. The igniter portion of the igniter assembly extends generally radially through the housing and the outer chamber. The firing tip portion of the igniter extends at least partially through an opening defined in the outer liner, and a ferrule or other sealing member extends around the igniter adjacent the opening to provide a seal to prevent fluid leakage through the opening. During operation of the gas turbine, such as during start-up or restart, the igniter may be energized to provide a spark at the ignition tip in order to ignite the fuel/air mixture within the combustion chamber.

More commonly, non-traditional high temperature materials, such as Ceramic Matrix Composite (CMC) materials, are being used in gas turbine applications. Components made from such materials have higher temperature capabilities than typical components (e.g., metal components), which may allow for improved component performance and/or increased engine temperatures. Thus, the use of CMC materials for the inner and outer liners of a combustor may improve the durability of the liners and allow for reduced impingement cooling or other types of cooling of the liners and increased combustion temperatures, which may improve engine performance. However, CMC materials typically have a much lower coefficient of thermal expansion than, for example, metals or metal alloys, such that CMC components have a much lower thermal growth rate than metal components.

Thus, for CMC combustor liners, the radial and/or axial positioning of the igniter assembly relative to the outer liner and/or the combustion chamber may change during gas turbine operation. For example, thermal growth rate variations of the outer shell and the CMC outer liner may cause the seal components to shift in position near the liner opening, which may result in undesirable fluid leakage through the opening, such as from the relatively cold side of the liner to the relatively hot combustion chamber. Accordingly, an improved ignition assembly for a gas turbine engine and an improved sealing system for an ignition assembly would be useful in the turbofan engine industry.

Disclosure of Invention

Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention.

In one exemplary embodiment of the present subject matter, a sealing system for a fuel ignition assembly of a gas turbine engine is provided. The fuel ignition assembly includes an igniter tube having a tip portion located adjacent to a combustor of the gas turbine engine. The sealing system includes a ferrule positioned on an exterior surface of a Ceramic Matrix Composite (CMC) liner of a combustor, proximate an aperture defined in the CMC liner. The sealing system also includes a sleeve positioned within the adapter of the fuel ignition assembly such that an inner end portion of the sleeve is in contact with the ferrule. The sleeve has an end wall forming an inner boundary of a cavity defined by the sleeve, and the adapter supports the igniter tube. The sealing system also includes a biasing member positioned within the cavity. The biasing member extends between the sleeve and the end wall of the sleeve. The sleeve is received through an outer end of an adapter opening defined by the adapter and has a shoulder extending around a periphery of the sleeve and abutting the outer end of the adapter. The biasing member continuously urges the sleeve into contact with the ferrule to seal the orifice to prevent fluid leakage through the orifice.

In another exemplary embodiment of the present subject matter, a combustion section of a gas turbine engine is provided. The combustion section includes an inner liner and an outer liner radially spaced from the inner liner, the outer liner defining an aperture therein; a combustion chamber defined between the inner and outer liners; a combustor casing extending circumferentially around the outer liner; and a fuel ignition assembly. The outer liner and the combustor case define an outer flow path therebetween, and the combustor case includes an aperture substantially aligned with the aperture of the outer liner. The fuel ignition assembly includes an igniter tube having a tip portion received in an aperture defined in an outer liner. The fuel igniter assembly also includes an adapter for supporting the igniter tube relative to the burner housing. The adapter defines an adapter opening for receiving the igniter tube, and the adapter opening has an inner end radially opposite the outer end. The fuel ignition assembly also includes a sleeve received through the outer end of the adapter opening. The sleeve has a shoulder extending around an outer periphery of the sleeve, and the shoulder abuts an outer end of the adapter. Additionally, the fuel ignition assembly includes a collar positioned on an outer surface of the outer liner, adjacent the aperture in the outer liner; a sleeve positioned within the adapter such that an inner end portion of the sleeve is in contact with the ferrule, the sleeve having an end wall forming an inner boundary of a cavity defined by the sleeve; and a biasing member positioned within the cavity, the biasing member extending between the sleeve and the end wall of the sleeve. The biasing member continuously urges the sleeve into contact with the ferrule to seal the orifice to prevent fluid leakage through the orifice.

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

Technical solution 1. a sealing system for a fuel ignition assembly of a gas turbine engine, the fuel ignition assembly including an igniter tube having a tip portion located adjacent a combustor of the gas turbine engine, the sealing system comprising:

a ferrule positioned on an exterior surface of a Ceramic Matrix Composite (CMC) liner of the combustor, the ferrule located proximate an aperture defined in the CMC liner;

an adapter supporting the igniter tube;

a sleeve having a sleeve shoulder extending around an outer periphery of the sleeve, the sleeve shoulder abutting an outer end of the adapter, the sleeve defining an opening therethrough and including a protrusion extending inwardly into the opening, the sleeve being received through the outer end of the adapter opening defined by the adapter;

a sleeve positioned within the adapter such that an inner end portion of the sleeve is in contact with the ferrule, the sleeve having an end wall that forms an inner boundary of a cavity defined by the sleeve; and

a biasing member positioned within the cavity, the biasing member extending between the sleeve and an end wall of the sleeve,

wherein the igniter tube includes an igniter tube shoulder that abuts the protrusion, and

wherein the biasing member continuously urges the sleeve into contact with the ferrule to seal the aperture to prevent leakage of fluid through the aperture.

