Patent Description:
Gas turbine engine typically include a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-energy exhaust gas flow. The high-energy exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.

Shafts coupled between the turbine and compressor sections are supported on bearings disposed within bearing compartments. The bearing compartments are isolated from high pressure and temperature regions within the engine. To maintain the environment within a bearing compartment a seal between a static and rotating part is often employed. Seals are typically biased against a rotating face at a defined pressure to provide the desired seal. Such seals may wear prematurely, causing decreased sealing performance.

<CIT> discloses cooling features for a hydrodynamic seal seat.

<CIT> discloses a compressor with associated impeller shaft seal.

From a first aspect of the invention, a seal assembly for a gas turbine engine as claimed in claim <NUM> is provided.

In various embodiments, the seal body contains graphite.

From a further aspect of the invention, a bearing assembly for a gas turbine engine as claimed in claim <NUM> is provided.

In various embodiments, the seal body comprises graphite.

In various embodiments, the shaft comprises a radially extending portion, and the recess is defined in the radially extending portion of the shaft. In various embodiments, the recess is defined in an axially aft facing surface of the radially extending portion. In various embodiments, the bearing assembly further includes a seal holder and a biasing member. The seal body may be supported by the seal holder, and the seal holder may be coupled to the engine static structure via the biasing member.

From a still further aspect, the invention provides a method of forming a seal between an engine static structure and a shaft of a gas turbine engine as claimed in claim <NUM>.

The detailed description of exemplary embodiments herein makes reference to the accompanying drawings, which show exemplary embodiments by way of illustration. While these exemplary embodiments are described in sufficient detail to enable those skilled in the art to practice the inventions, it should be understood that other embodiments may be realized and that logical changes and adaptations in design and construction may be made in accordance with the present inventions and the teachings herein. The scope of the present inventions is defined by the appended claims.

Disclosed herein, according to various embodiments, is a seal assembly for gas turbine engines that includes a seal seat insert coupled to the rotating side of the seal assembly. In various embodiments, the seal seat insert provides various benefits over coatings or other applied materials. More specifically, the seal seat insert disclosed herein may inhibit premature wear and may extend/increase sealing performance of the seal assembly. While numerous details are included herein pertaining to assemblies and method pertaining to implementing the seal seat insert in a gas turbine engine, the seal seat insert and the associated methods/systems may be used in other seal assemblies.

In various embodiments and with reference to <FIG>, a gas turbine engine <NUM> is provided. Gas turbine engine <NUM> may be a two-spool turbofan that generally incorporates a fan section <NUM>, a compressor section <NUM>, a combustor section <NUM> and a turbine section <NUM>. In operation, fan section <NUM> can drive fluid (e.g., air) along a bypass flow-path B while compressor section <NUM> can drive fluid along a core flow-path C for compression and communication into combustor section <NUM> then expansion through turbine section <NUM>. Although depicted as a turbofan gas turbine engine <NUM> herein, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.

Gas turbine engine <NUM> may generally comprise a low speed spool <NUM> and a high speed spool <NUM> mounted for rotation about an engine central longitudinal axis A-A' relative to an engine static structure <NUM> or engine case via several bearing systems <NUM>, <NUM>-<NUM>, and <NUM>-<NUM>. Engine central longitudinal axis A-A' is oriented in the z direction (axial direction) on the provided xyz axis. The y direction on the provided xyz axis refers to a radial direction, and the x direction on the provided xyz axis refers to a circumferential direction. It should be understood that various bearing systems <NUM> at various locations may alternatively or additionally be provided, including for example, bearing system <NUM>, bearing system <NUM>-<NUM>, and bearing system <NUM>-<NUM>.

