SEAL SUPPORT ASSEMBLY FOR A TURBINE ENGINE

A turbine engine is provided. The turbine engine includes: a rotor; a stator having a carrier; a seal support assembly coupled to the carrier; and a seal assembly disposed between the rotor and the stator and supported by the seal support assembly, the seal assembly defining a high pressure side and a low pressure side and including a plurality of seal segments, the plurality of seal segments having a first seal segment, the first seal segment having a seal face configured to form a fluid bearing with the rotor, a lip assembly, and a body, the lip assembly positioned on the high pressure side, the lip assembly including a seal lip having a high pressure surface defining a first angle with an axial direction and a low pressure surface defining a second angle with the axial direction, the second angle being greater than the first angle.

FIELD

The present disclosure relates to a seal support assembly for a turbine engine.

BACKGROUND

Gas turbine engines, such as turbofan engines, may be used for aircraft propulsion. A turbofan engine generally includes a bypass fan section and a turbomachine such as a gas turbine engine to drive the bypass fan. The turbomachine generally includes a compressor section, a combustion section, and a turbine section in a serial flow arrangement. Both the compressor section and the turbine section are driven by one or more rotor shafts and generally include multiple rows or stages of rotor blades coupled to the rotor shaft. Each individual row of rotor blades is axially spaced from a successive row of rotor blades by a respective row of stator or stationary vanes. A radial gap is formed between an inner surface of the stator vanes and an outer surface of the rotor shaft.

DETAILED DESCRIPTION

The term “at least one of” in the context of, e.g., “at least one of A, B, and C” refers to only A, only B, only C, or any combination of A, B, and C.

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

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

The term “combustion section” refers to any heat addition system for a turbomachine. For example, the term combustion section may refer to a section including one or more of a deflagrative combustion assembly, a rotating detonation combustion assembly, a pulse detonation combustion assembly, or other appropriate heat addition assembly. In certain example embodiments, the combustion section may include an annular combustor, a can combustor, a cannular combustor, a trapped vortex combustor (TVC), or other appropriate combustion system, or combinations thereof.

The terms “low” and “high”, or their respective comparative degrees (e.g., -er, where applicable), when used with a compressor, a turbine, a shaft, or spool components, etc. each refer to relative speeds within an engine unless otherwise specified. For example, a “low turbine” or “low speed turbine” defines a component configured to operate at a rotational speed, such as a maximum allowable rotational speed, lower than a “high turbine” or “high speed turbine” of the engine.

The term “spring extension” refers to an object that is configured to deform elastically and store mechanical energy as a result of such deformation. A spring extension may be configured to deform linearly through extension or compression; may be configured to deform in a twisting manner through rotation about its axis; or in any other suitable manner.

The term “adjacent” with respect to a relative position of two like components refers to there being no other like components positioned therebetween. The term “adjacent” with respect to a relative position of two different components refers to there being no intervening structure separating the two components.

The term “shape memory alloy material” and “shape memory alloy (SMA)” generally refer to a metal alloy that experiences a temperature-related or strain-related, solid-state, micro-structural phase change. An SMA material may change from one physical shape to another physical shape. The temperature at which a phase change occurs generally is called the critical or transition temperature of the SMA. The SMA material may be constructed of a single SMA or of various SMA materials. In an embodiment, high temperature SMA may define transition temperatures ranging between about 20 degrees Celsius and about 1400 degrees Celsius. The transition temperature of the SMA may be tunable to specific applications.

In some embodiments, a component said to be formed of a SMA may include the SMA material as a major constituent, e.g., in an amount greater than 50 weight percent (“wt. %”) of the component. In certain embodiments, the component may be essentially composed of the SMA material (e.g., at least 90 wt. %, such as at least 95 wt. %, such as 100 wt. %).

A SMA material is generally an alloy capable of returning to its original shape after being deformed. For instance, SMA materials may define a hysteresis effect where the loading path on a stress-strain graph is distinct from the unloading path on the stress-strain graph. Thus, SMA materials may provide improved hysteresis damping as compared to traditional elastic materials.

A SMA material may also provide varying stiffness, in a pre-determined manner, in response to certain ranges of temperatures. The change in stiffness of the shape memory alloy may be due to a temperature related, solid state micro-structural phase change that enables the alloy to change from one physical shape to another physical shape. The changes in stiffness of the SMA material may be developed by working and annealing a preform of the alloy at or above a temperature at which the solid state micro-structural phase change of the shape memory alloy occurs. Such may allow a component formed of a SMA to act as a spring extension having a desired stiffness profile.

In the manufacture of a component comprising SMA (also referred to as an SMA component) intended to change stiffness during operation of a gas turbine engine, the component may be formed to have one operative stiffness (e.g., a first stiffness) below a transition temperature and have another stiffness (e.g., a second stiffness) at or above the transition temperature.

The term “temperature-dependent shape memory alloy material” refers to a SMA characterized by a temperature-dependent phase change. These phases include a martensite phase and an austenite phase. The martensite phase generally refers to a lower temperature phase. Whereas the austenite phase generally refers to a higher temperature phase. The martensite phase is generally more deformable, while the austenite phase is generally less deformable. When the shape memory alloy is in the martensite phase and is heated to above a certain temperature, the shape memory alloy begins to change into the austenite phase. The temperature at which this phenomenon starts is referred to as the austenite start temperature (As). The temperature at which this phenomenon is completed is called the austenite finish temperature (Af). When the shape memory alloy, which is in the austenite phase, is cooled, it begins to transform into the martensite phase. The temperature at which this transformation starts is referred to as the martensite start temperature (Ms). The temperature at which the transformation to martensite phase is completed is called the martensite finish temperature (Mf). As used herein, the term “transition temperature” without any further qualifiers may refer to any of the martensite transition temperature and austenite transition temperature. Further, “below transition temperature” without the qualifier of “start temperature” or “finish temperature” generally refers to the temperature that is lower than the martensite finish temperature, and the “above transition temperature” without the qualifier of “start temperature” or “finish temperature” generally refers to the temperature that is greater than the austenite finish temperature.