Claim 2. the sealing system of any preceding claim, wherein at least one gasket is positioned between the igniter tube shoulder and the projection of the sleeve.

Claim 3. the sealing system of any preceding claim, wherein the igniter tube shoulder and the projection of the sleeve are each defined at a radial position to control a position of a tip portion of the igniter tube relative to an outer liner of a combustor.

Claim 4. the sealing system of any preceding claim, wherein at least one gasket is positioned between the sleeve shoulder and the outer end of the adapter.

Claim 5. the sealing system of any preceding claim, wherein a plate extends between the inner surface of the sleeve and the outer end of the biasing member.

Claim 6. the sealing system of any preceding claim, wherein the adapter defines a radial stop that limits radial movement of the sleeve within the adapter.

Claim 7. the sealing system of any preceding claim, wherein the sleeve defines a collar around an outer end portion of the sleeve, and wherein the collar abuts the radial stop when the sleeve is at a maximum radially inward position.

The seal system of any preceding claim, wherein the adapter defines a flange to support the adapter relative to the combustor case, and wherein a seal extends between the combustor case and the flange.

Claim 9. the sealing system of any preceding claim, wherein at least a portion of an outer surface of the inner end portion of the sleeve is spherical or arcuate.

Claim 10 the sealing system of any preceding claim, wherein the ferrule defines a pocket in which the inner end portion of the sleeve is received, and wherein an inner surface of the pocket is complementary in shape to an outer surface of the inner end portion of the sleeve.

Claim 11 the sealing system of any preceding claim, wherein the combustor casing is formed from a metal or metal alloy such that the combustor casing and the outer liner have different coefficients of thermal expansion.

Claim 12 the sealing system of any preceding claim, wherein the biasing member is a coil spring.

The invention of claim 13 is a combustion section of a gas turbine engine, comprising:

an inner liner and an outer liner radially spaced from the inner liner, the outer liner defining an aperture therein;

a combustion chamber defined between the inner and outer liners;

a combustor casing extending circumferentially around the outer liner, the outer liner and the combustor casing defining an outer flow passage therebetween, the combustor casing including an aperture substantially aligned with the aperture of the outer liner; and

a fuel ignition assembly, comprising:

an igniter tube having a tip portion received in an aperture defined in the outer liner;

an adapter for supporting the igniter tube relative to the burner housing, the adapter defining an adapter opening for receiving the igniter tube, the adapter opening having an inner end radially opposite an outer end;

a sleeve received through an outer end of the adapter opening, the sleeve having a sleeve shoulder extending around an outer circumference of the sleeve, the sleeve shoulder abutting the outer end of the adapter;

a ferrule positioned on an outer surface of the outer liner proximate an aperture in the outer liner;

wherein a position of the igniter tube shoulder is radially adjustable to control a position of the tip portion relative to an aperture defined in the outer liner, an

The burner section of any preceding claim, wherein the sleeve defines an opening therethrough and includes a protrusion extending inwardly into the opening, and wherein the igniter tube includes an igniter tube shoulder abutting the protrusion.

Claim 15 the combustion section of any preceding claim, wherein at least one gasket is positioned between the igniter tube shoulder and the projection of the sleeve.

The combustion section of any preceding claim, wherein at least one spacer is positioned between the shoulder of the sleeve and the outer end of the adapter.

Claim 17 the combustion section of any preceding claim, wherein a plate extends between the inner surface of the sleeve and the outer end of the biasing member.

The combustion section of any preceding claim, further comprising a heat shield positioned between the outer liner and the combustor casing.

The combustion section of any preceding claim, wherein at least the outer liner is formed from a ceramic matrix composite material.

Solution 20. the combustion section of any preceding solution, wherein at least a portion of an outer surface of the inner end portion of the sleeve is spherical or arcuate, wherein the ferrule defines a pocket in which the inner end portion of the sleeve is received, and wherein an inner surface of the pocket is complementary in shape to the outer surface of the inner end portion of the sleeve.

Drawings

A full 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 appended figures, in which:

FIG. 1 provides a schematic cross-sectional view of an exemplary gas turbine engine, according to various embodiments of the present subject matter.