Low speed spool <NUM> may generally comprise an inner shaft <NUM> that interconnects a fan <NUM>, a low pressure compressor <NUM> and a low pressure turbine <NUM>. Inner shaft <NUM> may be connected to fan <NUM> through a geared architecture <NUM> that can drive fan <NUM> at a lower speed than low speed spool <NUM>. Geared architecture <NUM> may comprise a gear assembly <NUM> enclosed within a gear housing. Gear assembly <NUM> couples inner shaft <NUM> to a rotating fan structure. High speed spool <NUM> may comprise an outer shaft <NUM> that interconnects a high pressure compressor <NUM> and high pressure turbine <NUM>.

A combustor <NUM> may be located between high pressure compressor <NUM> and high pressure turbine <NUM>. The combustor section <NUM> may have an annular wall assembly having inner and outer shells that support respective inner and outer heat shielding liners. The heat shield liners may include a plurality of combustor panels that collectively define the annular combustion chamber of the combustor <NUM>. An annular cooling cavity is defined between the respective shells and combustor panels for supplying cooling air. Impingement holes are located in the shell to supply the cooling air from an outer air plenum and into the annular cooling cavity.

A mid-turbine frame <NUM> of engine static structure <NUM> may be located generally between high pressure turbine <NUM> and low pressure turbine <NUM>. Mid-turbine frame <NUM> may support one or more bearing systems <NUM> in turbine section <NUM>. Inner shaft <NUM> and outer shaft <NUM> may be concentric and rotate via bearing systems <NUM> about the engine central longitudinal axis A-A', which is collinear with their longitudinal axes. As used herein, a "high pressure" compressor or turbine experiences a higher pressure than a corresponding "low pressure" compressor or turbine.

The core airflow C may be compressed by low pressure compressor <NUM> then high pressure compressor <NUM>, mixed and burned with fuel in combustor <NUM>, then expanded over high pressure turbine <NUM> and low pressure turbine <NUM>. Turbines <NUM>, <NUM> rotationally drive the respective low speed spool <NUM> and high speed spool <NUM> in response to the expansion.

In various embodiments, geared architecture <NUM> may be an epicyclic gear train, such as a star gear system (sun gear in meshing engagement with a plurality of star gears supported by a carrier and in meshing engagement with a ring gear) or other gear system. Geared architecture <NUM> may have a gear reduction ratio of greater than about <NUM> and low pressure turbine <NUM> may have a pressure ratio that is greater than about five (<NUM>). In various embodiments, the bypass ratio of gas turbine engine <NUM> is greater than about ten (<NUM>:<NUM>). In various embodiments, the diameter of fan <NUM> may be significantly larger than that of the low pressure compressor <NUM>, and the low pressure turbine <NUM> may have a pressure ratio that is greater than about five (<NUM>:<NUM>). Low pressure turbine <NUM> pressure ratio may be measured prior to inlet of low pressure turbine <NUM> as related to the pressure at the outlet of low pressure turbine <NUM> prior to an exhaust nozzle. It should be understood, however, that the above parameters are exemplary of various embodiments of a suitable geared architecture engine and that the present disclosure contemplates other gas turbine engines including direct drive turbofans. A gas turbine engine may comprise an industrial gas turbine (IGT) or a geared aircraft engine, such as a geared turbofan, or non-geared aircraft engine, such as a turbofan, or may comprise any gas turbine engine as desired.

Referring to <FIG>, a seal assembly <NUM> in an exemplary bearing compartment <NUM> is schematically shown. The bearing compartment <NUM> includes bearings <NUM> supporting rotation of a shaft <NUM>, according to various embodiments. The shaft <NUM> may be one of the inner shaft <NUM> and outer shaft <NUM> referenced above, or the shaft <NUM> may also be any other rotating shaft utilized within a gas turbine engine. A seal seat insert <NUM> is coupled to the shaft <NUM>, and a seal body <NUM> (e.g., a static seal body) is coupled to an engine static structure <NUM>. The seal body <NUM> is configured to be biased into direct contact with the seal seat insert <NUM> to provide a sealing pressure between the seal body <NUM> and the seal seat insert <NUM>.