In some embodiments, a SMA component (such as a spring extension formed of an SMA material) may define a first stiffness at a first temperature and define a second stiffness at a second temperature, wherein the second temperature is different from the first temperature. Further, in some embodiments, one of the first temperature or the second temperature is below the transition temperature and the other one may be at or above the transition temperature. Thus, in some embodiments, the first temperature may be below the transition temperature and the second temperature may be at or above the transition temperature. While in some other embodiments, the first temperature may be at or above the transition temperature and the second temperature may be below the transition temperature. Further, various embodiments of SMA components described herein may be configured to have different first stiffnesses and different second stiffnesses at the same first and second temperatures.

The term “strain dependent shape memory alloy material” refers to a SMA characterized by a strain-dependent phase change. These phases similarly include a martensite phase and an austenite phase, which function in a similar manner as with the temperature dependent shape memory alloy materials, but instead of defining a transition temperature, the strain dependent SMAs define a transition strain

Non-limiting examples of SMAs that may be suitable for forming various embodiments of the SMA components described herein may include nickel-titanium (NiTi) and other nickel-titanium based alloys such as nickel-titanium hydrogen fluoride (NiTiHf) and nickel-titanium palladium (NiTiPd). However, it should be appreciated that other SMA materials may be equally applicable to the current disclosure. For instance, in certain embodiments, the SMA material may include a nickel-aluminum based alloys, copper-aluminum-nickel alloy, or alloys containing zinc, zirconium, copper, gold, platinum, and/or iron. The alloy composition may be selected to provide the desired stiffness effect for the application such as, but not limited to, damping ability, transformation temperature and strain, the strain hysteresis, yield strength (of martensite and austenite phases), resistance to oxidation and hot corrosion, ability to change shape through repeated cycles, capability to exhibit one-way or two-way shape memory effect, and/or a number of other engineering design criteria. Suitable shape memory alloy compositions that may be employed with the embodiments of present disclosure may include, but are not limited to NiTi, NiTiHf, NiTiPt, NiTiPd, NiTiCu, NiTiNb, NiTiVd, TiNb, CuAlBe, CuZnAl and some ferrous based alloys. In some embodiments, NiTi alloys having transition temperatures between 5 degrees C. and 150 degrees C. are used. NiTi alloys may change from austenite to martensite upon cooling.

Moreover, SMA materials may also display superelastic properties. Superelasticity may generally be characterized by recovery of large strains, potentially with some dissipation. For instance, martensite and austenite phases of the SMA material may respond to mechanical stress as well as temperature induced phase transformations. For example, SMAs may be loaded in an austenite phase (i.e. above a certain temperature). As such, the material may begin to transform into the (twinned) martensite phase when a critical stress is reached. Upon continued loading and assuming isothermal conditions, the (twinned) martensite may begin to detwin, allowing the material to undergo plastic deformation. If the unloading happens before plasticity, the martensite may generally transform back to austenite, and the material may recover its original shape by developing a hysteresis.

The term “bimetallic material” refers to a material having a first layer formed of a first material and a second layer formed of a second material, with the first and second materials configured to expand differently in response to temperature, strain, or a combination thereof. For example, the first material may define a first coefficient of thermal expansion and the second material may define a second coefficient of thermal expansion different than the first coefficient of thermal expansion. Additionally or alternatively one of the first material or the second material may be a SMA material configured to expand differently than the other of the first material or the second material in response to operating conditions to which the bimetallic material is expected to be exposed.

The present disclosure is generally related to a seal member support system for a turbomachine of a gas turbine engine. A turbomachine generally 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 serial-flow order. Each of the low-pressure compressor, the high-pressure compressor, the high-pressure turbine and the low-pressure turbine include sequential rows of stationary or stator vanes axially spaced by sequential rows of rotor blades. The rotor blades are generally coupled to a rotor shaft and the stator vanes are mounted circumferentially in a ring configuration about an outer surface of the rotor shaft. Radial gaps are formed between the outer surface of the rotor shaft and an inner portion of each ring or row of stator vanes.

During operation, it is desirable to control (reduce or prevent) compressed air flow or combustion gas flow leakage through these radial gaps. Ring seals are used to form a film bearing seal to seal these radial gaps. Ring seals generally include a plurality of seal shoe or seal member segments. As pressure builds in the compressor section and/or the turbine section, the seal members are forced radially outwardly and form a bearing seal between the outer surface of the rotor shaft and the respective seal members. To reduce wear on the rotor shaft and/or the seal members, it is desirable to maintain a positive radial clearance between the seal members and the outer surface of the rotor shaft under all operating conditions of the turbomachine. However, at low delta pressure operating conditions and transients like during start-up, stall, rotor vibration events, or during sudden pressure surges within the turbomachine, the film bearing stiffness may be low or suddenly change thus leading to seal member/rotor rubs.