FIG. 2 provides a cross-sectional side view of a portion of the combustion section of the gas turbine engine, as shown in FIG. 1, in accordance with various embodiments of the present subject matter.

FIG. 3 is an enlarged cross-sectional view of a portion of the combustion section as shown in FIG. 2, including a fuel ignition assembly according to an exemplary embodiment of the present subject matter.

FIG. 4 is an enlarged cross-sectional view of a portion of a combustion section including a portion of the fuel ignition assembly as shown in FIG. 3 in accordance with at least one embodiment of the present subject matter.

Detailed Description

Reference now will be made in detail to embodiments of the invention, one or more examples of which are illustrated in the drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the invention. The terms "first," "second," and "third" as used herein may be used interchangeably to distinguish one element from another and are not intended to denote the position or importance of an individual element. The terms "upstream" and "downstream" refer to relative directions with respect to fluid flow in a fluid pathway. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction to which the fluid flows.

Referring now to the drawings, in which like numerals represent like elements throughout the several views, FIG. 1 is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure. More specifically, for the embodiment of FIG. 1, the gas turbine engine is a high bypass turbofan jet engine 10, referred to herein as "turbofan engine 10". As shown in fig. 1, the turbofan engine 10 defines an axial direction a (extending parallel to a longitudinal centerline 12 provided for reference) and a radial direction R. Generally, turbofan 10 includes a fan section 14 and a core turbine engine 16 disposed downstream of fan section 14.

The depicted exemplary core turbine engine 16 generally includes a substantially tubular casing 18 defining an annular inlet 20. The casing 18 encloses, in series flow relationship, a compressor section including a booster or Low Pressure (LP) compressor 22 and a High Pressure (HP) compressor 24; a combustion section 26; a turbine section including a High Pressure (HP) turbine 28 and a Low Pressure (LP) turbine 30; and a jet 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.

For the illustrated embodiment, fan section 14 includes a fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced apart manner. As shown, fan blades 40 extend outwardly from disk 42 in a generally radial direction R. Fan blades 40 and disk 42 are rotatable together about longitudinal axis 12 by LP shaft 36. In some embodiments, a power gearbox having multiple gears may be included to step down the rotational speed of LP shaft 36 to a more efficient fan speed.

Still referring to the exemplary embodiment of FIG. 1, disk 42 is covered by a rotatable forward nacelle 48 that is aerodynamically contoured to promote airflow across the plurality of fan blades 40. Additionally, exemplary fan section 14 includes an annular fan case or outer nacelle 50 that circumferentially surrounds at least a portion of fan 38 and/or core turbine engine 16. It should be appreciated that the nacelle 50 may be configured to be supported relative to the core turbine engine 16 by a plurality of circumferentially spaced outlet vanes 52. Moreover, a downstream section 54 of nacelle 50 may extend above an exterior of core turbine engine 16 so as to define a bypass airflow passage 56 therebetween.

During operation of the turbofan engine 10, a volume of air 58 enters the turbofan 10 via the nacelle 50 and/or an associated inlet 60 of the fan section 14. As the volume of air 58 passes by fan blades 40, a first portion of air 58, as indicated by arrow 62, is channeled or conveyed into bypass airflow path 56, and a second portion of air 58, as indicated by arrow 64, is channeled or conveyed into LP compressor 22. The ratio between the first portion 62 of air and the second portion 64 of air is commonly referred to as the bypass ratio. The pressure of the second portion 64 of the air then increases as it passes 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.

The combustion gases 66 are channeled through HP turbine 28 wherein a portion of thermal and/or kinetic energy from combustion gases 66 is extracted via successive stages of HP turbine stator vanes 68 coupled to casing 18 and HP turbine rotor blades 70 coupled to HP shaft or spool 34, thereby causing HP shaft or spool 34 to rotate, thereby supporting operation of HP compressor 24. The combustion gases 66 are then passed 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 or spool 36, thus causing the LP shaft or spool 36 to rotate, thereby supporting operation of the LP compressor 22 and/or rotation of the fan 38.

The combustion gases 66 are then passed through the jet exhaust nozzle section 32 of the core turbine engine 16 to provide propulsive thrust. At the same time, the pressure of the first portion of air 62 substantially increases as the first portion of air 62 passes through the bypass airflow passage 56 before it is exhausted from the fan nozzle exhaust section 76 of the turbofan 10, also providing propulsive thrust. HP turbine 28, LP turbine 30, and jet exhaust nozzle section 32 at least partially define a hot gas path 78 to route combustion gases 66 through core turbine engine 16.