The seal body <NUM> may be supported on a static structure <NUM> of the gas turbine engine. For example, the seal body <NUM> may be supported and/or retained by a seal holder <NUM>, and the seal holder <NUM> may be coupled to the engine static structure <NUM> via a biasing member <NUM>. The biasing member <NUM> may provide force to provide the sealing pressure between the seal body <NUM> and the seal seat insert <NUM>. The seal seat insert <NUM> is coupled to a radially extending portion <NUM> of the shaft <NUM>. Said differently, the shaft <NUM> may include radially extending portion <NUM>, and the seal seat insert <NUM> may be coupled thereto. The radially extending portion <NUM> is supported on the rotating shaft <NUM> such that it rotates relative to the fixed seal body <NUM>, according to various embodiments. The radially extending portion <NUM> may include a radial surface facing axially aft.

With momentary reference to <FIG>, the radially extending portion <NUM> of the shaft <NUM> defines a recess <NUM>, and the seal seat insert <NUM> is retained at least partially within the recess <NUM>. That is, the seal seat insert <NUM> is received within the recess <NUM> via an interference fit.

The seal seat insert <NUM> is retained within the recess <NUM>.

In various embodiments, the seal seat insert <NUM> may be an annular structure that is press fitted within the annular recess <NUM>. In various embodiments, the seal seat insert <NUM> may be a tight fit against the outer diameter of the recess <NUM>, thus maintaining a proper fit/engagement throughout thermal and centrifugal transient conditions. Further, by so engaging the seal seat insert <NUM> within the recess <NUM>, the seal seat insert <NUM> may maintain a favorable compressive stress, the seal seat insert <NUM> may be protected from damage during installation, and/or may facilitate retention of the seal seat insert <NUM> within the recess <NUM> in the event the seal seat insert <NUM> is fractured. In various embodiments, one or more fasteners disposed in the outer diameter of the recess <NUM> may be utilized to facilitate retention of the seal seat insert <NUM>.

In various embodiments, and with renewed reference to <FIG>, the seal body <NUM> is formed from a carbon material and provides a dry face seal that wears a predictable rate during operation of the gas turbine engine. In various embodiments, the seal body <NUM> provides sealing of the bearing compartment <NUM> against the environment surrounding the bearing compartment <NUM>. That is, the biasing member <NUM> may exert a force on the holder <NUM> and thereby the seal body <NUM> is forced against the seal seat insert <NUM> at a desired pressure. That is, the seal body <NUM> is a contact face seal, according to various embodiments. The pressure between the seal body <NUM> and the seal seat insert <NUM> may be within a desired range such that the seal body <NUM> and the seal seat insert <NUM> provide desired sealing performance (e.g., isolation of lubricant in the bearing compartment <NUM>).

The seal body <NUM> may be made from carbon materials, such as graphite. The seal seat insert <NUM> is made from a ceramic matrix composite. The seal seat insert <NUM> is made from a ceramic matrix composite including at least one of silicon carbide and/or silicon nitride. In various embodiments, the seal seat insert <NUM> comprises a plurality of stacked, layered, and/or wrapped matrix plies and/or weaves. In various embodiments, the ceramic matrix composite material may further comprise one or more of borides, carbides, oxides, and/or nitrides. In various embodiments the borides may be selected among a group comprising: ZrB2, HfB2, VB2, TiB2, TaB2, TaB, NbB2, NbB, VB2, TiB2, CrB2, Mo2B5, W2B5, Fe2B, FeB, Ni2B, NiB, LaB6, CoB, Co2B, or any other refractory boride. In various embodiments, the carbides may be selected among a group comprising: SiC, HfC, ZrC, C, B4C, SiOC, TiC, WC, Mo2C, TaC, NbC, or any other refractory carbide. In various embodiments, the oxides may be selected among a group comprising: HfO2, ZrO2, Al2O3, SiO2, class compositions including aluminosilicates, borosilicates, lithium aluminosilicates (LAS), magnesium aluminosilicates, barium magnesium aluminosilicates (BMAS), calcium aluminosilicates and other silica containing high temperature glasses, and/or other mixed metal oxides. In various embodiments, the nitrides may be selected among a group comprising: AlN, Si3N4, TaN, TiN, TiAlN, W2N, WN, WN2, VN, ZrN, BN, HfN, NbN, or any other refractory nitrides. In various embodiments, the CMC material may comprise mixed refractory nonoxides such as, for example, SiCN.