A seal support assembly may be provided that allows for the seal to move along the radial direction at various engine operating condition to accommodate, e.g., vibration events at low delta pressure operating conditions and further to establish relatively tight clearances with the rotor during high delta pressure operating conditions.

Disclosed herein is a turbine engine defining an axial direction. The turbine engine includes: a rotor; a stator comprising a carrier; a seal support assembly coupled to the carrier; and a seal assembly disposed between the rotor and the stator and supported by the seal support assembly. The seal assembly defines a high pressure side and a low pressure side and includes a plurality of seal segments. The plurality of seal segments includes a first seal segment, the first seal segment having a seal face configured to form a fluid bearing with the rotor. The first seal segment further includes a lip assembly and a body. The lip assembly is positioned on the high pressure side and includes a seal lip having a high pressure surface defining a first angle with the axial direction and a low pressure surface defining a second angle with the axial direction, the second angle being greater than the first angle. The configuration of the lip assembly having the seal lip may allow for the lip assembly to accommodate radial movement of the first seal segment during various operations of the turbine engine, while maintaining a relatively tight clearance with the rotor to function as an airflow seal with the rotor.

Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures,FIG.1is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment ofFIG.1, the gas turbine engine is a high-bypass turbofan jet engine, sometimes also referred to as a “turbofan engine.” As shown inFIG.1, the gas turbine engine10defines an axial direction A (extending parallel to a longitudinal centerline12provided for reference), a radial direction R. and a circumferential direction C extending about the longitudinal centerline12. In general, the gas turbine engine10includes a fan section14and a turbomachine16disposed downstream from the fan section14.

The exemplary turbomachine16depicted generally includes a substantially tubular outer casing18that defines an annular inlet20. The outer casing18encases, in serial flow relationship, a compressor section including a booster or low-pressure (LP) compressor22and a high-pressure (HP) compressor24; a combustion section26; a turbine section including a high-pressure (HP) turbine28and a low-pressure (LP) turbine30; and a jet exhaust nozzle section32. A high-pressure (HP) shaft34(which may additionally or alternatively be a spool) drivingly connects the HP turbine28to the HP compressor24. A low-pressure (LP) shaft36(which may additionally or alternatively be a spool) drivingly connects the LP turbine30to the LP compressor22. The compressor section, combustion section26, turbine section, and jet exhaust nozzle section32together define a working gas flowpath37.

For the embodiment depicted, the fan section14includes a fan38having a plurality of fan blades40coupled to a disk42in a spaced apart manner. As depicted, the fan blades40extend outwardly from disk42generally along the radial direction R R. Each fan blade40is rotatable relative to the disk42about a pitch axis P by virtue of the fan blades40being operatively coupled to a suitable pitch change mechanism44configured to collectively vary the pitch of the fan blades40, e.g., in unison. The gas turbine engine10further includes a power gear box46, and the fan blades40, disk42, and pitch change mechanism44are together rotatable about the longitudinal centerline12by LP shaft36across the power gear box46. The power gear box46includes a plurality of gears for adjusting a rotational speed of the fan38relative to a rotational speed of the LP shaft36, such that the fan38may rotate at a more efficient fan speed.

Referring still to the exemplary embodiment ofFIG.1, the disk42is covered by rotatable front hub48of the fan section14(sometimes also referred to as a “spinner”). The front hub48aerodynamically contoured to promote an airflow through the plurality of fan blades40.

Additionally, the exemplary fan section14includes an annular fan casing or outer nacelle50that circumferentially surrounds the fan38and/or at least a portion of the turbomachine16. It should be appreciated that the nacelle50is supported relative to the turbomachine16by a plurality of circumferentially-spaced outlet guide vanes52in the embodiment depicted. Moreover, a downstream section54of the nacelle50extends over an outer portion of the turbomachine16so as to define a bypass airflow passage56therebetween.

During operation of the gas turbine engine10, a volume of air58enters the gas turbine engine10through an associated inlet60of the nacelle50and fan section14. As the volume of air58passes across the fan blades40, a first portion of air62is directed or routed into the bypass airflow passage56and a second portion of air64as indicated by arrow64is directed or routed into the working gas flowpath37, or more specifically into the LP compressor22. The ratio between the first portion of air62and the second portion of air64is commonly known as a bypass ratio. A pressure of the second portion of air64is then increased as it is routed through the HP compressor24and into the combustion section26, where it is mixed with fuel and burned to provide combustion gases66.

It should be appreciated, however, that the exemplary gas turbine engine10depicted inFIG.1is by way of example only, and that in other exemplary embodiments, the gas turbine engine10may have any other suitable configuration. For example, although the gas turbine engine10depicted is configured as a ducted gas turbine engine (i.e., including the outer nacelle50), in other embodiments, the gas turbine engine10may be an unducted gas turbine engine (such that the fan38is an unducted fan, and the outlet guide vanes52are cantilevered from, e.g., the outer casing18). Additionally, or alternatively, although the gas turbine engine10depicted is configured as a geared gas turbine engine (i.e., including the power gear box46) and a variable pitch gas turbine engine (i.e., including a fan38configured as a variable pitch fan), in other embodiments, the gas turbine engine10may additionally or alternatively be configured as a direct drive gas turbine engine (such that the LP shaft36rotates at the same speed as the fan38), as a fixed pitch gas turbine engine (such that the fan38includes fan blades40that are not rotatable about a pitch axis P), or both. It should also be appreciated, that in still other exemplary embodiments, aspects of the present disclosure may be incorporated into any other suitable gas turbine engine. For example, in other exemplary embodiments, aspects of the present disclosure may (as appropriate) be incorporated into, e.g., a turboprop gas turbine engine, a turboshaft gas turbine engine, or a turbojet gas turbine engine.