In some embodiments, components of the turbofan engine 10, particularly components within the hot gas path 78, may include a Ceramic Matrix Composite (CMC) material that is a non-metallic material having high temperature capabilities. Exemplary CMC materials for such components may include silicon carbide (SiC), silicon, silica, or alumina matrix materials, and combinations thereof. Ceramic fibers may be embedded within a matrix, such as, oxidatively stable reinforcing fibers comprising monofilaments such as sapphire or silicon carbide (e.g., SCS-6 of Textron), as well as silica and yarns comprising silicon carbide (e.g., NICATON of Nippon Carbon, TYLRANNO of Ube Industries, and SYLRAMIC of Dow Corning), aluminum silicate (e.g., 440 and 480 of Nextel), and chopped whiskers and fibers (e.g., 440 and SAFFIL of Nextel), and optionally ceramic particles (e.g., oxides of Si, Al, Zr, Y, and combinations thereof), and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite). For example, in certain embodiments, the fiber bundles, which may include a ceramic refractory coating, are formed into reinforcing strips, such as unidirectional reinforcing strips. Multiple tapes may be stacked together (e.g., as plies) to form a prefabricated component. The fiber bundle may be impregnated with the slurry composite either before or after forming the preform. The preform may then be subjected to a thermal treatment, such as solidification or burnout, to produce a high coke residue in the preform, and subsequent chemical treatment, such as melt infiltration with silicon, to arrive at a component formed of the CMC material having the desired chemical composition. In other embodiments, the CMC material may be formed as, for example, a carbon fiber cloth rather than a tape.

More specifically, a method for forming a CMC component (such as a CMC outer liner of a combustor, as described below) may first include: a plurality of plies of CMC material are stacked to form a CMC preform having a desired shape or profile. It will be appreciated that the plurality of CMC plies forming the preform may be layered on a layering tool, a die, a mandrel, or another suitable device for supporting the plies and/or for defining the desired shape. The desired shape of the CMC preform may be a desired shape or profile of the resulting CMC component (e.g., annular CMC outer liner).

After the multiple plies are stacked to form the preform, the preform may be processed, e.g., consolidated and cured, in an autoclave. After processing, the preform forms a CMC component in an unprocessed state, such as an outer CMC liner in an unprocessed state. The CMC component in the green state is a single-piece component, i.e., multiple plies of the cured preform join the plies to produce the CMC component formed from a continuous piece of CMC material in the green state. The green state component may then undergo firing (or burn-out) and densification to produce a dense CMC component. For example, the component in the green state may be placed in a furnace to burn out any mandrel-forming materials and/or solvents used in forming the CMC plies, and to decompose the binder in the solvent, and then placed in the furnace with the silicon to convert the ceramic matrix precursors of the plies into the ceramic material of the matrix of the CMC component. The silicon melts and penetrates any porosity with the matrix due to binder decomposition during burn-up/firing; the melt infiltration of silicon into the CMC component densifies the CMC component. However, densification may be performed using any known densification technique, including but not limited to, Silcomp, Melt Infiltration (MI), Chemical Vapor Infiltration (CVI), Polymer Infiltration and Pyrolysis (PIP), and oxide/oxidation processes. In one embodiment, densification and firing may be performed in a vacuum furnace or an inert atmosphere with an established atmosphere at a temperature above 1200 ℃ to allow silicon or another suitable material or materials to melt infiltrate the component. Optionally, after firing and densification, the CMC component may be finished, if desired and as needed, and/or coated with one or more coatings, such as an Environmental Barrier Coating (EBC) or a Thermal Barrier Coating (TBC).

The foregoing methods of forming CMC components, such as CMC outer liners, are provided by way of example only. For example, other known methods or techniques may be used to compact and/or cure the CMC plies, as well as densify the CMC component in the green state. Alternatively, any combination of these or other known processes may be used.

As described above, components comprising CMC materials may be used within the hot gas path 78, such as within the combustion section and/or the turbine section of the engine 10. However, CMC components may also be used in other sections, such as compressor and/or fan sections. As particular examples described in more detail below, the outer liner of the combustor of the combustion section 26 may be formed of a CMC material, for example, to provide greater temperature capability of the combustor, to better protect the turbine casing from the combustion gas temperatures, and/or to reduce the amount of cooling fluid supplied to the outer liner.

FIG. 2 is a cross-sectional side view of a portion of combustion section 26. As shown in FIG. 2, the combustor section 26 generally includes an annular-type combustor 80 having an annular inner liner 82, an annular outer liner 84, and a dome end 86 extending between an upstream end 88 of the inner liner 82 and an upstream end 90 of the outer liner 84. Inner liner 82 is radially spaced from outer liner 84 and defines a generally annular combustion chamber 92 therebetween. As previously mentioned, the outer liner 84 is preferably formed of a CMC material; the inner liner 82 may also be formed from a CMC material.