With the seal seat insert <NUM> being formed of a ceramic matrix composite material, the life of the seal assembly is improved. That is, because the seal seat insert <NUM> is thicker than, a conventional wear coating, the seal assembly <NUM> has an improved wear life because the seal seat insert <NUM> may be less susceptible to cracking and/or particle liberation, according to various embodiments. In various embodiments, the seal seat insert <NUM> is configured to operate effectively as a seal counterface at temperatures from about -<NUM> degrees Fahrenheit (about -<NUM> degrees Celsius) to about <NUM> degrees Fahrenheit (about <NUM> degrees Celsius). As used herein, the term "about" refers to plus or minus <NUM>% of the indicated value. The interface between the seal seat insert <NUM> and the seal body <NUM> may generate substantial friction heat during operation of the gas turbine engine, and the ceramic matrix composite may be wellsuited for the operational temperatures.

In various embodiments, and with reference to <FIG>, a method <NUM> of forming a seal between an engine static structure and a shaft of a gas turbine engine is provided. The method <NUM> includes supporting a seal body on the engine static structure at step <NUM>, coupling a seal seat insert to the shaft at step <NUM>, and generating a sealing pressure via direct contact between the seal body and the seal seat insert at step <NUM>. Coupling the seal seat insert to the shaft at step <NUM> comprises retaining the seal seat insert within a recess defined by the shaft. In various embodiments, step <NUM> may include heating the shaft to open/enlarge the recess to allow for insertion of the seal seat insert. Upon cooling, the seal seat insert may be retained within the recess with a tight fit. For example, the seal seat insert may be an annular structure, and the seal seat insert may be moved axially to be received into an annular recess defined by the shaft (e.g., a radially extending portion of the shaft). The seal seat insert is made from ceramic matrix composite including at least one of silicon carbide and siliscon nitride. In various embodiments, the seal seat insert is installed so as to be flush with the radially extending portion within which the seal seat insert is retained. Accordingly, the method may include polishing or otherwise grinding the seal seat insert to make the sealing interface surface flush with the radially extending portion.

Also, any reference to attached, fixed, connected, coupled or the like may include permanent (e.g., integral), removable, temporary, partial, full, and/or any other possible attachment option.

In the detailed description herein, references to "one embodiment", "an embodiment", "various embodiments", etc., indicate that the embodiment described may include a particular feature, structure, or characteristic, but every embodiment may not necessarily include the particular feature, structure, or characteristic.

Claim 1:
A seal assembly (<NUM>) for a gas turbine engine (<NUM>), the seal assembly (<NUM>) comprising:
a seal body (<NUM>) coupled to an engine static structure (<NUM>) of the gas turbine engine (<NUM>); and
a seal seat insert (<NUM>) coupled to a radially extending portion (<NUM>) of a shaft (<NUM>) of the gas turbine engine (<NUM>);
wherein the seal body (<NUM>) is configured to be biased into contact with the seal seat insert (<NUM>) to provide a sealing pressure between the seal body (<NUM>) and the seal seat insert (<NUM>), characterised by:
the seal seat insert (<NUM>) being made from a ceramic matrix composite including at least one of silicon carbide and silicon nitride;
wherein the seal seat insert (<NUM>) is retained within a recess (<NUM>) defined by the shaft (<NUM>) using an interference fit.