Referring now toFIG.2, a cross sectional, schematic view of a portion of the turbomachine16ofFIG.1is provided. As will be appreciated, the exemplary turbomachine16generally includes a rotor100, a stator102having a carrier104, a seal assembly106disposed between the rotor100and the stator102, and a seal support assembly108. The rotor100may be any rotor of the turbomachine16, such as the LP shaft36, the HP shaft34, etc. By way of example, referring briefly back toFIG.1, Circles SA have been added toFIG.1to provide example locations that the seal assembly106and seal support assembly108of the present disclosure may be incorporated into a turbomachine of the present disclosure.

Referring still toFIG.2, and as will be explained in more detail below, the exemplary seal assembly106includes a plurality of seal segments110arranged along the circumferential direction C. Each seal segment110of the plurality of seal segments110has a seal face112configured to form a fluid bearing with the rotor100, and more specifically a radial fluid bearing (i.e., configured to constrain the rotor100along the radial direction R).

The seal support assembly108includes a spring arrangement114extending between the carrier104and a first seal segment110A of the plurality of seal segments110to support the plurality of seal segments110of the seal assembly106. The seal support assembly108may further include similar spring arrangements114extending between the carrier104and the other seal segments110of the plurality of seal segments110.

Further, referring now toFIG.3, a close-up, schematic, cross-sectional view is depicted, taken along Line3-3andFIG.2. In particular,FIG.3depicts the first seal segment110A of the plurality of seal segments110positioned between the rotor100and the carrier104of the stator102.

As will be appreciated, the stator102further includes a stator vane116and the seal assembly106is, in the embodiment depicted, positioned at an inner end of a stator vane116along the radial direction R of the turbomachine16. The turbomachine16further includes a first stage118of rotor blades120and a second stage122of rotor blades120spaced along the axial direction A of the gas turbine engine10. The seal assembly106is positioned between the first stage118of rotor blades120and the second stage122of rotor blades120along the axial direction A.

In the embodiment depicted, the seal assembly106is positioned within a turbine section of the gas turbine engine10, such as within the HP turbine28or the LP turbine30. In such a manner, it will be appreciated that the rotor100may be a rotor coupled to the HP turbine28, such as the HP shaft34, or a rotor coupled to the LP turbine30, such as the LP shaft36. More specifically, still, in the embodiment affected, the rotor100is a connector extending between a disk124of the first stage118of rotor blades120and a disk124of the second stage of rotor blades120.

It will further be appreciated that the seal assembly106defines a high-pressure side126and a low-pressure side128. The seal assembly106is operable to prevent or minimize an airflow from the high-pressure side126to the low-pressure side128between the rotor100and the seal assembly106. In particular, it will be appreciated that the first seal segment110A depicted includes the seal face112configured to form a fluid bearing with the rotor100to support the rotor100along the radial direction R and prevent or minimize the airflow from the high-pressure side126to the low-pressure side128between the rotor100and the seal assembly106.

As will be appreciated, the first seal segment110A may be in fluid communication with a high-pressure air source to provide a high-pressure fluid flow to the seal face112to form the fluid bearing with the rotor100. In at least certain exemplary aspects, the high-pressure air source may be the working gas flowpath37through the gas turbine engine10and the seal assembly106, and more specifically the first seal segment110A, may be in fluid communication with the high-pressure air source, e.g., at the high-pressure side126of the seal assembly106.

In particular, for the embodiment depicted, referring back briefly also toFIG.1, the gas turbine engine10further includes a high-pressure air duct130extending from the high-pressure air source and in fluid communication with scal assembly106. As noted, the high-pressure air source is the working gas flowpath37, and more specifically is a portion of the working gas flowpath defined by the HP compressor24of the compressor section (seeFIG.1). The high-pressure air duct130extends to and through the stator vane116and to a high-pressure cavity132defined at the high-pressure side126of the seal assembly106(e.g., between the stator102and the rotor100). A high-pressure airflow from the high-pressure air duct130may pressurize the high-pressure cavity132to prevent gasses from the working gas flowpath37(which may be combustion gasses) from entering the high-pressure cavity132and damaging one or more components exposed thereto. The high-pressure airflow may also feed the seal assembly106. For example, although not depicted, it will be appreciated that the exemplary first seal segment110A may define a plurality of air ducts extending therethrough, extending between one or more inlets in airflow communication with the high-pressure cavity132and one or more outlets in airflow communication with the seal face112to provide a necessary high-pressure airflow to form the fluid bearing with the rotor100.

It will be appreciated, however, that in other exemplary embodiments, the seal assembly106may be integrated into, e.g., a compressor section of the gas turbine engine10. In such a case, the high-pressure side126may be positioned on a downstream side or aft side of seal assembly106, and the low-pressure side128may be positioned on an upstream side forward side of the seal assembly106.