The inner and

outer liners

82, 84 are encapsulated within a combustor or casing 94, i.e., the combustor casing 94 extends circumferentially around the outer liner 84. The heat shield 96 is positioned against the inner surface 94a of the combustor casing 94, for example, to help prevent the temperature of the hot gases within the combustion zone 26 from causing creep in the combustor casing 94. Heat shield 96 is formed from any suitable material; in one embodiment, the heat shield 96 is formed from a high temperature metal having a honeycomb pattern formed therein. Further, an outer flow path 98 may be defined between the combustor casing 94 and the outer liner 84. The inner and

outer liners

82, 84 extend from the dome end 86 toward the turbine nozzle 100. Further, a fuel injector or nozzle 102 extends at least partially through the dome end 86 and provides a fuel-air mixture 104 to the combustion chamber 92.

In various embodiments, as shown in FIG. 2, the combustion section 26 includes a fuel ignition system 200 for igniting the fuel-air mixture 102 within the combustion chamber 92. The fuel ignition system 200 generally includes at least one fuel igniter assembly 202 electrically/electronically coupled to a controller or ignition source 204. The ignition assembly 202 may be attached to the outer surface 94b of the burner housing 94. Fuel ignition assembly 202 includes a sealing system 206 that, for example, prevents fluid from outer flow path 98 (i.e., the cold side of outer liner 84) from flowing into combustion chamber 92 (i.e., the hot side of outer liner 84).

FIG. 3 provides an enlarged cross-sectional view of a portion of the combustion section 26 including the fuel ignition assembly 202 as shown in FIG. 2, according to an exemplary embodiment of the present subject matter. In the exemplary embodiment shown in fig. 3, the fuel ignition assembly 202 includes an igniter tube 208 that extends generally radially through the aperture 106 defined by the burner housing 94. A firing tip or tip portion 210 of the igniter tube 208 extends at least partially through the aperture 108 defined within the outer liner 84 such that the tip portion 210 is located adjacent to the combustor 80, as shown in FIG. 2. In a particular embodiment, the tip portion 210 may be concentrically aligned with respect to the aperture 108 and with respect to a longitudinal centerline CL of the igniter tube 208. As shown in fig. 3, the outer bushing bore 108 may be sized (i.e., may have sufficient cross-sectional area) to allow some axial and/or circumferential movement of the igniter tip portion 210 within the bore 108.

In the exemplary embodiment shown in fig. 3, the fuel ignition assembly 202 includes an outer housing or adapter 212 that facilitates supporting the igniter tube 208, as well as other components of the fuel ignition assembly 202 positioned within the adapter 212 as further described herein with reference to the burner housing 94. The adapter 212 has an interior 214 and an exterior 216. The opening 218 extends through the adapter 212; an inner end 218a of the opening 218 is defined in the inner portion 214, and an outer end 218b of the opening 218 is defined in the outer portion 216, such that the outer end 218b is diametrically opposed to the inner end 218 a. The opening 218 may be sized and/or shaped for receiving one or more components of the fuel ignition assembly 202, each component surrounding the igniter tube 208, as described in more detail below. For example, the opening 218 may be substantially cylindrical or have any other suitable shape. As shown in fig. 3, a portion of the igniter tube 208 may extend through the opening 218 and radially outward from the opening 218. Further, the adapter 212 extends radially through the aperture 106 in the combustor casing 94 and includes a flange 213 that supports the adapter 212 relative to the combustor casing 94. Moreover, the adapter 212 may be configured to be coupled to the combustor casing 94. For example, the adapter 212 may be coupled to the combustor casing 94 at the adapter flange 213 using one or more bolts, screws, or other suitable attachment or fastening mechanisms. A seal 220 is positioned between the adapter flange 213 and the combustor casing 94 to help prevent fluid from leaking through the apertures 106. It will be appreciated that the seal 220 may be a high temperature, high pressure seal suitable for use in the combustion section 26 of the gas turbine engine 10. Further, the adapter interior 214 extends through the aperture 110 in the heat shield 96 and toward the outer liner 84.

Still referring to fig. 3, the fuel ignition assembly 202 also includes a sleeve 222 that supports the igniter tube 208. More specifically, an inner end 222a of the sleeve 222 is received through an outer end 218b of the opening 218 in the adapter 218. The sleeve 222 includes a shoulder 224 between the sleeve inner end 222a and the outer end 222 b. Shoulder 224 extends around the outer circumference of the sleeve; the shoulder 224 abuts or abuts the exterior 216 of the adapter 212 such that the outer end 222b of the sleeve 222 extends radially outward from the adapter exterior 216. As shown in fig. 3, one or more shims 226 may be positioned between the shoulder 224 and the adapter exterior 216, for example, to control the radial position of the igniter tip portion 210.

Similar to the opening 218 through the adapter 212, the opening 228 extends through the sleeve 222 and may be sized and/or shaped to receive the igniter tube 208. The opening 228 may be substantially cylindrical or have any other suitable shape. The protrusion 230 extends into the opening 228 through the sleeve 222, and the shoulder 232 of the igniter tube 208 rests on the protrusion 230. Thus, movement of the sleeve 222 generally parallel to the centerline CL of the igniter tube 208 is also sufficient to move the igniter tube 208. For example, replacing the spacer 226 with a thicker or thinner spacer, using more than one spacer 226, or removing the spacers 226 altogether, repositions the sleeve 222 relative to the adapter 212 and thus the combustor casing 94 and the outer liner 84, which likewise repositions the igniter tube 208 relative to the adapter 212, the combustor casing 94, and the outer liner 84. Thus, the position of the sleeve 222 is substantially adjustable in the radial direction R (i.e., the sleeve position is substantially radially adjustable), and thus, by allowing for substantially radial adjustment of the tip portion 210, control of the position of the igniter tip portion 210 is facilitated.

Further, it will be understood that the igniter tube shoulder 232 may be formed as part of the igniter tube 208 or may be fixedly attached to the igniter tube 208. For example, the shoulder 232 may be formed by an outer housing, nut, washer, or the like that is fixedly attached to the igniter tube 208. In some embodiments, one or more shims 226 may be included between the igniter tube shoulder 232 and the protrusion 230, which may have a variable thickness as described herein, to reposition the igniter tube 208 relative to the sleeve 222 and the adapter 212, which repositions the igniter tube 208 and its tip 210 relative to the combustor casing 94 and the outer liner 84. In other embodiments, the shoulder 232 may be thicker or thinner, or otherwise extend a longer or shorter distance along the centerline CL, to change the position of the igniter tube 208 relative to the sleeve 222, adapter 212, combustor casing 94, and outer liner 84. Accordingly, the position of the igniter tube 208 is substantially adjustable in the radial direction R (i.e., the igniter tube position is substantially radially adjustable), which may help control the position of the igniter tip portion 210 by allowing the tip portion 210 to be substantially radially adjusted.

The location or position of the igniter tip portion 210 may also be controlled and/or adjusted in other ways. For example, a thicker or thinner plate 238 positioned between the sleeve inner surface 222c and the outer end 236b of the biasing member 236 in fig. 3 and discussed in more detail below, the use of multiple plates 238 between the sleeve 222 and the biasing member 236, or the omission of plates 238 may be used to radially change the position of the igniter tube 208 and, thus, the igniter tip portion 210. In other embodiments, a thicker or thinner seal 220, the use of multiple seals 220 between the adapter 212 and the burner housing 94, or the omission of seals 220 may be used to alter the position of the igniter tube 208 and, thus, the igniter tip portion 210. Other radial variations in the igniter assembly 202, such as a combination of two or more ways of the radial position of the protrusion 230 within the sleeve 222, the radial height of the adapter 212 (upon which the sleeve shoulder 224 rests), and/or for changing the radial position of the igniter tube 208, may also be used to control and/or adjust the positioning or position of the igniter tip portion 210.

Continuing with fig. 3, a generally annular sleeve 234 is positioned within the adapter 212. More specifically, the sleeve 234 extends within the opening 218 in the adapter 212, and an inner end portion 234a of the sleeve 234 extends through the inner end 218a of the opening 218 and toward the outer bushing 84. An outer end portion 234b of the sleeve 234 is positioned adjacent the sleeve inner end 222a within the adapter opening 218. As shown in fig. 3, the sleeve 234 is free to move along the adapter opening 218, but the adapter 212 may define a lip or radial stop 219 to limit the radially inward movement of the sleeve 234. That is, the sleeve 234 defines a collar 233 about its outer end portion 234b that snaps over or abuts the radial stop 219 when the sleeve 234 is in the maximum radially inward position. The radial stop 219 may be defined within the sleeve 234 such that when the adapter 212 is assembled with the combustor casing, the radial stop is radially outward from or above the combustor casing 94.

A biasing member 236, such as a spring or the like, is positioned within the sleeve 234 and extends circumferentially around a portion of the igniter tube 208. Biasing member 236 is disposed between end wall 234c of sleeve 234 and inner surface 222c of sleeve 222. End wall 234c forms an inner boundary of a cavity 235 defined by sleeve 234, and a biasing member 236 is positioned within cavity 235 such that biasing member 236 extends between end wall 234c and sleeve 222. In one embodiment, the biasing member 236 may be a coil spring, and in another embodiment, the biasing member 236 may be a wave spring. The sleeve 234 surrounding the biasing member 236 helps prevent binding or bending of the biasing member 236, for example, under operating conditions of the engine 10.