Referring now also toFIG.4, a close-up, schematic, cross-sectional view is provided of the rotor100, carrier104, first seal segment110A, and seal support assembly108ofFIG.3. As will be appreciated, the exemplary first seal segment110A depicted further includes a lip assembly136and a body138. The lip assembly136extends from the body138along the axial direction A of the gas turbine engine10on the high-pressure side126of the seal assembly106. The lip assembly136includes an outer pressurization surface140along the radial direction R of the gas turbine engine10. For the embodiment depicted, the outer pressurization surface140is in airflow communication with the working gas flowpath37of the gas turbine engine10, and more specifically is exposed to the high-pressure cavity132and thus is in fluid communication with the working gas flowpath37of the gas turbine engine10from the high-pressure side126of the seal assembly106. The outer pressurization surface140is a radially outer surface of the lip assembly136, and as will be appreciated, as a pressure within the high-pressure cavity132increases, a radially-inward force exerted on the outer pressurization surface140(and the first seal segment110A) correspondingly increases.

The seal support assembly108, more specifically, the spring arrangement114of the seal support assembly108, extends between the carrier104and the first seal segment110A to counter a pressure on the outer pressurization surface140during operation of the gas turbine engine10, while allowing for passive control of a radial clearance gap defined between the seal face112and the rotor100during operation of the turbine engine.

For example, the seal support assembly108may generally define a resistance along the radial direction R of the gas turbine engine10. The gas turbine engine10may define a first high-pressure at the high-pressure side126of the seal assembly106(i.e., within the high-pressure cavity132) when the gas turbine engine10is operated in the high power operating mode, and may further define the second high-pressure at the high-pressure side126of the seal assembly106(i.e., within the high-pressure cavity132) when the gas turbine engine10is operated in the low power operating mode. The seal support assembly108is configured to hold the first seal segment110A at a radial distance away from the rotor100when the gas turbine engine10defines the second high-pressure. By contrast, the seal support assembly108is configured to move the first seal segment110A (or rather, allow the first seal segment110A to move) towards the rotor100when the gas turbine engine10defines the first high-pressure. In such a manner, the seal support assembly108may allow for the first seal segment110A to be moved closer to the rotor100during the high-pressure operating mode as compared to during the low-pressure operating mode.

Such a configuration may allow for a higher radial clearance between the first seal segment110A and the rotor100during low-pressure operating conditions and transients, which may allow for accommodation of, e.g., rotor vibrations with a reduced amount of rub between the rotor100and the first seal segment110A. Such a configuration may also allow for a lower radial clearance between the first seal segment110A and the rotor100during high-pressure operating conditions when, e.g., rotor100vibrations may be less severe. As will be appreciated, the first seal segment110A may be more effective at preventing or minimizing airflow from the high-pressure side126to the low-pressure side128with a lower radial clearance.

Referring still toFIG.4, and now also toFIG.5, a close-up view of the lip assembly136of the first seal segment110A is provided. The seal assembly106generally includes a main lip body150, a seal body152, and a connector154extending between the seal body152and the main lip body150.

The main lip body150is coupled to, or formed integrally with, the body138(seeFIG.4) of the first seal segment110A. The main lip body150defines the outer pressurization surface140in the exemplary embodiment depicted.

As noted above, the seal body152is coupled to the main lip body150through the connector154. For the embodiment shown, the connector154includes a section extending along the radial direction R of the turbine engine. The seal body152extends from the connector154in a direction towards the low-pressure side128of the seal assembly106. In such a manner, it will be appreciated that the seal body152of the lip assembly136is cantilevered from the main lip body150and body138of the first seal segment110A.

Referring particularly toFIG.5, the seal body152includes a seal lip156. The seal lip156includes a high-pressure surface158and a low-pressure surface160. The high-pressure surface158is located closer to the high-pressure cavity132than the low-pressure surface160.

Briefly, it will be appreciated that the seal body152of the lip assembly136further includes a dust lip162located forward of the seal lip156, and more specifically, located upstream of the seal lip156in the embodiment shown. The dust lip162defines a first gap164with the rotor100and the seal lip156defines a second gap166with the rotor100. The first gap164is larger than the second gap166during an operating condition of the turbine engine. In such a manner, an airflow from the high-pressure cavity132may flow into a seal body cavity168defined between the seal body152and the rotor100during operation of the gas turbine engine.

Further, it will be appreciated that the second gap166of the seal lip156defined with the rotor100may be relatively small so as to prevent at least a portion of the airflow from the high-pressure cavity132(and seal body cavity168) from passing between the seal lip156and the rotor100during operation. In such manner, it will be appreciated that the seal lip156, and more specifically, the first seal segment110A defines a low-pressure cavity170opposite the seal lip156from the high-pressure cavity132.

Further, still, it will be appreciated that the seal assembly106is designed to deflect along the radial direction R in response to an increase in a pressure differential between the high-pressure cavity132and the low-pressure cavity170. Such may generally prevent a rubbing between the seal lip156and the rotor100during such an operating condition.

More specifically, it will be appreciated that the high-pressure surface158of the seal lip156defines a first angle172with the axial direction A and the low-pressure surface160defines a second angle174with the axial direction A. The second angle174is greater than the first angle172. Both the first angle172and the second angle174are greater than 0 degrees and less than 90 degrees.

Notably, the radial direction R and the axial direction A together define a reference plane (the view depicted inFIG.5), and the first angle172and the second angle174are each defined within the reference plane.

Due at least in part to the differences in a magnitude of the first angle172and the second angle174, the high-pressure surface158defines a high-pressure surface length176and the low-pressure surface160defines a low-pressure surface length178. The high-pressure surface length176and the low-pressure surface length178are each defined within the reference plane.