Further, in some embodiments, plate 238 extends between sleeve inner surface 222c and outer end 236b of biasing member 236 such that biasing member 236 contacts plate 238 rather than inner surface 222c of sleeve 222. Thus, the plate 238 protects the sleeve 222 from wear that would otherwise occur due to contact by the biasing member. It will be appreciated that the plate 238 preferably has a sufficient cross-sectional area to prevent any portion of the biasing member 236 from contacting the bushing 222. Further, in addition to controlling the position of the igniter tip portion 210, one or more shims 226 may be used to control the working height or length of the biasing member 236, which may affect the load provided by the biasing member against the end wall 234 c.

The biasing member 236 generally provides a radially inward force relative to the sleeve 234 (i.e., the end wall 234 c) to bias or continuously urge the sleeve inner end portion 234a relative to the ferrule 240 positioned adjacent the outer bushing aperture 108 and thereby seat the inner end portion 234a relative to the ferrule 240. As described in more detail below, the collar 240 provides a seal around the igniter tube 208 and the outer liner bore 108 to, for example, prevent fluid from leaking from the outer flow path 98 into the hot gas path 78, i.e., from the cold side to the hot side of the outer liner 84. By biasing or urging the sleeve 234 relative to the collar 240, the biasing member 236 helps maintain the seal provided by the collar despite relative axial and circumferential movement between the collar 240 and the outer liner 84 and relative radial movement between the outer liner 84 and the combustor casing 94.

FIG. 4 is an enlarged cross-sectional view of a portion of the combustion section 26 as shown in FIG. 3 and at least partially includes a portion of the outer liner 64, the tip portion 210 of the igniter tube 208, the inner end portion 234a of the sleeve 234, and the ferrule 240. In the exemplary embodiment shown in fig. 4, the collar 240 is generally concentrically aligned with the opening 108 in the outer bushing 84, but as described above, the collar 240 may move axially and/or circumferentially relative to the opening 108. For example, during operation of the gas turbine engine 10, the collar 240 and/or the outer liner 84 may move axially and/or circumferentially relative to the other of the collar and the outer liner. The collar 240 may have a shape and/or size suitable to ensure that a collar 240a of the collar 240 extending around the collar 240 still surrounds the opening 108 despite any relative axial and/or circumferential movement between the collar 240 and the outer bushing 84.

As further shown in fig. 4, the ferrule 240 defines a pocket 242, and the inner end portion 234a of the sleeve 234 is formed and/or shaped to fit within the pocket 242. The collar 240 may include a lip 244 or other similar feature or device for locking or retaining the sleeve inner end portion 234a within the pocket 242. For example, in other embodiments, the collar 240 may define one or more tabs 244 to retain the sleeve inner end portion 234a within the pocket 242. In particular embodiments, at least a portion of an outer surface 234d of sleeve inner end portion 234a may be shaped or formed to complement inner surface 242a of pocket 242. For example, in one embodiment, a portion of the sleeve outer surface 234d and a portion of the pocket inner surface 242a may be spherical and/or arcuate to form a ball and socket type joint therebetween, thereby allowing relative movement between the outer liner 84 and the collar 240 and between the outer liner 84 and the combustor casing 94 during operation of the gas turbine engine 10. That is, during engine operation, the outer liner 84 may move radially and/or axially relative to the combustor casing 94; this relative movement may be caused by a number of factors, including varying rates of thermal growth between the outer liner 84 and the combustor casing 94, and/or gravitational forces on the gas turbine engine 10, such as takeoff, landing or general maneuvering of the aircraft to which the engine 10 is attached. Further, the collar 240 may move axially and/or circumferentially relative to the outer liner 84, for example, due to the aforementioned reasons of combustion dynamics and relative movement between the outer liner 84 and the combustor casing 94. By locking or retaining the sleeve inner end portion 234a in the collar pocket 242, contact between the sleeve 234 and the collar 240 may be maintained because the sleeve 234 will travel radially, axially, and/or circumferentially with the collar 240 as the sleeve 234 moves relative to the outer liner 84, and as the outer liner 84 moves relative to the combustor casing 94. As a result, a seal between the ferrule 240 and the igniter tube 208 and the outer liner bore 108 may be maintained to prevent fluid leakage between the ferrule 240 and the igniter tube 208 and between the ferrule 240 and the bore 108.