The smaller size of the first angle172relative to the second angle174and the larger size of the high-pressure surface length176relative to the low-pressure surface length178may together facilitate a deflection of the seal body152, and more specifically of the seal lip156, inwardly along the radial direction R in response to an increase in pressure of the high-pressure cavity132(and seal body cavity168) relative to the low-pressure cavity170. For example, as will be appreciated, the above configuration provides for an increased amount of surface area on the high-pressure surface158relative to the low-pressure surface160, which may allow for a pressure of an air within the seal body cavity168defined between the high-pressure surface158and the rotor100to generate a larger amount of outward force along the radial direction R.

Furthermore, the inclusion of the dust lip162located upstream of the seal lip156may further allow for the seal body cavity168defined between the high-pressure surface158and the rotor100to maintain a pressure therein, which may further allow for the seal body152to form a fluid bearing with the rotor100located inward of the high-pressure surface158of the seal body152of the lip assembly136. In the embodiment depicted, the rotor100defines a cylindrical shape at such location.

Moreover, referring particularly toFIG.4, it will be appreciated that the second gap166defined between the seal lip156and the rotor100along the radial direction R is less than a gap180defined between the seal face112of the first seal segment110A and the rotor100along the radial direction R. In such manner, in the event of a rotor deflection, the seal lip156may still contact the rotor100prior to the seal face112of the first seal segment110A, despite the above-mentioned deflection of the seal body152. Although a contact is not desired, such a configuration may allow for single point of contact with the rotor100and may further prevent a contact between the seal face112and the rotor100, which may cause damage to the seal face112.

In certain exemplary embodiments, the connector154, the main lip body150, and the seal body152may be coupled to one another in any suitable manner. For example, referring briefly toFIGS.6and7, two exemplary embodiments of the present disclosure are provided. The exemplary embodiments ofFIGS.6and7may be configured in similar manner as the exemplary embodiment described above with referenceFIGS.4and5.

Referring particularly toFIG.6, in the embodiment depicted, a connector154of a lip assembly136of the embodiment shown is depicted coupled to a main lip body150of the lip assembly136by brazing or welding, the brazing or welding joint indicated at182. Such a configuration may provide for a relatively strong connection between these components.

Referring particularly toFIG.7, in the embodiment depicted, a connector154of a lip assembly136of the embodiment shown is depicted coupled to a main lip body150of the lip assembly136using a mechanical fastener184, such as a screw, bolt, or the like. Such a configuration may provide for an assembly that may be removed for maintenance and/or repair operations.

Notably, referring back briefly toFIGS.4and5, in order to provide the lip assembly136with a desired amount of resistance along the radial direction R, the connector154of the lip assembly136may be designed to provide a certain amount of stiffness. In at least certain exemplary embodiments, the connector154, the main lip body150, and the seal body152may be formed of a common material, such as the same metal material.

However, in other exemplary embodiments, the connector154may be formed of a material unique from a material of the seal body152, the main lip body150, or both. For example, in one exemplary embodiment, the connector154may be formed of a shape memory alloy material. The shape memory alloy material may be a temperature dependent shape memory alloy, or a strain-dependent shape memory alloy. In such a manner, a shape and/or stiffness of the connector154may change in response to an operating condition of the gas turbine engine.

Further, still, and other exemplary embodiments, the connector154may be formed of a bimetallic material.

Moreover, in still other exemplary embodiments, the lip assembly136may include other suitable structure to provide the desired amount of resistance along the radial direction R.

For example, referring briefly toFIG.8, an embodiment of the present disclosure is provided, which may be configured in a similar manner as one or more of the exemplary embodiments described hereinabove. In the embodiment depicted, a lip assembly136includes a radial spring member186operable with a seal body152of the lip assembly136to bias a seal lip156of the seal body152inwardly along a radial direction R of the turbine engine. In particular, for the embodiment shown, the radial spring member186is a helical spring extending between a main lip body150of the lip assembly136and the seal body152.

Additionally, or alternatively, referring now brieflyFIG.9, another embodiment of the present disclosure is provided, which again may be configured in a similar manner as one or more of the exemplary embodiments described hereinabove. The embodiment depicted inFIG.9includes a lip assembly136having a garter spring188positioned outward of a seal body152of the lip assembly136to bias a seal lip156of the seal body152inwardly along a radial direction R of the turbine engine.

One or more of these exemplary embodiments may be utilized to provide a lip assembly136having a desired amount of stiffness along the radial direction R during operation of the turbine engine.

In at least certain embodiments, the configuration of the lip assembly136may be selected to define a stiffness that is less than a stiffness of the fluid bearing formed between a seal face112of the first seal segment110and the rotor100of the turbine engine. In such a manner, the lip assembly136, and more specifically, the connector154and/or the seal body152of the lip assembly136may be configured to absorb stresses along the radial direction R, minimizing or reducing an amount of stress to be absorbed by the fluid bearing formed between the seal face112of the first seal segment110A and the rotor100of the turbine engine.

As will be appreciated, a functionality of the seal body152, and more specifically of the seal lip156, to restrict an airflow from the high-pressure cavity132to low-pressure cavity170may be enabled at least in part by minimizing airflow leakage between adjacent lip assemblies136along a circumferential direction C of the turbine engine. For example, in only certain exemplary embodiments, wherein a plurality of seal segments are provided along the circumferential direction C of the gas turbine engine, with each seal segment including a respective lip assembly136, it may be important to minimize a leakage of airflow from the high-pressure cavity132to the low-pressure cavity170between adjacent seal segments along the circumferential direction C.

In order to reduce such a leakage, referring still toFIG.9, it will be appreciated that the lip assembly136depicted defines a circumferential end190, and further defines a portion of a spline seal groove192at the circumferential end190. Referring toFIG.10, an example spline194is depicted that may be received within the portion of the spline seal groove192defined by the lip assembly136ofFIG.9. As will be appreciated, the spline194ofFIG.10is a multi-piece spline, formed of a plurality of individual spline segments195. Such may allow for relatively easy installation.

However, in other embodiments, the spline194may be a single-piece spline, extending continuously along its length. Such may provide for better airflow sealing capabilities. The exemplary splines194depicted inFIGS.6through8are single-piece splines.

Referring now toFIG.11, a schematic view is depicted of a first lip assembly136of the first seal segment110A and a second lip assembly136of a second seal segment110B in a view similar to the view indicated by Line11-11ofFIG.9. Referring also toFIG.12, a schematic view is depicted of the first lip assembly136of the first seal segment110A and the second lip assembly136of the second seal segment ofFIG.11, in a view similar to the view indicated by Line12-12FIG.9.

As will be appreciated, the first seal segment110A is positioned adjacent to the second seal segment110B in the circumferential direction C, more specifically, the lip assembly136of the first seal segment110A is positioned adjacent to the lip assembly136of the second seal segment110B in the circumferential direction C. A circumferential end of the lip assembly136of the first seal segment110A and a circumferential end of the lip assembly136of the second seal segment110B together define the spline seal groove192. The spline194is positioned within the spline seal groove192to prevent or minimize leakage from a high-pressure cavity132to a low-pressure cavity170(seeFIG.11) between adjacent lip assemblies136of adjacent seal segments110A,110B. Such a configuration may facilitate certain operations of the lip assembly136, such as one or more of the operations described hereinabove.

It will be appreciated, however, that in still other exemplary embodiments, a seal assembly106for a turbine engine may be provided in accordance with another exemplary embodiment of the present disclosure. For example, referring now toFIGS.13and14, a section of a turbine engine is provided in accordance with another embodiment. The exemplary embodiment ofFIGS.13and14may be configured in a similar manner as one or more of the exemplary embodiments described hereinabove.FIG.13depicts the assembly when the turbine engine is in a first operating condition, such as a low power operating condition; andFIG.14depicts the assembly when the turbine engine is in a second operating condition, such as a high power operating condition.

For example, the exemplary embodiment includes a rotor100, a stator102having a carrier104, a seal support assembly108coupled to the carrier104, and a seal assembly106positioned between the rotor100and the stator102and supported by the seal support assembly108. As with the embodiments described above, the seal assembly106defines a high-pressure side126and a low-pressure side128and includes a first seal segment110A. The first seal segment110A may be one of a plurality of seal segments110of the seal assembly106(see, e.g.,FIG.2). The first seal segment110A includes a seal face112configured to form a fluid bearing with the rotor100, a body138, and a lip assembly136. The lip assembly136is fixedly coupled to the body138on the high-pressure side126of the seal assembly106.

However, for the embodiment depicted, the lip assembly136includes a lip seal extension196and a main lip body150, with the lip seal extension196extending from the main lip body150. The lip seal extension196defines a height198along a radial direction R of the turbine engine. The lip seal extension196is formed of a material configured to change shape in a desired manner in response to anticipated operating conditions of the turbine engine. For example, the lip seal extension196defines a first shape when the turbine engine is operated in the first operating condition (FIG.13) and a second shape turbine engines operated in the second operating condition (FIG.14).

More specifically, for the embodiment depicted the lip seal extension196is formed of a shape memory alloy material, a bimetallic material, or both to reduce the height198of the lip seal extension196in response to an increase in a pressure to which it is exposure, an increase in a temperature to which it is exposure, or both.

More specifically, still, for the embodiment depicted, the lip seal extension196is formed of a shape memory alloy material and is configured to change from the shape depicted inFIG.13to the shape depicted inFIG.14in response to an increase in temperature of an airflow within a high-pressure cavity132at the high-pressure side126of the seal assembly106, in response to an increase in a pressure of the airflow within the high-pressure cavity132at the high-pressure side126of the seal assembly106, or both. In such a manner, it will be appreciated that the lip seal extension196may be formed of a temperature-dependent shape memory alloy material, or a strain-dependent shape memory alloy material, or both.

With such a configuration, the lip seal extension196may be configured to reduce a flow of air from the high-pressure cavity132to a low-pressure cavity170defined by the seal assembly106during, e.g., the low power operating condition of the turbine engine. However, as the turbine engine transitions to the high power operating condition of the turbine engine (during which a temperature and a pressure of an airflow with thin the high-pressure cavity132increases), the lip seal extension196is configured reduce its height198to allow the first seal segment110A to move inwardly along the radial direction R.

Accordingly, an inner edge200of the lip seal extension196along the radial direction R is configured to move outwardly along the radial direction R during such a transition. Notably, it will be appreciated that the main lip body150includes an inner edge202along the radial direction R at the high-pressure side126. The lip seal extension196is movable between the first position (FIG.13) in which the inner edge200of the lip seal extension196is positioned inward of the inner edge202of the main lip body150along the radial direction R, and the second position (FIG.14) in which the inner edge200of the lip seal extension196is positioned outward of the inner edge202of the main lip body150along the radial direction R.