Further, the sleeve 234 transfers a generally uniform load from the biasing member 236 to the ferrule 240, e.g., the biasing member 236 presses against the sleeve end wall 234c, which in turn transfers the load from the biasing member 236 to the sleeve inner end portion 234a, and thus to the ferrule 240, in a generally uniform manner. Thus, the biasing member 236 helps ensure a generally consistent contact between the sleeve 234 and the ferrule 240, which helps ensure a good seal between the ferrule 240 and the outer bushing 84 and the ferrule 240 and the igniter tube 208. Further, the biasing member 236 is sized and/or selected to provide sufficient load to the ferrule 240 via the sleeve 234 at any engine cycle temperature or combustion dynamics. More specifically, during operation of engine 10, biasing member 236 may be exposed to relatively high temperatures, such as in excess of about 1300 ° f. Therefore, an appropriate biasing member 236 must be selected to apply sufficient load to the ferrule 240 over a temperature range that includes such relatively high temperatures. Additionally, the combustion dynamics of engine 10 may include vibrations within combustor 80, which may also cause outer liner 84 to vibrate. Thus, an appropriate biasing member 236 must be selected to apply sufficient load to the ferrule 240 to maintain the ferrule in contact with the outer bushing 84 even when the outer bushing 84 vibrates or moves. Further, each of the sleeve 234 and the ferrule 240 is preferably a lightweight member, e.g., formed of a lightweight material or as neatly formed as possible, to help reduce dynamic loads of the moving parts. Lightweight components may be, for example, those that maintain combustor dynamics within an acceptable range and thus do not push combustor dynamics outside of the acceptable range.

Further, as previously mentioned, in particular embodiments, the outer liner 84 is formed from a CMC material, and thus may be referred to as a CMC outer liner 84. However, the burner housing 94 may be formed from a different material, such as a metal or metal alloy. Thus, the CMC outer liner 84 and the combustor casing 94 may have different coefficients of thermal expansion or different rates of thermal growth, and in embodiments where the combustor casing 94 is formed of a metal or metal alloy material, the combustor casing 94 may thermally expand faster than the CMC outer liner 84 or at a greater rate than the CMC outer liner 84. Due to the different thermal growth rates, the combustor casing 94 may move radially relative to the outer liner 84.

As described above and shown in fig. 3, the burner housing 94 supports the adapter 212, the sleeve 234 and the biasing member 236 are received in the adapter, and the adapter supports the igniter tube 208 and the sleeve 222. Thus, radial movement of the burner housing 94 causes radial movement of the adapter 212 and, thus, the igniter tube 208 and the sleeve 222. Thus, the biasing member 236 must supply sufficient load to maintain the inner end portion 234a of the sleeve 234 in contact with the ferrule 240 despite any radial movement of the adapter 212, the igniter tube 208, and the sleeve 222, and thus maintain the seal between the ferrule 240, the outer liner 84, and the igniter tube 208.

The embodiments as described herein and as shown in fig. 3 and 4 provide various improvements and/or technical advantages over known or existing spark ignition systems and sealing features for such systems. For example, the biasing member 236 maintains the inner end portion 234a of the sleeve 234 in contact with the collar 240 as the outer liner 84 moves relative to the combustor casing 94 and as the collar 240 moves relative to the outer liner 84, i.e., across various static and dynamic loads and despite different thermal growth rates between components. Maintaining a suitable seal between the cold side and the hot side of the outer liner 84 may, for example, improve engine performance, etc.

Additionally or alternatively, the sleeve 234 facilitates preventing binding or buckling of the biasing component 236 due to relatively high temperatures and/or due to radial and/or axial growth differences between the CMC outer liner 84 and the combustor casing 94. Additionally or alternatively, the use of one or more shims 226 minimizes tolerance stack-up issues of biasing member 236 and sleeve 234, and helps eliminate sticking issues with respect to biasing member 236. Other improvements and/or technical advantages may also be realized from the embodiments described herein.

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

Claims

1. A sealing system for a fuel ignition assembly of a gas turbine engine, the fuel ignition assembly including an igniter tube having a tip portion located near a combustor of the gas turbine engine, the sealing system comprising:

an adapter supporting the igniter tube;

2. The sealing system of claim 1, wherein at least one gasket is positioned between the igniter tube shoulder and the sleeve protrusion.

3. The sealing system of claim 1, wherein the igniter tube shoulder and the projection of the sleeve are each defined at a radial position to control a position of a tip portion of the igniter tube relative to an outer liner of a combustor.

4. The sealing system of claim 1, wherein at least one gasket is positioned between the sleeve shoulder and the outer end of the adapter.

5. The sealing system of claim 1, wherein a plate extends between an inner surface of the sleeve and an outer end of the biasing member.

6. The sealing system of claim 1, wherein the adapter defines a radial stop that limits radial movement of the sleeve within the adapter.

7. The sealing system of claim 6, wherein the sleeve defines a collar around an outer end portion of the sleeve, and wherein the collar abuts the radial stop when the sleeve is in a maximum radially inward position.

8. The sealing system of claim 1, wherein the adapter defines a flange to support the adapter relative to the combustor case, and wherein a seal extends between the combustor case and the flange.

9. The sealing system of claim 1, wherein at least a portion of an outer surface of the inner end portion of the sleeve is spherical or arcuate.

10. The sealing system of claim 9, wherein the ferrule defines a pocket in which the inner end portion of the sleeve is received, and wherein an inner surface of the pocket is complementary in shape to an outer surface of the inner end portion of the sleeve.