In such a manner, it will be appreciated that the inner edge202of the main lip body150may function to reduce an airflow from the high-pressure cavity132to the low-pressure cavity170for the first seal segment110A when the turbine engines operated in the high power condition.

A turbine engine defining an axial direction, comprising: a rotor; a stator comprising a carrier; a seal support assembly coupled to the carrier; and a seal assembly disposed between the rotor and the stator and supported by the seal support assembly, the seal assembly defining a high pressure side and a low pressure side and comprising a plurality of seal segments, the plurality of seal segments having a first seal segment, the first seal segment having a seal face configured to form a fluid bearing with the rotor, a lip assembly, and a body, the lip assembly positioned on the high pressure side, the lip assembly including a seal lip having a high pressure surface defining a first angle with the axial direction and a low pressure surface defining a second angle with the axial direction, the second angle being greater than the first angle.

The turbine engine of any preceding clause, wherein the lip assembly includes a dust lip located upstream of the seal lip, wherein the dust lip defines a first gap with the rotor, wherein the seal lip defines a second gap with the rotor, and wherein the first gap is larger than the second gap at an operating condition of the turbine engine.

The turbine engine of any preceding clause, wherein the lip assembly includes a seal body having the seal lip, and wherein the seal body is cantilevered from the body of the first seal segment.

The turbine engine of any preceding clause, wherein the lip assembly includes a connector extending along a radial direction of the turbine engine, and wherein the seal body extends from the connector in a direction towards the low pressure side.

The turbine engine of any preceding clause, wherein the turbine engine defines a radial direction and a reference plane defined by the axial direction and the radial direction, and wherein the low pressure surface defines a length in the reference plane less than a length of the high pressure surface in the reference plane.

The turbine engine of any preceding clause, wherein the turbine engine defines a radial direction and a reference plane defined by the axial direction and the radial direction, wherein the first angle and the second angle are each defined in the reference plane.

The turbine engine of any preceding clause, wherein the rotor defines a cylindrical surface inward of the low pressure surface and inward of the high pressure surface of the lip assembly.

The turbine engine of any preceding clause, wherein the lip assembly includes a connector formed of a shape memory alloy material.

The turbine engine of any preceding clause, wherein the lip assembly includes a seal body having the seal lip and a radial spring extension operable with the seal body to bias the seal lip inwardly along a radial direction of the turbine engine.

The turbine engine of any preceding clause, wherein the lip assembly further includes a main lip body and a connector extending between the main lip body and the seal body, and wherein the radial spring extension is a helical spring extending between the main lip body and the seal body.

The turbine engine of any preceding clause, wherein the lip assembly includes a garter spring positioned outward of the seal body to bias the seal lip inwardly along a radial direction of the turbine engine.

The turbine engine of any preceding clause, wherein the lip assembly is a first lip assembly, wherein the plurality of seal segments further includes a second seal segment positioned adjacent to the first seal segment, wherein the second seal segment comprises a second lip assembly, wherein the first and second lip assemblies together define a spline seal groove, and wherein the seal support assembly includes a spline seal positioned within the spline seal groove.

The turbine engine of any preceding clause, wherein the spline seal is formed of a shape memory alloy material.

The turbine engine of any preceding clause, wherein the lip assembly is coupled to the body by brazing, using mechanical fasteners, or both.

A seal assembly for a turbine engine, the turbine engine having a rotor, a carrier, and a seal support assembly coupled to the carrier, the seal assembly comprising: a first seal segment configured to be disposed between the rotor and the carrier and supported by the seal support assembly, the first seal segment defining a high pressure side and a low pressure side and comprising a seal face configured to form a fluid bearing with the rotor, the first seal segment further comprising a lip assembly and a body, the lip assembly positioned on the high pressure side, the lip assembly including a seal lip having a high pressure surface configured to define a first angle with an axial direction of the turbine engine and a low pressure surface configured to define a second angle with the axial direction, the second angle being greater than the first angle.

The seal assembly of any preceding clause, wherein the lip assembly includes a dust lip located upstream of the seal lip, wherein the dust lip is configured to define a first gap with the rotor, wherein the seal lip is configured to define a second gap with the rotor, and wherein the first gap is larger than the second gap at an operating condition of the turbine engine.

The seal assembly of any preceding clause, wherein the turbine engine defines a radial direction and a reference plane defined by the axial direction and the radial direction, and wherein the low pressure surface defines a length in the reference plane less than a length of the high pressure surface in the reference plane.

A turbine engine defining a radial direction, the turbine engine comprising: a rotor; a stator comprising a carrier; a seal support assembly coupled to the carrier; and a seal assembly disposed between the rotor and the stator and supported by the seal support assembly, the seal assembly defining a high pressure side and a low pressure side and comprising a plurality of seal segments, the plurality of seal segments having a first seal segment, the first seal segment having a seal face configured to form a fluid bearing with the rotor, a lip seal extension, and a body, the lip seal extension fixedly coupled to the body on the high pressure side and defining a height along the radial direction, the lip seal extension formed or a shape memory alloy material, a bimetallic material, or both to reduce the height in response to an increase in pressure exposure, an increase in temperature exposure, or both.

The turbine engine of any preceding clause, wherein the lip seal extension is formed of a shape memory alloy material.

The turbine engine of any preceding clause, wherein the body includes an inner edge along the radial direction at the high pressure side, wherein the lip seal extension is moveable between a first position in which an end of the lip seal extension is positioned inward of the inner edge along the radial direction and a second position in which the end of the lip seal extension is positioned outward of the inner edge along the radial direction.